IPPW8 Abstract Book June 1

Abstract Guide

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WELCOME

Welcome to the 8th International Planetary Probe Workshop, and to historic Portsmouth, Virginia, USA. This year’s event is resuming an annual schedule, after a remarkable workshop in Barcelona last year. We have a full roster of participants, a varied program, and we are excited about the possibilities for collaboration. This year’s theme is technology development, as reflected in our Short Course and many of the oral and poster presentations. Our community has been very busy over the past year; all of our work has generated an outstanding set of presentations and posters that you will encounter in the next four and a half days. We are pleased to welcome an international group of scientists, technologists, engineers, mission designers, and policy makers to IPPW-8. Our committees have worked very hard in organizing the logistics for the workshop, planning the program, soliciting and evaluating nominees for the Al Seiff Award, and coordinating opportunities for student participation. We are delighted to host the meeting in the maritime city of Portsmouth, near the NASA-Langley Research Center. We recommend that all participants enjoy several vantage points throughout the week, and we hope you will take advantage of the exciting cultural and culinary experiences that await you here. We encourage you to attend as many oral and poster sessions as possible, in order to benefit from the world-wide planetary probe mission experts who are attending IPPW-8. We have scheduled a relaxing poster session on Tuesday evening. To better associate the submitted posters with their sessions, we will also have posters available in conjunction with each session. In keeping with agendas at previous IPPWs, we have scheduled parallel oral sessions only on Thursday. Our conveners will coordinate their timing so it will be possible to move back and forth between the parallel sessions in the morning and afternoon. Of interest to our student and early career attendees is a professional development session, also scheduled for Thursday. Since IPPW-8 is indeed a workshop, we also urge you to take advantage of the numerous opportunities during coffee breaks, lunches and social activities to build collaborative partnerships with other workshop participants. If you are joining us on the Wednesday afternoon tour of NASA-Langley, you will have the opportunity to see some unique, world-class facilities. In addition, the IPPW-8 sponsors have funded a significant number of students who would be interested in meeting the working planetary probe participants to gain a better understanding of how to build a future career in this exciting field. We are very encouraged to have a sizeable student population with us! On Friday, 10 June, there will be a presentation on the plans for IPPW-9 in 2012, in Europe. We encourage you to attend this talk to learn about your next opportunity to join our community. In this time of transition for many of our Agencies, it is all the more valuable for us to reconnect with our colleagues and celebrate our strong planetary probe foundations--please enjoy. Let’s make it a great week! Bernie Bienstock NASA Jet Propulsion Laboratory IPPW-8 International Organizing Committee US Co-Chair

Michelle Munk NASA-Langley Research Center IPPW-8 Local Organizing Committee Chair

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ACKNOWLEDGEMENTS

IPPW-8 Sponsors

Supporting Organizations

Microelectronics Research and Communications Institute College of Engineering

Dept. of Electrical and Computer Engineering

Dept. of Mechanical Engineering Department of Physics

NASA Idaho Space Grant Consortium Univ. of Idaho Office of Research and

Economic Development

Ablatives Laboratory

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IPPW-8 COMMITTEE MEMBERS

International Organizing Committee

Bernie Bienstock – CHAIR NASA Jet Propulsion Laboratory, USA

Ed Chester – CO-CHAIR Aevo GmbH, Germany

Mark Adler

NASA Jet Propulsion Laboratory, USA

Michael Amato NASA Goddard Space Flight Center,

USA

Marla Arcadi ERC/NASA Ames Research Center, USA

James Arnold

NASA Ames Research Center, USA

David Atkinson University of Idaho, USA

Tibor Balint

NASA Headquarters, USA

Eric Blood NASA Jet Propulsion Laboratory, USA

Jean-Marc Bouilly

Astrium Space Transportation, France

Robert Braun NASA Headquarters, USA

Neil Cheatwood

NASA Langley Research Center, USA

Athena Coustenis Observatorie de Paris, France

Jody Davis


Karl Edquist NASA Langley Research Center, USA

Kristin Gates-Medlock

Global Aerospace Corporation, USA

Rodrigo Haya Ramos Deimos Space, Spain

Jean-Pierre Lebreton

ESA/ESTEC

Michelle Munk NASA Langley Research Center, USA

Periklis Papadopoulos San Jose State University, USA

Stephen Ruffin

Georgia Space Grant Consortium/Georgia Institute of Technology, USA

Steve Sandford


Anita Sengupta NASA Jet Propulsion Laboratory, USA

Alexandre Solé

Open University, United Kingdom

Christine Szalai NASA Jet Propulsion Laboratory, USA

Ethiraj Venkatapathy


Michael Wright NASA Ames Research Center, USA

Program Organizing Committee

Karl Edquist – CHAIR NASA Langley Research Center, USA

Rodrigo Haya Ramos – CO-CHAIR

Deimos Space, Spain

David Atkinson

University of Idaho, USA

Bernie Bienstock NASA Jet Propulsion Laboratory, USA

Ed Chester

Aevo GmbH, Germany

Athena Coustenis Observatorie de Paris, France

Ioana Cozmuta

ERC/NASA Ames Research Center, USA

Kristin Gates-Medlock Global Aerospace Corporation, USA

Michelle Munk


Anita Sengupta


Thomas Spilker NASA Jet Propulsion Laboratory, USA

Christine Szalai


Al Seiff Award Committee

Ethiraj Venkatapathy – CHAIR NASA Ames Research Center, USA

James Arnold – CO-CHAIR



Bernie Bienstock


Jean-Marc Bouilly Astrium Space Transportation, France

Athena Coustenis

Observatorie de Paris, France

Short Course Organizing Committee

Tibor Balint – CHAIR NASA Headquarters, USA

Mark Adler


Dave Atkinson University of Idaho, USA

Bernard Bienstock


Michelle Munk


Mike Wright NASA Ames Research Center, USA

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Local Organizing Committee

Michelle Munk – CHAIR NASA Langley Research Center, USA

Neil Cheatwood – CO-CHAIR


Karl Edquist NASA Langley Research Center, USA

Jody Davis


Stacy Dees National Institute of Aerospace, USA

Bob Moses


Charlene Gleaton


Shannon Verstynen National Institute of Aerospace, USA

Michael Wagner

National Institute of Aerospace, USA

Student Organizing Committee

Stephen Ruffin – CHAIR (USA) Georgia Space Grant Consortium/Georgia Institute of Technology, USA

Alexandre Solé – CHAIR (EUROPE)

Open University, United Kingdom


Chris Carter

Virginia Space Grant Consortium, USA

Denise Dublin Virginia Space Grant Consortium, USA

Kristin Gates-Medlock Global Aerospace Corporation, USA

Michelle Munk


Periklis Papadopoulos San Jose State University, USA

Greg Swanson

Santa Clara University, USA

Christopher Tanner Georgia Institute of Technology, USA

CONTENTS

ABSTRACTS ............................................................................................................................................................ 14 SESSION 1 - OUTLOOK FOR PROBE MISSIONS ............................................................................................ 15 2011 AL SEIFF AWARD LECTURE..................................................................................................................................16 THE HUYGENS STORY.........................................................................................................................................................16 Jean-Pierre Lebreton...........................................................................................................................................................16

PLANETARY PROBES AND THE PLANETARY DECADAL SURVEY ...................................................................17 Amy Simon-Miller .................................................................................................................................................................17

25 YEARS OF DEEP SPACE EXPLORATION AT ESA.................................................................................................18 Marcello Coradini.................................................................................................................................................................18

NASA INVESTMENTS IN OUR FUTURE: EXPLORING SPACE THROUGH INNOVATION AND TECHNOLOGY.........................................................................................................................................................................19 Dr. Michael Gazarik .............................................................................................................................................................19

ESA EXPLORATION PROGRAMMES FROM ISS TO THE LUNAR LANDER MISSION ..................................20 Bruno Gardini.........................................................................................................................................................................20

PROGRESS TOWARD A COMPLETE RESPONSE TO THE PLANETARY DECADAL SURVEY....................21 Jim Adams ................................................................................................................................................................................21

SESSION 2 - PROBE MISSIONS .......................................................................................................................... 22 MARS SCIENCE LABORATORY ENTRY, DESCENT AND LANDING SYSTEM DESIGN, DEVELOPMENT AND PRELAUNCH STATUS................................................................................................................................................23 Adam Steltzner ......................................................................................................................................................................23

EXOMARS EDM MISSION AND DESIGN OVERVIEW...............................................................................................24 Olivier Bayle*, Leila Lorenzoni*, Thierry Blancquaert*, Stephane Langlois*, Thomas Walloschek*, S. Portigliotti§, M. Capuano§ ...........................................................................................................................................24

END TO END MISSION PERFORMANCES OF EXOMARS 2016 EDM.................................................................25 Rodrigo Haya-Ramos1 Mariano Sanchez Nogales1, Juan Luis Cano1 David Riley2, David Northey2, Stefano Portigliotti3, Olivier Bayle4 ..............................................................................................................................25

FUTURE MISSIONS AND TECHNOLOGIES WITHIN THE MARS ROBOTIC EXPLORATION PREPARATION (MREP) PROGRAMME ........................................................................................................................27 K. Geelen1, D. Agnolon1, P. Falkner1, J. Larranaga1, ...............................................................................................27 S. Vijendran1, D. Rebuffat1, MC. Perkinson2, F. Mura3 ..........................................................................................27

THE MISSION MIRIAM-‐2: PUTTING A GOSSAMER BALLUTE THROUGH AN ATMOSPHERIC ENTRY FLIGHT TEST...........................................................................................................................................................................29 H.S.Griebel1*, R.Foerstner2, C.Mundt2, J.Polkko3, H.Teodorescu4, G.Herdrich5, T.Marynowski5, A.Stamminger6.......................................................................................................................................................................29

VENUS DEEP ATMOSPHERE DESCENT PROBE (VDAP) .......................................................................................30 James B. Garvin, Lori Glaze, Paul Mahaffy, Natasha Johnson, Michael Amato, Tim Van Sant............30

VENUS PATHFINDER – A COMPACT LONG-‐LIVED LANDER MISSION ...........................................................32 Ralph Lorenz ..........................................................................................................................................................................32

TITAN AERIAL EXPLORER (TAE): EXPLORING TITAN BY BALLOON.............................................................34 Jeffery L. Hall1, Jonathan Lunine2, Christophe Sotin3, Kim Reh4, Andre Vargas5 and Patrice Couzin6......................................................................................................................................................................................................34

AN ADVANCED DESIGN FOR A TITAN BALLOON....................................................................................................36 Julian Nott1, Don Cameron2, Don Day3, Greg Mungas4.........................................................................................36

8 MISSION CONCEPT FOR ENTRY PROBES TO THE FOUR OUTER PLANETS BASED ON E-‐SAIL PROPULSION...........................................................................................................................................................................39 Jean-Pierre Lebreton1, Pekka Janhunen2, Sini Merikallio2, Petri Toivanen2...............................................39

SESSION 3 - SCIENCE FROM PROBES AND PENETRATORS...................................................................... 41 NEW TOOLS AND METHODS TO FULLY CHARACTERIZE THE ATMOSPHERIC ENVIRONMENT FOR A MARTIAN EDL APPLICATION TO THE 2016 EXOMARS DESCENT MODULE ..........................................42 F. Forget1, A. Spiga1, L. Montabone1, E. Millour1, A. Colaitis1, V. Bourrier1 F. Gonzalez-Galindo2 S. R. Lewis3, S. Portigliotti4. ........................................................................................................................................................42

ENTRY TRAJECTORY RECONSTRUCTION USING PHOENIX RADIO LINK ............................................................................44 Ö. Karatekin1 and S. W. Asmar2 ......................................................................................................................................44

AIRBORNE OBSERVATION OF THE HAYABUSA SAMPLE RETURN CAPSULE RE-‐ENTRY .....................45 Jay H. Grinstead1, Peter M. Jenniskens2, Alan M. Cassell3, Jim Albers2, Michael Winter4........................45

RADIATION MODELING FOR THE REENTRY OF THE HAYABUSA SAMPLE RETURN CAPSULE.........48 Michael W. Winter1, Ryan D. McDaniel2, Yih-Kanq Chen2, Yen Liu2, David Saunders3...........................48

GIANT PLANET FORMATION, SATURN AND URANUS ENTRY PROBES, AND THE DECADAL .............50 Sushil Atreya...........................................................................................................................................................................50

2012 DECADAL SURVEY GIANT PLANET ENTRY PROBE SCIENCE.................................................................52 Thomas R. Spilker(1), David H. Atkinson(2)..................................................................................................................52

OUTER PLANET DOPPLER WIND MEASUREMENTS .............................................................................................54 D.H. Atkinson(1), S.W. Asmar(2), T.R. Spilker(2) ..........................................................................................................54

TITAN AERIAL EXPLORER ................................................................................................................................................56 Jonathan I Lunine1, Christophe Sotin2 .........................................................................................................................56

SESSION 4 - EDL TECHNOLOGY DEVELOPMENT......................................................................................... 58 GOING BEYOND RIGID AEROSHELLS: ......................................................................................................................................59 ENABLING VENUS IN-‐SITU SCIENCE MISSIONS WITH DEPLOYABLES................................................................................59 Ethiraj Venkatapathy1, Todd White2, Gary Allen3, and Dinesh Prabhu4 ......................................................59

A COMPARISON OF INFLATABLE AND SEMI-‐RIGID DEPLOYABLE AERODYNAMIC DECELERATORS FOR FUTURE AEROCAPTURE AND ENTRY MISSIONS ..........................................................................................62 Reuben R. Rohrschneider, Jim Masciarelli, and Kevin L. Miller ........................................................................62

EXOMARS 2016 – GNC APPROACH FOR ENTRY DESCENT AND LANDING DEMONSTRATOR ............64 S. Portigliotti1, P.Martella1, M.Capuano1, O.Bayle2, T.Blancquaert2...............................................................64

THE MARS SCIENCE LABORATORY ENTRY DESCENT AND LANDING MODE COMMANDER..............66 Paul Brugarolas, Kim Gostelow, A. Miguel San Martin, Fred Serricchio, and Gurkipal Singh............66

SUPERSONIC RETRO-‐PROPULSION FLIGHT TEST CONCEPTS .........................................................................67 Ethan Post(1), Artem Dyakonov(2), Ashley Korzun(3), Ian Dupzyk(4), Jeremy Shidner(5), Arturo Casillas(6), Karl Edquist(7)..............................................................................................................................................67

MAXIMUM ATTAINABLE DRAG LIMITS FOR ATMOSPHERIC ENTRY VIA SUPERSONIC RETROPROPULSION............................................................................................................................................................69 No ̈el M. Bakhtian1, Michael J. Aftosmis2.....................................................................................................................69

ROTARY WING DECELERATOR USE ON TITAN .......................................................................................................70 Ted Steiner1, Larry Young2...............................................................................................................................................70

SMALL PROBE REENTRY INVESTIGATION FOR TPS ENGINEERING (SPRITE) (IPPW-‐8).....................72 Daniel M. Empey1, Kristina A. Skokova2, Parul Agrawal2, Gregory T. Swanson2, Dinesh K. Prabhu2, Keith H. Peterson2 and Ethiraj Venkatapathy3........................................................................................................72

THE DEVELOPMENT OF A CO2 TEST CAPABILITY IN THE NASA JSC ARCJET FOR MARS REENTRY SIMULATION...........................................................................................................................................................................74 Steven V. Del Papa1, Leonard Suess2, Brian Shafer3 ..............................................................................................74

SESSION 5 - SCIENCE INSTRUMENTATION................................................................................................... 76 PAYLOAD OPTIONS FOR FUTURE ENTRY PROBE MISSIONS ............................................................................77 Thomas R. Spilker.................................................................................................................................................................77

TITAN LAKE PROBE: SCIENCE REQUIREMENTS AND INSTRUMENTATION..............................................78 J. Hunter Waite1, Tim Brockwell1, John Elliott2, Patricia Beauchamp2.........................................................78

INSTRUMENTS FOR IN SITU TITAN MISSIONS ........................................................................................................79 Patricia M Beauchamp1, Jonathan Lunine2...............................................................................................................79

9 Athena Coustenis LESIA3, Peter Willis1, George Cody4, Kim R. Reh1 ...............................................................79

SPACECRAFT-‐TO-‐SPACECRAFT RADIO LINKS INSTRUMENTATION FOR PLANETARY GRAVITY, ATMOSPHERIC AND SURFACE SCIENCES ..................................................................................................................81 Sami W. Asmar.......................................................................................................................................................................81

THE MARS MICROPHONE 2016 EXPERIMENT ........................................................................................................82 D. Mimoun1, Jean-Pierre Lebreton2, and the Mars Microphone 2016 team3..............................................82

LIDAR INSTRUMENT FOR GLOBAL MEASUREMENT OF MARS ATMOSPHERE.........................................85 Farzin Amzajerdian1, George Busch2, Norman Barnes1, Robert Tolson3, and Diego Pierrottet2 ......85

MARTIAN SONIC ANEMOMETER...................................................................................................................................87 Don Banfield ...........................................................................................................................................................................87

THE CHEMCAM INSTRUMENT FOR THE 2011 MARS SCIENCE LABORATORY MISSION: SYSTEM REQUIREMENTS AND PERFORMANCE.......................................................................................................................88 R. Perez1, B.L. Barraclough2, S.C. Bender2, A. Cousin3, A. Cros3, N. Le Roch4, S. Maurice3, A. Paillet1, L. Pares3, Y. Parot3, M. Saccoccio1, R.C. Wiens2 .............................................................................................................88

MEADS CALIBRATION AND MSL TRAJECTORY RECONSTRUCTION ..............................................................90 Mark Schoenenberger1, Chris Karlgaard2, Michelle Munk1 ...............................................................................90

OPTICAL EMISSION SPECTROSCOPIC EXPERIMENTS FOR IN-‐FLIGHT ENTRY RESEARCH.................91 Sebastian Lein1, Georg Herdrich1, Monika Auweter-Kurtz2 and Stefanos Fasoulas1..............................91

SESSION 6A - NEW TECHNOLOGIES ................................................................................................................ 93 PEDALS: EVOLVED DESIGN OF EDL ARCHITECTURES ........................................................................................94 Ed Chester, João Graciano.................................................................................................................................................94

CHALLENGES OF THE INSTRUMENTATION FOR HIGH SPEED ENTRY VEHICLES...................................95 Ali Gülhan, Frank Siebe, Thomas Thiele .....................................................................................................................95

SYSTEM DEVELOPMENT FOR MARS ENTRY IN-‐SITU RESOURCE UTILIZATION......................................97 Svetozar Popovic1, Robert W. Moses2, Leposava Vuskovic3................................................................................97

TERMINAL DESCENT AND LANDING SYSTEM ARCHITECTURES FOR A MARS PRECISION LANDER.......................................................................................................................................................................................................99 Lisa Peacocke1, Marie-Claire Perkinson1, Jaime Reed1, Tobias Lutz2, Marco Wolf2, Joerg Boltz2 .....99

OVERVIEW OF HYPERSONIC INFLATABLE AERODYNAMIC DECELERATOR LARGE ARTICLE GROUND TEST CAMPAIGN....................................................................................................................................................................................................101 Alan M. Cassell, Gregory T. Swanson, R. Keith Johnson, Stephen J. Hughes, F. McNeil Cheatwood 101

LOW-‐DENSITY SUPERSONIC DECELERATOR SYSTEM .....................................................................................103 Mark Adler, Chuck Player, Juan Cruz, Ian Clark, Adam Steltzner, and Tom Rivellini.......................... 103

CO2 PROPULSION FOR A MARS SURFACE HOPPER............................................................................................104 Christopher Perry & Robert L. Ash ............................................................................................................................. 104

COMPUTATIONAL STUDY OF ROUGHNESS-‐INDUCED TRANSITION..........................................................105 Seokkwan Yoon, Michael D. Barnhardt and Emre Sozer ................................................................................. 105

THREE DIMENSIONAL RADIATION IN MARTIAN ATMOSPHERE ................................................................109 Daniil Andrienko1,2, Sergey Surzhikov2 .................................................................................................................... 109

SESSION 6B - AEROASSIST, EXPERIMENTAL MISSIONS AND EDL MISSION DESIGN.....................111 OVERVIEW OF THE NASA ENTRY, DESCENT AND LANDING SYSTEMS ANALYSIS EXPLORATION FEED-‐FORWARD STUDY.................................................................................................................................................112 Alicia D. Cianciolo1, Thomas A. Zang1, Ronald R. Sostaric2, M. Kathy Mcguire3..................................... 112

AEROFAST: MARTIAN AEROCAPTURE FOR FUTURE SPACE TRANSPORTATION – MISSION OVERVIEW............................................................................................................................................................................113 T. Salmon*1, F. Bonnefond1, J-M. Bouilly1, P. Augros2, T. Lutz3 ....................................................................... 113

MISSION ANALYSIS AND FLIGHT MECHANICS OF EARTH EXPERIMENTAL MISSIONS .....................115 Rodrigo Haya-Ramos1, Davide Bonetti1, Cristina Parigini1, Jorge Serna1, Gabriele de Zaiacomo1, Federico Massobrio2......................................................................................................................................................... 115

HAYABUSA REENTRY: TRAJECTORY ANALYSIS AND OBSERVATION MISSION DESIGN...................118 Alan M. Cassell1, Gary A. Allen1, Jay H. Grinstead1, Manny E. Antimisiaris2, Jim Albers3, Petrus M. Jenniskens3............................................................................................................................................................................ 118

A SIMPLE ANALYTICAL EQUATION TO ACCURATELY CALCULATE THE ATMOSPHERIC DRAG DURING AEROBRAKING CAMPAIGNS VALIDATION IN THE MARTIAN CASE .........................................120 F. Forget, M. Capderou .................................................................................................................................................... 120

10 AEROBRAKING PERIAPSIS CONTROL STRATEGIES...........................................................................................122 M. Sánchez*, F. Cichocki.................................................................................................................................................. 122

PLANNED FLIGHT OF THE INFLATABLE REENTRY VEHICLE EXPERIMENT 3 (IRVE-‐3) ...................124 Robert A Dillman, F Neil Cheatwood, Stephen J Hughes, Joseph Del Corso, Richard J Bodkin, and Aaron Olds ............................................................................................................................................................................ 124

DIMENSIONLESS PARAMETERS FOR ESTIMATING MASS OF INFLATABLE AERODYNAMIC DECELERATORS .................................................................................................................................................................125 Jamshid A. Samareh.......................................................................................................................................................... 125

SESSION 7A - ADVANCES IN TPS TECHNOLOGY FOR PLANETARY PROBE DESIGN.......................126 CHALLENGES WITH THERMAL PROTECTION MATERIAL DEVELOPMENT AND IMPLEMENTATION: LESSONS LEARNED FROM RECENT NASA EXPERIENCE.......................................127 D. Ellerby ............................................................................................................................................................................... 127

ONGOING EUROPEAN DEVELOPMENTS ON ENTRY HEATSHIELDS AND TPS MATERIALS .............128 H. Ritter1, O. Bayle1, Y. Mignon2, P. Portela3, J-M. Bouilly2, R. Sharda4 ....................................................... 128

MEDLI AEROTHERMAL ENVIRONMENT RECONSTRUCTION EFFORTS ..........................................................................129 Todd White ........................................................................................................................................................................... 129

ORION FLIGHT TEST-‐1 THERMAL PROTECTION SYSTEM INSTRUMENTATION..................................130 T. John Kowal1..................................................................................................................................................................... 130

FLEXIBLE ABLATORS: APPLICATIONS AND ARCJET TESTING......................................................................132 James O. Arnold1, Ethiraj Venkatapathy1, Robin Beck1, Kathy M. McGuire1, Dinesh K. Prabhu2 and Sergey Gorbunov3 .............................................................................................................................................................. 132

OVERVIEW OF INITIAL DEVELOPMENT OF FLEXIBLE ABLATORS FOR MARS EDL .......................................................134 Robin A.S. Beck, Susan White, James Arnold, Wenhong Fan, Mairead Stackpoole, Parul Agrawal................................................................................................................................................................................................... 134

AEROFAST: DEVELOPMENT OF CORK TPS MATERIAL AND A 3D COMPARATIVE THERMAL/ABLATION ANALYSIS OF AN APOLLO & A BICONIC SLED SHAPE FOR AN AEROCAPTURE MISSION................................................................................................................................................136 G. Pinaud1 & A.J. van Eekelen2 ...................................................................................................................................... 136

MODULAR MANUFACTURING OF HONEYCOMB-‐REINFORCED CHARRING ABLATOR SYSTEMS FOR THE AEROSHELLS OF LARGE EDL VEHICLES........................................................................................................138 William M. Congdon ......................................................................................................................................................... 138

SESSION 7B - AIRLESS BODY SURFACE MISSIONS....................................................................................139 ROBOTIC AND HUMAN SPACE EXPLORATION OF NEAR-EARTH OBJECTS ............................................ 140 D.D. Mazanek....................................................................................................................................................................... 140

EUROPEAN GNC TECHNOLOGY DEVELOPMENT AND PERSPECTIVE FOR AIRLESS BODIES EXPLORATION ....................................................................................................................................................................141 A. CARAMAGNO .......................................................................................................................................................................141 MAGIC (MOBILE AUTONOMOUS GENERALIZED INSTRUMENT CARRIER) ......................................................................142 T. van Zoest1, T.-M. Ho1, C. Lange1, L. Witte1, S. Wagenbach1, C. Krause1, S. Ulamec1, ........................ 142 J. Biele1, Florian Herrmann1, Joachim Block1, and Pierre Bousquet2 .......................................................... 142 1 DLR – Deutsches Zentrum f. Luft- und Raumfahrt, Germany ..................................................................... 142 2 CNES – Centre National d'Études Spatiales, Toulouse, France.................................................................. 142

THE ESA LUNAR LANDER MISSION ...........................................................................................................................143 CAMERA-‐AIDED INERTIAL NAVIGATION FOR PINPOINT PLANETARY LANDING ON RUGGED TERRAINS..............................................................................................................................................................................144 Jeff Delaune1, Guy Le Besnerais1, Martial Sanfourche3, Aurélien Plyer4, Jean-Loup Farges5, Clément Bourdarias6, Thomas Voirin7 and Alain Piquereau8 .......................................................................................... 144

MARCO POLO-‐R: AN ASTEROID SAMPLE RETURN MISSION..........................................................................146 Mark Adler, Andy Cheng, Tom Randolph, and Rob Maddock......................................................................... 146

WHAT MOONRISE LUNAR SAMPLE RETURN CAN TEACH US ABOUT MARS SAMPLE RETURN.....147 George Chen1, Eric Blood2 .............................................................................................................................................. 147

FARSIDE EXPLORER: UNIQUE SCIENCE FROM A MISSION TO THE FARSIDE OF THE MOON..........148 David Mimoun1, Mark Wieczorek2, and the Farside Explorer Team3......................................................... 148

VLBI TRACKING OF PHOBOS-‐GRUNT PROBE........................................................................................................151 Guifré Molera Calvés1,2, S.V. Pogrebenko3, G. Cimò3, D.A. Duev3,4, L.I. Gurvits3,5 ..................................... 151

11 SESSION 8 - CLOSING .........................................................................................................................................153 AUSTERITY IN THE AGE OF INNOVATION..............................................................................................................154 Bethany Johns...................................................................................................................................................................... 154

NASA-‐LANGLEY RESEARCH CENTER’S ENGINEERING DIRECTORATE.....................................................155 Stephen P. Sandford.......................................................................................................................................................... 155

POSTERS................................................................................................................................................................156 POSTER SESSION 2 – PROBE MISSIONS.......................................................................................................157 STUDY OF PLANETARY ENTRY PROBES (PEP) FOR VENUS AND OUTER PLANETS: SATURN, URANUS AND NEPTUNE .................................................................................................................................................158 Denis Rebuffat, Peter Falkner, Jonan Larranaga, Jens Romstedt, Kelly Geelen...................................... 158

POSTER SESSION 3 – SCIENCE FROM PROBES AND PENETRATORS ..................................................160 ACCOMMODATION STUDY FOR AN ANEMOMETER ON A MARTIAN LANDER.......................................161 Benjamin Lenoir1, Don Banfield1 ................................................................................................................................ 161

THE MARS CLIMATE DATABASE, CURRENT STATUS AND FUTURE IMPROVEMENTS ......................163 E. Millour(1), F. Forget(1), A. Spiga(1), S. Lebonnois(1), S.R. Lewis(2), L. Montabone(3), P.L. Read(3),M.A. López-Valverde(4), F. González-Galindo(4), F. Lefèvre(5), F. Montmessin(5), M.-C. Desjean(6), J.-P. Huot(7) and the MCD/GCM development team ................................................................. 163

ARMADILLO – A DEMONSTRATION FOR LOW-‐COST IN-‐SITU INVESTIGATIONS OF THE UPPER ATMOSPHERE OF PLANETARY BODIES ..................................................................................................................167 Rene Laufer(1,3), Glenn Lightsey(2), Georg Herdrich(3,1), Ralf Srama(3,4,1), Gregory Earle(5), Carsten Wiedemann(6), Ed Chester(7), Hugh Hill(8), Troy Henderson(9), Rainer Sandau(10,11,1), Lorin Matthews(1), Truell Hyde(1) ........................................................................................................................... 167

POSTER SESSION 4 – EDL TECHNOLOGY DEVELOPMENT.....................................................................169 ONGOING VALIDATION OF COMPUTATIONAL FLUID DYNAMICS FOR SUPERSONIC RETRO-‐PROPULSION........................................................................................................................................................................170 Daniel G. Schauerhamer,* Kerry A. Trumble†, William Kleb†, Jan-Renee Carlson§, Pieter G. Buning,** Karl Edquist††, and Emre Sozer‡‡ ...................................................................................................... 170

ENTRY AND POWERED DESCENT GUIDANCE FOR MARS ROBOTIC PRECURSORS .............................173 Sostaric, Ronald R.; Garcia-Llama, E. Powell, R. W............................................................................................. 173

MULTI-‐MISSION EARTH ENTRY VEHICLE DESIGN TRADE SPACE AND CONCEPT DEVELOPMENT STATUS (VERSION 2.0)....................................................................................................................................................175 Robert W. Maddock .......................................................................................................................................................... 175

THERMAL SOAK ANALYSIS OF SPRITE PROBE.....................................................................................................176 P. Agrawal1, Y.K. Chen2, D.K. Prabhu1 D. Empey3, E. Venkatapathy2, J. Arnold2 ..................................... 176

DESIGN CHOICE CONSIDERATIONS FOR VEHICLES UTILIZING SUPERSONIC RETROPROPULSION....................................................................................................................................................................................................178 Ashley M. Korzun(1), Ian G. Clark(2), Robert D. Braun(3) .............................................................................. 178

POSTER SESSION 5 – SCIENCE INSTRUMENTATION ...............................................................................181 THE STUDENT RAINDROP DETECTOR (SRD): AN INSTRUMENT FOR MEASURING METHANE RAIN ON TITAN...............................................................................................................................................................................182 Allison Tucker1, Gabriel Wilson1, Hieu Truong1, Tim Kunz1, Kysen Palmer1, Colton Therrian1, Jason W. Barnes1, David H. Atkinson1................................................................................................................................... 182 Ralph D. Lorenz2 ................................................................................................................................................................ 182

PLANETARY POLARIZATION NEPHELOMETER ..................................................................................................183 Don Banfield(1), Adam Saltzman(1) ........................................................................................................................ 183

SCIENCE AND EDUCATION WITH MARS EXPRESS' VISUAL MONITORING CAMERA ..........................185 (VMC) ......................................................................................................................................................................................185 H.S.Griebel*1, T.Ormston1, M.Denis2, J.Landeau-Constantin2, D.Scouka2,3, L.Griebel4, C.Scorza5, M.Frommelt5 ........................................................................................................................................................................ 185

DEVELOPMENT OF INSTRUMENTATION FOR HYPERSONIC INFLATABLE AERODYNAMIC DECELERATOR CHARACTERIZATION......................................................................................................................187 Gregory T. Swanson1,2, Alan M. Cassell 2 .................................................................................................................. 187

12 MARS MICROPHONE 2016: A UNIQUE OPPORTUNITY FOR STUDENT INVOLVEMENT.....................189 *A. Minier1, *W. Rapin1, D. Perez Escobar1, D. Mimoun1 and the Mars Microphone Team 2 ............ 189

POSTER SESSION 6A – NEW TECHNOLOGIES ............................................................................................191 TDNR: A MODULAR NANO-‐ROVER PLATFORM FOR NETWORKED PLANETARY MISSIONS ...........192 Abraham Rademacher (1), Amardeep Singh(2), Jasvir Singh(3), Jose Cortez (4), Kavinda Wittahachchi(5), Mikhail Paremski(6), Yawo Ezunkpe(7), Dr. Periklis Papadopoulos(8), Marcus S. Murbach(9), Bob Feretich(10)..................................................................................................................................... 192

ANALYSIS OF ANOMALOUS VARIATIONS IN HIGH ALTITUDE BALLOON ASCENT RATES NEAR THE TROPOPAUSE ......................................................................................................................................................................194 Walter Taresh*, Kevin Ramus, Kim Baird, Carlos Gonzalez, Gabe Wilson, Rory Riggs, George Korbel, David H. Atkinson, and the Idaho Near Space Engineering Team............................................... 194

DEVELOPMENT OF AN AUTONOMOUS HIGH ALTITUDE BALLOON CUTDOWN SYSTEM.................195 Kevin Ramus*, Kim Baird, Carlos Gonzalez, Gabe Wilson, Walter Taresh, Rory Riggs, George Korbel, David H. Atkinson, and the Idaho Near Space Engineering Team............................................... 195

THE TITAN SKY SIMULATORTM NEW LOW COST CRYOGENIC TEST FACILITY AVAILABLE.............196 DEVELOPED FOR TITAN BALLOONS BUT SUITABLE FOR MANY APPLICATIONS ................................196 J. Nott ...................................................................................................................................................................................... 196

ONE-‐WAY UPLINK RANGING FOR ENHANCING PLANETARY WIND MEASUREMENTS.....................198 K. Oudrhiri(1), D.H. Atkinson(2), S.W. Asmar(1), S. Bryant(1), T.R. Spilker(1)...................................... 198

POSTER SESSION 6B - AEROASSIST, EXPERIMENTAL MISSIONS AND EDL MISSION DESIGN....199 SATURN SYSTEM MISSION OPPORTUNITIES USING A TITAN AEROGRAVITY ASSIST FOR ORBITAL CAPTURE ...............................................................................................................................................................................200 Robert M. Booher(1) and J.E. Lyne(2)....................................................................................................................... 200

AERODYNAMIC STABILITY OF BLUNTED-‐CONE ENTRY VEHICLES ...........................................................203 Daniel R. Ladiges*, Eleanor C. Button, Charles R. Lilley, Nicholas S. Mackenzie, Edward Ross, And John E. Sader........................................................................................................................................................................ 203

DETERMINATION OF AERODYNAMIC DAMPING COEFFICIENTS OF ENTRY VEHICLES IN TRANSONIC REGIME ........................................................................................................................................................205 S. Paris, O. Karatekinn, A. Karitonov+, J. Ouvrard* ............................................................................................. 205

STATISTICAL ENTRY, DESCENT AND LANDING PERFORMANCE RECONSTRUCTION OF THE MARS PHOENIX LANDER.............................................................................................................................................................207 Soumyo Dutta(1), Ian G. Clark(2), Ryan P. Russell(3), Robert D. Braun(4)............................................. 207

VERTICAL STRUCTURE AND WIND SHEAR IN A SIMULATED TRITON ATMOSPHERE ......................209 Charles Miller(1), Nancy J. Chanover(1), James R. Murphy(1) ...................................................................... 209

POSTER SESSION 7A - ADVANCES IN TPS TECHNOLOGY FOR PLANETARY PROBE DESIGN......211 DEVELOPMENT OF A THERMAL PROTECTION SYSTEM MASS ESTIMATING RELATIONSHIP BASED ON FIAT PREDICTIONS....................................................................................................................................................212 S. Sepka1, J. O. Arnold2, E. Venkatapathy3 and K. Trumble4............................................................................. 212

RASTAS SPEAR : RADIATION-‐SHAPES-‐THERMAL PROTECTION INVESTIGATIONS FOR HIGH SPEED EARTH RE-‐ENTRY ..............................................................................................................................................213 J-M Bouilly1, A. Pisseloup1, O. Chazot2, G. Vekinis3, A. Bourgoing4, B. Chanetz5, O. Sladek6 ............... 213

RESIN IMPREGNATED CARBON ABLATOR (RICA): A NEW THERMAL PROTECTION SYSTEM MATERIAL FOR HIGH-‐SPEED PLANETARY ENTRY VEHICLES ......................................................................215 Jaime Esper (1), Hans-Peter Roeser (2), Georg Herdrich (2) ......................................................................... 215

PERFORMANCE CHARACTERIZATION, SENSITIVITY AND COMPARISON OF A DUAL LAYER THERMAL PROTECTION SYSTEM...............................................................................................................................217 Cole D. Kazemba(1), Mary Kathleen McGuire(2), Austin Howard(3), Ian G. Clark(4), Robert D. Braun(5)................................................................................................................................................................................ 217

EDL HEATSHIELD EXPERIMENTS WITH DUAL-‐LAYER ABLATORS, ADVANCED MATERIALS AND VARIABLE HONEYCOMBS..............................................................................................................................................218 Jennifer N. Congdon.......................................................................................................................................................... 218

LOW DENSITY FLEXIBLE CARBON PHENOLIC ABLATORS .............................................................................219 Mairead Stackpoole1, Jeremy Thornton1, Wendy Fan1 and Parul Agrawal1, Evan Doxtad2, Robin Beck3........................................................................................................................................................................................ 219

13 ROTATING ARC JET TEST MODEL: TIME-‐ACCURATE TRAJECTORY HEAT FLUX REPLICATION IN A GROUND TEST ENVIRONMENT...................................................................................................................................220 Bernard Laub and Jay Grinstead1, Artem Dyakonov2, Ethiraj Venkatapathy1 ....................................... 220

ADVANCED RIGID ABLATIVE TPS ..............................................................................................................................223 Matt Gasch............................................................................................................................................................................ 223

MODELING OF THE MATERIAL RESPONSE OF THERMAL PROTECTION SYSTEMS IN HYPERSONIC FLOWS ....................................................................................................................................................................................224 Jonathan Wiebenga(1), Iain D. Boyd(2), Alexandre Martin(2)..................................................................... 224

14

ABSTRACTS

15

Session 1 - Outlook for Probe Missions

16

2011 AL SEIFF AWARD LECTURE

THE HUYGENS STORY

Jean-Pierre Lebreton

ESA/ESTEC, SolarSystem Missions Division, Noordwijk, The Netherlands. [email protected],

ABSTRACT

In this lecture, I will tell the story of (Cassini-)Huygens with emphasis on the role played by the young scientists, engineers and students I directly worked with in the Planetary Missions Division, and later the Solar System Missions Division at ESA/ESTEC. I will illustrate how their work contributed to i) getting Huygens ready for its historical descent in Titanʼs atmosphere on January 14 2005, ii) the post-flight analysis and interpretation of Huygens data and iii) Huygens legacy.

17

PLANETARY PROBES AND THE PLANETARY DECADAL SURVEY

Amy Simon-Miller

NASA Goddard Space Flight Center, [email protected]

ABSTRACT

On March 7, 2011, the National Research Council released Vision and Voyages for Planetary Science in the Decade 2013-2022. This report represents broad community input via white paper submissions, presentations to panels and the Steering Group, and participation in mission concept studies. Careful independent assessment of mission concepts was used to validate cost, technical feasibility, and risk, prior to ranking each mission by the discipline panels. The final priorities of large and medium class missions, as well as technology development, research and analysis, and small missions were made across disciplines based on programmatic balance, costs and science value. The recommended missions include opportunities for atmospheric probes to Venus, Saturn and Uranus, as well as the study of future probe technology. This presentation will give a brief overview of the survey process and recommendations, with specific emphasis on planetary probe opportunities.

18

25 YEARS OF DEEP SPACE EXPLORATION AT ESA

Marcello Coradini

ESA Programs Coordinator at JPL, [email protected]

ABSTRACT

The European adventure in the deep space started in 1985 with the launch of the GIOTTO mission to comet Halley. This first, rather simple, deep space probe was followed in the course of the years, by a large number of satellites and probes directed almost in any corner of the Solar System. From Mercury, with the BepiColombo composite spacecraft, to Venus with Venus Express, the Moon with SMART-1, the red planet with Mars Express, the Saturn System with Cassini/Huygens, Comet 67P/Churyumov-Gerasimenko and the asteroids Steins and Lutetia with the Rosetta, ESA and the European scientists are present and actively carrying out scientific observations. In the near future the Robotic Exploration Programme, defined in a close collaboration with NASA, will allow ESA to be again in Mars orbit with the TGO satellite and on the surface of the red planet with the ExoMars rover, while a new interplanetary probe might be travelling to Jupiter to explore its satellite system. In about 25 years ESA, the European Scientists and the European Industry have reached a high level of competitiveness, competence and experience. New worlds have been explored, a wealth of new technologies and systems have been developed. Exploration of the Solar System with robots, and one day with humans, will continue to offer challenges and excitement for many years to come.

19

NASA INVESTMENTS IN OUR FUTURE: EXPLORING SPACE THROUGH INNOVATION AND TECHNOLOGY

Dr. Michael Gazarik

Deputy Chief Technologist at NASA, [email protected]

ABSTRACT NASA Deputy Chief Technologist Dr. Michael Gazarik will provide an overview of NASA's planned research, innovation and technology investments that focus on enabling bold robotic and human exploration of the solar system while providing broadly-applicable benefits on Earth. Dr. Gazarik will highlight NASA's push for disruptive technologies that may enable exploration of deep space, including near-Earth asteroids and eventually Mars. This technology-enabled exploration strategy will allow NASA to explore beyond low Earth orbit more efficiently, safely, and expeditiously. Central to this approach, NASA's new Space Technology Program seeks to create the technological knowledge and capability needed to enable a new generation of NASA aeronautics, science, and exploration missions. By taking informed risks and focusing on high-payoff technologies, the Space Technology Program will provide the answers to the Agency's future technological needs. Developing these new transformational technologies and capabilities will require the best of academia, industry, and our government labs. Dr. Gazarik will also highlight NASA’s technology development roadmaps, as provided to the National Research Council (NRC), that describes the research and technology development investments required for tomorrow’s great discoveries.

20

ESA EXPLORATION PROGRAMMES FROM ISS TO THE LUNAR LANDER MISSION

Bruno Gardini

European Space Agency – ESA-ESTEC, Noordwijk, The Netherlands, [email protected]

ABSTRACT

The extension of the ISS operation to 2020 is providing new opportunities for Exploration preparatory activities in a representative environment in the field of human spaceflight. To this end ESA has recently issued a call for ideas and is preparing new activities to be implemented in the near future. In the same time industrial activities to design and develop the first ESA Lunar Lander continue to progress at a fast pace and with an increasing support of ESA Members states, setting the ground for a full development proposal being presented for approval at the next ESA Council at Ministerial level in 2012. While providing a good opportunity for scientific experiments on the surface of the Moon, the Lunar Lander’s primary goal is to develop precision landing technology. Mandated by the requirement to land on a rough terrain at the Moon South Pole the mission will develop for Europe the new generation of guidance, navigation and control sensors, algorithms and software including visual navigation and hazard avoidance. As such the technology can be applied to Moon and asteroid landing as well as to the terminal phase of a Mars landing mission. With the prospective of human presence being the ultimate goal of Exploration, the presentation will include an overview of the ESA present and planned activities in the area of human spaceflight.

21

PROGRESS TOWARD A COMPLETE RESPONSE TO THE PLANETARY DECADAL SURVEY

Jim Adams

NASA Headquarters, Planetary Science Division, Washington, DC USA, [email protected]

ABSTRACT

March 7, 2011, the National Research Council released Vision and Voyages for Planetary Science in the Decade 2013-2022. This report represents broad community input via white paper submissions, presentations to panels and the Steering Group, and participation in mission concept studies. The NASA Headquarters Planetary Science Division is in the process of preparing a response to the 200+ recommendations contained the Decadal Survey. Just two weeks earlier the President’s fiscal year 2012 budget request was released. This presentation will give a brief overview of the response process, the realities and limitations of the 2012 budget and progress toward completing a reconciled response.

Session 2 - Probe Missions

23

MARS SCIENCE LABORATORY ENTRY, DESCENT AND LANDING SYSTEM DESIGN, DEVELOPMENT AND

PRELAUNCH STATUS

Adam Steltzner

NASA JPL, [email protected]

ABSTRACT

In late November of 2011 the Mars Science Laboratory mission will launch from Kennedy Space Center in Florida headed for Mars. This mission will deliver the largest ever extraterrestrial lander. Further this landed system is a mobile rover and the MSL entry descent and landing system delivers it on its wheels ready for commissioning. This MSL EDL system provides more performance, in mass delivery, altitude capability, landed accuracy than any Mars EDL system before it. This paper will document the MSL EDL design and development and its status for launch readiness.

24

EXOMARS EDM MISSION AND DESIGN OVERVIEW

Olivier Bayle*, Leila Lorenzoni*, Thierry Blancquaert*, Stephane Langlois*, Thomas Walloschek*, S. Portigliotti§, M. Capuano§

* European Space Agency ESTEC, Noordwijk [email protected]

[email protected], [email protected], [email protected], [email protected]

§Thales Alenia Space Italy Turin, Italy [email protected], [email protected]

ABSTRACT

The ExoMars Programme is a joint ESA-NASA initiative to explore Mars and will make use of the 2016 and 2018 launch opportunities. Among the ExoMars objectives, the 2016 mission shall provide a demonstration of key technologies required to safely land a payload on the surface of Mars: -‐ Heat Shield -‐ Parachute System -‐ Doppler Radar System for ground relative altitude and relative velocity

measurement -‐ Liquid Propulsion System for attitude control and final braking -‐ Crushable material for impact loads attenuation These technologies will be embarked into the ExoMars EDL Demonstrator Module (EDM), a 600 kg 2.4 m diameter vehicle. In order to maximise the lessons learnt from this demonstration mission, the EDM will carry a package of sensors that will allow a detailed reconstruction of the flown trajectory as well as the assessment of the EDL subsystems performance. The paper provides an overview of the EDM mission and design and describes the early test activities that have already been carried out in order to raise the technology readiness level of the key EDL technologies (aerothermodynamics, TPS, Parachute system, Doppler radar, propulsion, crushable structure).

25

END TO END MISSION PERFORMANCES OF EXOMARS 2016 EDM

Rodrigo Haya-Ramos1 Mariano Sanchez Nogales1, Juan Luis Cano1 David Riley2, David Northey2, Stefano Portigliotti3, Olivier Bayle4

DEIMOS Space S.L.U1, [email protected]

Tessella plc2, [email protected] Thales Alenia Space3, Stefano.portigliotti@ thalesaleniaspace.com

European Space Agency4, [email protected]

ABSTRACT

The EXOMARS programme foresees two missions: the first, to be launched in 2016, consisting of an Orbiter plus an Entry, Descent and Landing Demonstrator (EDM) and the second, with a launch date in 2018, consisting of two rovers. Both missions will be carried out in cooperation with NASA. This scenario is the result of a 4 year evolution during the Phase B of the Exomars programme. The objective of this paper is to present the Mission Performances (dispersion analysis) of the Exomars 2016 Mission from launch to splashdown for the Baseline presented at the System PDR with focus on the EDM element. The present 2016 mission baseline is based on launch with Atlas V (421) in 2016 of a spacecraft Composite bearing a Orbiter Module and the EDM which is directed towards Mars through a direct type T2 transfer orbit, which includes a Deep Space Manoeuvre (DSM). The EDM is released from the arrival hyperbola 3 days before reaching Mars atmosphere. The EDM performs a ballistic entry and deploys a single stage Disk-gap-band parachute at Mach 1.95. After 40 s the frontshield is jettisoned and the rest of the EDM continues descent while radar is activated. The backshell is separated and the lander performs a powered landing with a g-turn manoeuvre to cancel the vertical velocity at 2 m above ground, where the retrorockets are switched off and the surface platform lands using its crushable structure on the Meridiani region. One of the objectives of the Mission Design and Analysis activity has been the consideration of a continuous end to end profile from launcher injection to touchdown in order to couple the arrival with the EDL phases since the first steps of the mission design. This coupling is relevant in several areas; among others: arrival epoch and local time with Mars environment, arrival hyperbola with reachability of landing latitudes, entry orientation (pro/retrograde) with aerothermodynamics and parachute deployment conditions, mapping of Navigation uncertainties and manoeuvres dispersions into landing accuracy and Entry Corridor size… This end-to-end philosophy has been applied in all of the mission design steps and in particular in the Mission performances evaluation. A single continuous mission timeline and trajectory from launcher separation to touchdown has been built as reference for the

26 assessment. This evaluation of the performances has been carried out through high fidelity simulation of the mission phases and events. Different levels of simulations have been planned and executed depending on the performances to be assessed. All of them are based on large Monte Carlo campaigns where initial states, atmosphere, mechanisms and vehicle characteristics are perturbed using high fidelity models for the environment and suitable performance models for the different elements and functions (separation mechanism, GNC…). First, the 3 DoF Monte Carlo campaign provides the overall mission performances with end-2-end simulations from the DSM to the touchdown where the compliance of the entry constraints (heat flux, load factor..), descent constraints (verticalisation, terminal velocities, inflation loads…) and landing constraints (impact velocity, consumed fuel…) are assessed in a multiphase process. This is the reference Mission Performance simulation which is complemented and tuned with more detailed performance assessments. Thus, 6 DoF simulations from the EDM separation down to the parachute triggering are carried out to assess the vehicle attitude dynamics and coupling with trajectory performances. Multibody simulations of the EDM under the parachute as a continuation of the perturbed 6DoF entry trajectories are executed for a detailed assessment of the dynamics under parachutes and the impact of the frontshield jettison on the EDM dynamics. The paper will present a summary of the mission design status, the selection of the reference Exomars Mission timeline and the results and discussion of the different campaigns (3DoF, 6DoF, Multibody) against the Mission and System Requirements. These results constitute the reference Mission performances for the Exomars Mission. A comparison of these performances with the Entry Corridor predictions will be presented and discussed. The s/w environment used for this assessment is the Endoatmospheric Simulator (EndoSim) of the Planetary Entry Toolbox and the Parachute System Design and Analysis Tool (PASDA), which represent the reference and validation sources for the official Exomars simulator.

27

FUTURE MISSIONS AND TECHNOLOGIES WITHIN THE MARS ROBOTIC EXPLORATION PREPARATION

(MREP) PROGRAMME

K. Geelen1, D. Agnolon1, P. Falkner1, J. Larranaga1,

S. Vijendran1, D. Rebuffat1, MC. Perkinson2, F. Mura3

European Space Agency1, ESTEC, Noordwijk The Netherlands,

[email protected] Astrium Limited, Stevenage, United Kingdom Thales Alenia Space Italia3, Torino, Italy

ABSTRACT

The European Mars Robotic Exploration Preparation (MREP) programme has the general approach to consider a Mars Sample Return mission in collaboration with NASA as a long-term objective and to progress step by step towards this mission through short and medium term technology developments. In parallel, long term generic enabling technologies are being developed with respect to propulsion and nuclear power systems. Intermediate missions would validate these technologies wherever possible. The 2018 joint NASA-ESA mission includes a sampling and caching rover, which will prepare cached samples to be retrieved and returned by an MSR lander mission in the early 2020’s. As such, the 2018 mission can be considered as the first component of the joint ESA/NASA MSR mission. In addition to this first step, MSR will include at least three main elements:

• A NASA-led MSR Lander (delivered to the Mars surface via the sky-crane concept), including a sample fetching rover which will retrieve the cached sample and transfer it into a sample container and a Mars Ascent Vehicle which will insert the sample container into Mars orbit,

• An ESA-led MSR Orbiter which will capture the sample container into Mars orbit and insert it into the Earth re-entry capsule which is brought back to Earth,

• A sample receiving facility as a key ground component. A preliminary scheme and schedule of the ESA and NASA shares for these building blocks and their components is presented here together with a preliminary design of the ESA undertakings. In order to maintain the robustness of the programme, ESA currently foresees four mission candidates for the post-ExoMars launch slots (2020/2022). The candidate missions currently being considered are: A. Mars network science mission, possibly including a high precision landing

28 demonstration, B. Sample return from a moon of Mars (Deimos or Phobos), C. Mars Precision lander (< ~10 km) with sampling/fetching rover, as a self- standing mission or a possible extra MSR segment if the NASA-led MSR lander cannot accommodate the Mars Ascent Vehicle and a Sample Fetching Rover, D. Mars Sample Return orbiter. Missions A. to C. are scientifically rewarding alternatives to cope with possible MSR delays, while mission D., and possibly mission C., may become MSR segments under Europe lead. Parallel phase 0/A studies are ongoing for the latter two missions whereas missions A. and B. have already been subject to system assessment studies in the past, which will be consolidated in 2011. These missions require a wide range of enabling technologies, for which development is ongoing within the MREP programme, such as:

• Mars Entry, Descent and Landing of small or medium-sized landers: o Improved navigation prior to Mars atmospheric entry o Guided entry to compensate known dispersions at entry and minimise

errors introduced by atmospheric uncertainty o Smart parachute deployment triggers o Hazard avoidance system: lidar and/or camera-based o Different landing systems such as legs and airbags

• Sampling, fetching and sample transfer techniques, • Precision landing on low-gravity bodies, • High-speed Earth re-entry, including thermal protection system and

aerothermodynamics, etc. • Autonomous rendezvous and capture in Mars orbit, including GNC, capture

mechanisms, etc. • Planetary protection, including bio-sealing, monitoring, etc.

The ongoing systems studies and technology development relating to the ESA MREP candidates missions are presented here and will help prepare the required inputs for the next Ministerial Council for enabling the down-selection of two of these missions for further definition phase (Phase B1). A decision on the implementation of MSR, i.e. the MSR orbiter, lander and sample receiving facility, should be taken at the Ministerial Council in 2015, together with NASA.

29

THE MISSION MIRIAM-2: PUTTING A GOSSAMER BALLUTE THROUGH AN ATMOSPHERIC ENTRY

FLIGHT TEST

H.S.Griebel1*, R.Foerstner2, C.Mundt2, J.Polkko3, H.Teodorescu4, G.Herdrich5, T.Marynowski5, A.Stamminger6

1Mars Society Deutschland e.V. & VEGA Space GmbH, 2Universitaet der Bundeswehr Muenchen, 3Finnish Meteorological Institute, 4Technical University of Iasi, 5IRS TU

Stuttgart, 6DLR Mobile Rocket Base Oberpfaffenhofen e-mail: [email protected], [email protected] [email protected],

[email protected], [email protected] [email protected], [email protected]

ABSTRACT

MIRIAM, short for ‘Main Inflated Re-entry Into the Atmosphere Mission test’, is a validation concept designed for the Mars ballute technology development programme ARCHIMEDES. This development programme is a joint effort of the Mars Society Germany and the University of the Federal Armed Forces of Germany in Munich, with further support by research institutes throughout Europe, the DLR and several industrial companies. The scientific objective of ARCHIMEDES is to obtain measurements of the Martian atmosphere, magnetic environment and surface throughout almost the entire altitude range reaching from outer space to ground. This is facilitated by an instrument carrier attached to a large and gossamer thin film ballute. MIRIAM was designed to validate the theory behind such a vehicle, test the newly developed technology in flight and to gain experience related to manufacturing, handling, flight operations, and gathering new scientific data on Mars and its atmosphere. A first MIRIAM test was flown in October 2008 from ESRANGE. In this paper we focus on the second MIRIAM test, named MIRIAM-2, which is currently slated for launch in October 2014 on a two-stage Taurus-Improved Orion rocket, again from ESRANGE. The main objective of MIRIAM-2 is to obtain realistic flight data for the re-entry of such a low ballistic coefficient design. Aerothermodynamic studies based on different methods have been performed for the entry into both Mars and Earth atmospheres, and flight data is needed to validate those methods. Presented will be the underlying theory, the mission and spacecraft design, the scientific experiments aboard, and expected improvements of the current body of knowledge. We will conclude with an outlook on further development, particularly the transfer of MIRIAM-2 results to a more accurate description of the atmospheric entry on Mars.

30

VENUS DEEP ATMOSPHERE DESCENT PROBE (VDAP)

James B. Garvin, Lori Glaze, Paul Mahaffy, Natasha Johnson, Michael Amato, Tim Van Sant

NASA Goddard, Email: [email protected]

ABSTRACT

The deep atmosphere of Venus remains largely unexplored in terms of the details of its trace gas chemistry. The history of key volatile reservoirs and surface-atmosphere-interior exchange processes is also poorly established on the basis of existing data. Noble gases within the bulk Venus atmosphere, as well as isotopes of hydrogen, oxygen, and sulfur, are essentially unmeasured to the degree required to address fundamental questions about the evolution of the planet, as prioritized in the latest planetary Decadal Survey by the US National Academy of Sciences. For these reasons, we have developed a mission concept for a Venus deep atmosphere descent probe (VDAP) that leverages existing and emergent technologies associated with entry-descent-touchdown, instrumentation, and flight system avionics. The intent is to define a low risk, cost-effective mission concept for achieving the recently outlined priorities for Venus atmospheric chemistry and dynamics, many of which date back to the early 1980’s in the aftermath of the US Pioneer Venus mission (PV) and contemporaneous Soviet Venera landers. Science objectives for the VDAP concept require state-of-the-art neutral mass spectrometer capabilities to achieve seminal measurements of noble gas isotopes including xenon (Xe), while also allowing for high mass resolution and time rate sampling of trace gases from beneath the Venus cloud deck to the surface. Instruments recently developed at NASA’s Goddard Space flight Center are well-suited to achieve these pivotal observations. Key to measuring the in situ chemistry of the atmosphere is a robust approach for sampling that avoids the clogging issues that befell the Pioneer Venus Large Probe, and which permits access to atmospheric gas samples from within the clouds as well as repeatedly within the lowermost scale height (i.e., from 16 km to the surface). Coupled to the requirement for in situ mass spectroscopy is the desire for direct observation of isotopes of hydrogen, oxygen, and sulfur. This is optimally accomplished by means of tunable laser spectrometer instrumentation, similar in capability to that which is part of NASA’s Mars Science Laboratory (MSL). The physical context for the required atmospheric chemistry measurements is an essential part of the scientific measurement strategy and can take advantage of the current state-of-the-art in atmospheric structure instrumentation for pressure, temperature, and accelerations. The photometry of the atmosphere beneath the cloud deck, as well as imaging of surfaces in regions not explored by the Soviet Venera landers (i.e., such as highlands) represents another opportunity for new science, and is enabled by descent imaging systems such as

31 those flown on NASA’s Phoenix Mars polar lander and on the upcoming MSL mission. These scientific measurement approaches can be combined into an optimized “descent sphere” within a probe flight system that includes an aero-entry capsule with a thermal protection system and parachutes to provide approximately an hour worth of pioneering observations about Venus to form the basis from which future missions can be designed. The VDAP architecture enables observations that were not possible during the first era of in situ Venus reconnaissance (i.e., PV, Venera), and which go beyond what orbital or flyby remote sensing can achieve. Such observations would form essential boundary conditions and constraints for models of atmospheric and climate evolution, as well as some aspects of surface-atmosphere-interior interactions. Furthermore, the suggested VDAP approach is a natural pathfinder for larger-scale landed missions or to Flagship-scale missions involving orbiters, balloons, and landed probes. Finally, direct, in situ observation of the chemistry of the atmosphere, as well as of the meter-scale morphology of the surface in rugged regions is only possible with extant technology from a deep atmosphere descent probe. Our concept for VDAP is based upon NASA Goddard Space Flight Center investments in combination with those of mission concept partners within the US. We believe the VDAP approach is the lowest risk and most cost-effective approach to resolving key scientific issues for Venus within the context of competed mission programs at NASA. Now is the time for a next-generation probe mission to Venus if future Flagship-class missions to our sister planet are to be implemented in the 2020’s by the world community.

32

VENUS PATHFINDER – A COMPACT LONG-LIVED LANDER MISSION

Ralph Lorenz

John Hopkins University Applied Physics Laboratory, [email protected]

ABSTRACT

The interior of Earth’s sister planet, and its surface meteorology, is largely unknown. To win an understanding of Venus’ seismicity and surface diurnal cycles even comparable to that we gained at Mars from Viking 35 years ago, requires a new technological capability, namely that of long-duration survival on the torrid Venus surface. Consideration of the diurnal cycle, and communications windows with the Earth, suggest a 50-day minimum requirement for surface duration, with a 200 day goal. This requires that the vehicle operate in a thermal steady state, which requires active cooling. We propose a compact equipment vault, protected by a robust dewar, with minimal heat dissipation inside. The nominal scenario is that electrical power, most of which would be devoted to cooling, would be provided by a Radioisotope Stirling Generator (we acknowledge that the present ASRG design does not tolerate either the ambient temperatures at Venus, or the likely >200g entry loads during delivery). Alternative power sources, such as batteries and fuel cells are also considered, but these fail to meet the minimum duration above. Communications would be direct-to-Earth (DTE), enabling this mission to be self-contained i.e. without an orbiter for communications support (although of course such support could substantially augment the science return). DTE capability of 350 bps would permit a total return of ~270 Mbit over 50 days. The lander design implications for a mission that includes communication through an orbiter are also discussed. The lander concept is a hybrid design including a thermally protected enclosure with exposed sensors. The internal power dissipation is limited to about 4W. Heat leaks into the thermal vault add another 20W, making the total cooling required by the protected area about 30 kW-hr. This Venus Pathfinder mission has a substantial technology development associated with it, and should be seen largely as a technology validation mission with some unique science capabilities, rather than as a Flagship-class science mission. The payload is therefore somewhat austere, and focuses on the science that is uniquely enabled by a long-lived lander. Primary instrumentation would include a descent camera (which would not be cooled,

33 and whose function would end shortly after landing), a meteorology and atmospheric optics package, and a seismometer. Possible augmentations might include a Gamma ray spectrometer and a magnetometer. Key payload issues are the deployment of the seismometer onto the ground (including decoupling it from the lander and protecting it from wind induced noise), and tolerance of the ambient conditions (temperature, pressure and composition) by the seismometer and anemometer.

34

TITAN AERIAL EXPLORER (TAE): EXPLORING TITAN BY BALLOON

Jeffery L. Hall1, Jonathan Lunine2, Christophe Sotin3, Kim Reh4, Andre Vargas5 and Patrice Couzin6

1Jet Propulsion Laboratory, [email protected] 2Dipartimento di Fisica, Università degli Studi di Roma “Tor Vergata”, [email protected] 3Jet Propulsion Laboratory, [email protected] 4Jet Propulsion Laboratory, [email protected] 5CNES

National d’Etudes Spatiales [email protected] 6Thales Alenia Space, [email protected]

ABSTRACT

Titan Aerial Explorer (TAE) is a mission concept for the exploration of Titan through use of a helium superpressure balloon. The 4.6 m diameter spherical balloon would cruise at a nominal altitude of 8 km just south of the equator and travel around the planet carried by the prevailing wind. The mission science floor is accomplished with a 3 month navigation, with a goal of complete circumnavigation that, at an estimated speed of 1 m/s, would require 6 months. The total floating mass is estimated to be 200 kg (including design margin and helium gas) of which 20 kg is science instruments carried in the gondola suspended below the balloon. The TAE mission would acquire in situ measurements of Titan’s troposphere and conduct imaging and sounding of the surface and subsurface at high resolution. The instrument suite would consist of three remote sensors—a camera (VISTA-B), near-infrared spectrometer (BSS) and radar sounder (TRS)—and three in situ experiments—an aerosol collector and analyzer (TCAA), meteorology package (ASI/MET), and a device for measuring electric and magnetic fields and conductivity (TEEP-B). In addition, tracking of the balloon’s radio signals would allow for determination of atmospheric circulation patterns at the cruising altitude. Collectively, these measurements would address the two scientific goals of the mission: (1) to explore how Titan functions as a system in the context of the complex interplay of the geology, hydrology, meteorology and aeronomy present there; and (2) to understand the nature of Titan’s organic chemistry in the atmosphere and on the surface. The linkage between the scientific goals and the measurements to be performed flows through a detailed science traceability matrix. Delivery of the balloon and gondola into the atmosphere would be via a Huygens-like entry system with a 3 m diameter aeroshell, that is itself released from a carrier spacecraft after a several year interplanetary trip. The balloon would be aerially deployed and inflated while under parachute descent. The helium inflation gas would be carried in a set of high pressure storage tanks mounted inside the aeroshell. 240 W of electrical power would be provided by 2 Advanced Stirling Radioisotope Generators (ASRG) mounted on the gondola. Waste heat from the ASRGs would be used to keep the gondola interior temperature near 20 °C. Direct-to-Earth

35 telecommunications would be provided by a 20 W X-band transmitter and a 0.75 m diameter steerable high gain antenna mounted on the gondola. It is estimated that an average of 170 Mbits of data would be transmitted to Earth during each Titan sol using ESA and/or NASA 35/34 m ground antennas. The balloon would be fabricated from a polyester film and fabric laminate. A vent valve and a few kilograms of ballast would be carried to enable a limited number of altitude excursions during the mission. Otherwise, the superpressure design will result in constant altitude flight with very small deviations of tens of meters from the nominal 8 km float altitude.

36

AN ADVANCED DESIGN FOR A TITAN BALLOON

Julian Nott1, Don Cameron2, Don Day3, Greg Mungas4

Nott Technology LLC1, President Cameron Balloons Ltd2, President Day Weather Inc.3, CTO Firestar Engineering LLC4

ABSTRACT

Ever since Cassini arrived at Saturn and the Huygens Probe descended onto Titan, Titan has emerged as an ever more interesting place with corresponding continuously growing interest in a follow-on mission. A balloon appears to be an ideal vehicle to explore Titan. It would carry cameras and instruments like a Mars Rover. But while the two existing Mars Rovers have, combined, traveled less than thirty miles in six years, a balloon could cover thousands or tens of thousands of miles. In addition is has emerged that Titan has weather and other conditions that are dramatically better for balloon flight than provided by the Earth's weather and conditions. Balloons emerge as very attractive for Titan in-situ exploration. Conditions are of course still partly uncertain, see below. It is very beneficial for a Titan balloon to be able to change altitude at will. It is obviously a major advantage for science observations if the balloon can view wide panoramas at altitude and descend to take close-up pictures and to lower instruments to touch both solid and liquid surfaces directly. Being able to change altitude also means the balloon can use light winds to travel slowly at low altitude for observations or climb into stronger upper winds to travel long distances. And of great importance it allows for substantial steering. The extent to which contemporary terrestrial balloons are steered very effectively, simply by changing altitude is not fully appreciated outside the field. It will be impossible to know exactly what Titan conditions will be offer until the balloon actually arrives. So the more flexibility the balloon can have the better, perhaps to fly above certain weather or fly below icing conditions or avoid bad weather altogether by steering. Balloons have long been proposed for Titan, but serious interest in hot air balloons began followed the seminal 2005 paper [Jones, Fairbrother et al] which suggested that a Titan hot air balloon could be heated by the surplus heat from the radioisotope thermoelectric generators used to power all craft at the outer planets where sunlight is too weak for solar cells to be effective. But since then there has been only limited change in the basic concepts for such a balloon. This paper describes in detail a system that hopefully substantially improves over the 2005 proposal.

37 The paper will include:

• A detailed description of a highly insulated balloon envelope which used multiple fabric layers giving sufficient insulation that it can be heated by the surplus heat from the newly developed Advanced Stirling Radioisotope Generator. This has several major advantages. It allows for a lighter balloon system and requires only one eighth [depending on the design] of the amount of radioisotope material, a very scarce and expensive resource. In addition a smaller highly insulated balloon has a higher buoyancy per unit volume. This of itself gives greater resistance to gusting. In the design described the full hydrostatic pressure at the top of the balloon is carried to the mouth following the classic Cameron "Coke Can" design. This concept has been tested in innumerable balloons flown over three decades. This give pressure at the mouth and this too is very valuable to resist any atmospheric gusting. As yet another feature to resist gusting, the balloon will incorporate a "Base Parachute", a fabric check valve mimicking the extremely reliable crown parachute in universal use in hot air balloons for several decades. Finally a smaller, hotter balloon has a smaller displacement and correspondingly lower inertia, yet another beneficial quality when encountering a gust or other unexpected weather. Also this smaller inertia and cross section area mean that it can more easily be moved sideways if it is fitted with propellers as has sometimes been proposed. While such a balloon might be thought to be complex, it is no more so than such balloons as the two piloted balloons which successfully flew around the world and trivial in complexity compared to a space craft. Moreover this kind of design can be quickly and very inexpensively prototyped.

As well as all these advantages, substantially smaller balloons might allow a mission with a smaller rocket or allow a balloon as a "Hitchhiker Payload" on a large mission or allow two or three balloons to be sent on a mission where one was originally planned.

• A detailed thermal analysis of the balloon envelope design, based on the extensive physical and theoretical thermal modeling already completed [Colonius, Nott, et. al. 2009].

• Insights into Titan weather extrapolated from terrestrial experience. Currently the best assumption is that Titan is rather earthlike. So the practical experience gained by forecasters specializing in balloons from the more than four million piloted flights made over the last five decades by terrestrial hot air balloons is invaluable to draw on.

• A description of an emergency heating system. As mentioned there will be uncertainties about Titan conditions even after a balloon is flying there. The balloon described is very much better to survive unexpected weather than other concepts. But the paper will also include a detailed description of an emergency heat source using hydrazine or a non-toxic, low temperature tolerant, very high energy density, NOFBX monopropellant (in flight experiment development for launch to Space Station in 2012 and in prototype ascent engine development and

38 test for the Mars Sample Return Mars Ascent Vehicle) to allow the balloon even better ability to survive encounters with turbulence, downdrafts and other unexpected conditions. Despite almost 230 years practical experience, terrestrial balloon operators still encounter weather which has never been experienced previously. Assuming, as is currently anticipated, weather like everything else on Titan is very earthlike, unexpected events will be encountered. With or without the features to add robustness to the balloon described above, an emergency heat source is seen by some terrestrial operators as improving reliability perhaps by an order of magnitude or more, although this is not fully quantifiable.

• A detailed description of a method of air-launch inflation, meaning that the balloon fills, heats and flies away while falling through the atmosphere. This extrapolates from 50 years experience of contemporary hot air balloon operations including thousands of hot air balloons that have been successfully air-launched.

In all it is hoped to present a design with substantial benefits over previous Titan hot air balloon proposals, over any proposal where the balloon flies at a fixed altitude and any design such as an AM, mixed gas and hot air balloon, where the balloon uses a lifting gas which will inevitably suffer lifting gas loss over time.

39

MISSION CONCEPT FOR ENTRY PROBES TO THE FOUR OUTER PLANETS BASED ON E-SAIL PROPULSION

Jean-Pierre Lebreton1, Pekka Janhunen2, Sini Merikallio2, Petri Toivanen2

ESA/ESTEC, Solar System Missions Division, Noordwijk, The Netherlands.1

[email protected], Finnish Meteorological Institute, FMI, Helsinki, Finland2, [email protected], [email protected], [email protected]

ABSTRACT

The Electric Solar Wind Sail (E-sail) is a new propulsion method that uses long, thin, positively charged tethers to convert solar wind momentum flux into thrust. The E-sail concept was invented in 2006 (http://www.electric-sailing.fi/) and its development is partly funded by the European Unionʼs Seventh Framework Programme for Research and Technological Development, EU FP7. According to current estimates, the E-sail can be 2-3 orders of magnitude more efficient than traditional propulsion methods (chemical rockets and ion engines) in terms of produced lifetime-integrated impulse per propulsion system mass. In an E-Sail, the “screen” that reflects the solar wind protons is made by a network of the electrostatic sheaths forming around each of the highly positively charged tethers. Despite the fact that the solar wind dynamic pressure is smaller than the radiation pressure of solar photons, the E-Sail can be more efficient than the photonic Solar Sail as the electrostatic screen that reflects the solar wind protons can be orders of magnitude larger -when using long, highly-charged tethers-, than that of a solar sail. Although the solar wind flux and the solar radiation flux both vary with the squared distance to the sun, the thrust produced by an E-Sail is inversely proportional to the distance from the sun (F α 1/r) as the sheath size around each tether increases when the solar wind density decreases, while the thrust produced by a photonic solar sail is directly proportional to the solar radiation flux (F α 1/r2). This makes the E-Sail a propellantless method very attractive for outer solar system missions. The science case for entry probes in the four outer planets has been made by several authors (e.g. Owen T. C., Atmospheric Probes: Needs and Prospects, in International Workshop on Planetary Probes, ESA SP-544, 2004; Atreya et al., Multiprobe exploration of the giant planets- shallow probes, Proceedings, International Planetary Probe Workshop, IPPW-3, ESA SP-WPP263, 2006). In this paper, we describe the concept of a multi-probe mission to each of the four outer planets that is based on a common concept of a carrier-entry probe composite propelled by an E-Sail to each destination for a direct entry into the atmosphere of the planets. The E-sail technology would allow significantly reduced travel times and reduced launch costs compared to traditional propulsion techniques. The concept of a standard 1-N E-Sail has been recently

40 studied in detail (Janhunen et al., Electric solar wind sail: Towards test missions (Invited article), Rev. Sci. Instrum., 81, 111301, 2010, doi:10.1063/1.3514548). It requires hundred 20 km long tethers charged to several 10ʼs of kV. It would allow to propel a 500 kg spacecraft (carrier-probe composite, but excluding the E-Sail propulsion stage) to Jupiter in a mere 1 year, to Saturn in 1.7 year, to Uranus in 3 years and to Neptune in 4.6 years. The four probes could either be launched independently by a small launcher or together by more powerful launcher on a trajectory that would place them in the solar wind. The constraints of a planetary launch window would not apply, thus providing increased launch flexibility compared to a classical planetary mission launch window. The arrival velocity of the probes would be relatively large, but it would not significantly affect the entry speed as this key parameter would essentially be governed by acceleration due to planet gravity (the same would not be true at Titan). This makes the E-sail concept a very appealing propellant-less method to conduct a multi-probe mission to the four outer planets at an affordable cost, especially if a similar entry probe design would be used for all four planets. The scientific return that would be allowed by identical probes to each of the four outer planets would need to be evaluated carefully to confirm the attractiveness of the proposed approach. Alternatively, for a higher cost, each probe could be tailored for optimizing the science return at each of the four planets.

Session 3 - Science from Probes and Penetrators

42

NEW TOOLS AND METHODS TO FULLY CHARACTERIZE THE ATMOSPHERIC ENVIRONMENT

FOR A MARTIAN EDL APPLICATION TO THE 2016 EXOMARS DESCENT MODULE

F. Forget1, A. Spiga1, L. Montabone1, E. Millour1, A. Colaitis1, V. Bourrier1 F. Gonzalez-Galindo2 S. R. Lewis3, S. Portigliotti4.

Laboratoire de Météorologie Dyna- mique, IPSL, Paris, France ([email protected])

1, Instituto de Astrofísica de Andalucía, Granada, Spain2, Department of Physics and Astronomy, The Open University, Milton Keynes, UK3, Thales Alenis Space – Italia,

Tonio, Italy4

ABSTRACT

Because of the presence of a relatively thin and highly variable atmosphere, the Entry Descent and Landing of Martian probes is a difficult task which requires the best possible characterization of the Martian atmosphere, from the upper atmosphere to the boundary layer. Ideally, one want to predict as accurately as possible density, temperature, pressure and winds (including their variability and perturbations) as well as possible updraft and downdraft in the boundary layer (for the parachute phase) and the aerosols mixing ratio (in particular for the heatshield erosion), etc. We will review the various tools that are now available to address these questions (with a focus on the tools and data that have been made available in the last couple of years). This includes:

• Spacecraft observations, which have been the reference source of information to prepare an IDL. A climatology of the Martian atmosphere has been collected since 1999 (beginning of mapping mission of Mars Global Surveyor) by various instruments. In particular, data from the Mars Climate Sounder on MRO which monitor the atmosphere from the surface to above 70 km are now available. Altogether, we have now detailed climatologies on the Martian weather for 7 years, which allows us for the first time to derive reliable statistics on the year-to-year variability. At some seasons, it seems that the Martian atmosphere is very repeatable from year to year. This can give a lot of confidence in the prediction.

• Global Climate Model (GCM) and derived tools. GCMS have been extensively used to provide reliable climatologies of the Martian climate. They are constantly improving and are now able to predict the Martian weather anywhere and at any season with a striking accuracy. GCM simulations can be used directly or through tools that exploit the GCM outputs to provide engineering tools like Mars Gram or the Mars Climate database suitable to combine outputs from the GCM with variability models suitable for Monte-Carlo EDL simulations. The MCD is for instance designed to simulate a variety of possible entry profile for various dust loading and meteorological conditions. It also include a tool designed to predict surface pressure with the high- est possible accuracy by combining 1) reference pressure measure of Viking lander 1 site at a giv-

43 en seasonal date, 2) large scale spatial variation due to meteorology (including thermal tides at a given local time) from the GCM and 3) small scale variations due to topography, using 1/32 degree MOLA data (~2km horizontal resolution). Overall this tool is thought to predict pressure with an error of less than a couple of percent.

• Data assimilation are obtained by optimally combining observations (obtained at various locations and time) with the a-priori knowledge from a GCM. State of the art techniques that are used on the Earth to construct reference climatologies (“re-analysis”) are now available on Mars from various group using the MGS TES and MRO MCS data.

• Meso-scale models with a resolution of a few kilometers are also necessary to complement the GCMS, in particular to predict the local winds resulting from the topography below 10 km and the landing conditions.

• Large Eddy Simulation models (LES) are new kind of tools with a resolution of a few tens of- meters which are able to simulate the convective boundary layer environment (during daytime) at a landing site, and in particular the strong convective updraft and downdraft which may be dangerous for a probe under a parachute.

Figure: an example of “Large Eddy simulations” of the convective boundary layer used to model the parachute descent of a probe. In our presentation, we will review all the tools available, and illustrate the kind of results that can be obtained with the case of the ESA Exomars Descent module schedule to land on Mars in October 2016.

44

Entry Trajectory Reconstruction Using Phoenix Radio Link

Ö. Karatekin1 and S. W. Asmar2

Royal Observatory of Belgium1Jet Propulsion Laboratory, California Institute of

Technology2

ABSTRACT

The Phoenix Mars Lander entered the Martian atmosphere on May 25, 2008. All ensuing communications during Entry Descent and Landing (EDL) path were via an UHF uplink to a Mars orbiting spacecraft. The Odyssey orbiter relayed the Phoenix data to the Deep Space Network station (DSN) at Goldstone. In addition, objective of this activity was to monitor the state of the lander during critical stages of the EDL. The data can now be explored for utility to reconstruct the entry trajectory provided that the received UHF signal is not too noisy. The recorded signal profile from Phoenix EDL is processed to quantify the accuracy of the reconstructed trajectory and the atmospheric profiles (density, pressure, and temperature) determined along this trajectory.

45

AIRBORNE OBSERVATION OF THE HAYABUSA SAMPLE RETURN CAPSULE RE-ENTRY

Jay H. Grinstead1, Peter M. Jenniskens2, Alan M. Cassell3, Jim Albers2, Michael Winter4

NASA Ames Research Center1, The SETI Institute2, ERC Incorporated3, University

Affiliated Research Center/University of California4

ABSTRACT

The Japan Aerospace Exploration Agency (JAXA) recently completed their Hayabusa asteroid exploration mission. Launched in 2003, Hayabusa made contact with, and retrieved a sample from, the near-Earth asteroid Itokawa in 2005. The sample return capsule (SRC) re-entered over the Woomera Test Range (WTR) in southern Australia on June 13, 2010, at approximately 11:21 pm local time (09:51 UTC). The SRC re-entry velocity was 12.2 km/s, making it the second-fastest Earth return velocity behind NASA1s Stardust sample return capsule re-entry in 2006. From a space technology development perspective, Hayabusa’s re-entry functioned as a rare flight experiment of an entry vehicle and its thermal protection system. In collaboration with the SETI Institute, NASA deployed its DC-8 airborne laboratory and a team of international researchers to Australia to observe the re-entry of the SRC. The use of an airborne platform enables observation above most clouds and weather and greatly diminishes atmospheric absorption of the optical signals. The DC-8’s flight path was engineered and flown to provide a view of the spacecraft that bracketed the heat pulse to the capsule. A suite of imaging instruments on board the DC-8 successfully recorded the luminous portion of the re-entry event. For approximately 70 seconds, the spectroscopic and radiometric instruments acquired images and spectra of the capsule, its wake, and destructive re-entry of the spacecraft bus. Figure 1 shows a perspective view of the WTR, the SRC re-entry trajectory, and the flight path of the DC-8. The SRC was jettisoned from the spacecraft bus approximately 3 hours prior to entry interface. Due to thruster failures on the spacecraft, it could not be diverted from the entry path and followed the trajectory of the SRC, where it burned up in the atmosphere between approximately 100 and 50 km altitude. Fortuitously, the separation distance between the spacecraft and SRC was sufficient to clearly resolve the SRC from the debris field of the burning spacecraft. Figure 2 shows a frame from a high-definition television camera on board the aircraft and denotes the locations of the SRC and spacecraft bus debris. Most instruments had the capability to spectrally resolve the emission of the SRC and spacecraft debris fragments. The spectral range covered by the instruments spanned from the near ultraviolet (approximately 300 nm) to the short wave infrared (approximately 1700 nm). The instruments were calibrated before and after the observation flight. Reference standard irradiance source lamps were used for calibration to absolute spectral radiance. Atomic line source lamps were used for wavelength calibration. Atmospheric

46 absorption will be corrected for using extinction calculations based on an atmosphere model and range-to-target distances. Figure 3 shows a preliminary spectrum recorded simultaneously by four separate instrument platforms; signal diminution due to atmospheric absorption in the infrared by H2O and O2 has not been corrected. Ground- based observation teams from the US, Australia, and Japan also recorded the re-entry. The ground and airborne observation data have been used to reconstruct the as-flown trajectory of the SRC. The Hayabusa observation campaign’s objectives and methods were similar to that of the Stardust re-entry observation. However, unique technical and programmatic challenges were encountered arising from coordination and cooperation with JAXA and the Australian authorities. A brief summary of the Hayabusa mission, the airborne observation campaign, data, and analysis will be presented.

Figure 1. Perspective view of the Woomera Test Range in South Australia showing the re-entry trajectory of the Hayabusa SRC and the flight path of the DC-8 observation aircraft. Peak heating, predicted to occur at 58 km altitude, is noted.

47

Figure 2. Single-frame image from a high-definition television camera aboard the DC-8 observation aircraft. The Hayabusa SRC is well separated from the burning debris of the spacecraft bus.

Figure 3. Composite spectrum of the SRC emission at one point in time as seen with four different instruments. Atomic and molecular emission features in the shock layer are noted. Absorption due to atmospheric O2 and H2O has not been corrected for. 2 Apparent flux (W/m/nm)

48

RADIATION MODELING FOR THE REENTRY OF THE HAYABUSA SAMPLE RETURN CAPSULE

Michael W. Winter1, Ryan D. McDaniel2, Yih-Kanq Chen2, Yen Liu2, David Saunders3

University Affiliated Research Center UARC, UC Santa Cruz NASA Ames Research

Center, [email protected], NASA Ames Research Center2, ERC, Incorporated, NASA Ames Research Center3

ABSTRACT

On June 13, 2010 the Japanese Hayabusa capsule performed its reentry into the Earth’s atmosphere over Australia after a seven year journey to the asteroid Itokawa. The reentry was studied by numerous imaging and spectroscopic instruments onboard NASA's DC-8 Airborne Laboratory and from three sites on the ground, in order to measure surface and plasma radiation generated by the Hayabusa Sample Return Capsule (SRC). Before flight, computations of the flow field around the forebody were performed using the in- house code DPLR [1, 2] assuming an 11-species (N2, O2, NO, NO+, N2+, O2+, N, O, N+, O+, and e–) air in thermochemical nonequilibrium at peak heating. The results were used as input for the material response code FIAT [3] to calculate surface temperatures of the heat shield. Finally, the thermal radiation of the glowing heat shield was computed based on these temperatures and propagated to the predicted observation position taking into account the influence of the observation angle and of atmospheric extinction yielding estimates of thermal radiation to be measured by the observing instruments during reentry. These estimates were used to provide calibration sources of appropriate brightness. Post flight, the flow solutions were recomputed to include the whole flow field around the capsule at 11 points along the reentry trajectory using updated trajectory information. Again, material response was taken into account to obtain most reliable surface temperature information. These data will be used to compute thermal radiation of the glowing heat shield and plasma radiation by the shock/post-shock layer system to support analysis of the experimental observation data. For this purpose, lines of sight data are being extracted from the flow field volume grids and plasma radiation will be computed using NEQAIR [4] which is a line-by-line spectroscopic code with one-dimensional transport of radiation intensity. The procedures being used were already successfully applied to the analysis of the observation of the Stardust reentry [5].

49 Details of the numerical procedures and the calibration approach will be provided in the full-length paper.

Acknowledgments The observation campaign was funded and managed by the Orion Thermal Protection System Advanced Development Project and the NASA Engineering and Safety Center. The present work was supported by NASA Contract NAS2-03/44 to UARC, UC Santa Cruz, and by NASA ContractNNA10DE12C to ERC Incorporated. The authors would like to thank Dr. George Raiche (Chief, Thermophysics Facilities Branch, NASA ARC) and Dr. Aga Goodsell (Chief, Reacting Flow Environments Branch, NASA ARC) for support of modelling and simulation aspects of the present work. Furthermore, the authors wish to acknowledge the support of Jay Grinstead, NASA Ames, Peter Jenniskens, SETI Institute, Alan Cassell, ERC, and Jim Albers, for mission planning and for providing trajectory information, and Nicholas Clinton and Jeffrey Myers, UARC, for compiling Modtran computations for atmospheric extinction.

References [1] Wright, M. J., Candler, G. V., and Bose, D., Data-Parallel Line Relaxation Method of the Navier-Stokes Equations, AIAA Journal, Vol. 36, No. 9, 1998, pp. 1603–1609. [2] Wright, M.W., White, T., and Mangini, N., Data Parallel Line Relaxation (DPLR) Code User Manual Acadia – Version 4.01.1, NASA/TM-2009-215388, October 2009. [3] Chen, Y.-K., and Milos, F.S., Ablation and Thermal Analysis Program for Spacecraft Heatshield Analysis, Journal of Spacecraft and Rockets, Vol. 36, No. 3, 1999, pp. 475-483. [4] Whiting, E. E., Park, C., Liu, Y., Arnold, J. O., and Paterson, J. A., NEQAIR96, Nonequilibrium and Equilibrium Radiative Transport and Spectra Program: User’s Manual, NASA RP-1389, NASA, December 1996. [5] Yen Liu, Dinesh Prabhu, Kerry A. Trumble, David Saunders, and Peter Jenniskens, Radiation Modeling for the Reentry of the Stardust Sample Return Capsule, Journal of Spacecrafts and Rockets, Vol. 47, No. 5, September– October 2010.

50

GIANT PLANET FORMATION, SATURN AND URANUS ENTRY PROBES, AND THE DECADAL

Sushil Atreya

University of Michigan, www.umich.edu/~atreya

ABSTRACT This talk will focus on the past, present and the future of the origin and evolution of the giant planets and their atmospheres problem. The core accretion model has been the conventional model of the formation of the giant planets for four decades [1]. According to this model, a core formed first from grains of ice, rock, metals and refractory material of the protoplanetary nebula. Gases were trapped in these solids. Upon reaching a critical mass of 10-15 Earth Mass, the core gravitationally captured the most volatile of the gases, neon, hydrogen and helium from the surrounding nebula. This led to the gravitational collapse of the protoplanetary nebula. These last volatiles and the gases released from the core during accretional heating were the origin of the atmosphere. Thus the elemental abundance in the atmosphere would reflect that in the protoplanetary nebula, i.e. solar composition with the same abundance ratio to hydrogen as in the sun. Surprisingly, the Galileo probe found the abundance of heavy elements (relative to H) in Jupiter’s atmosphere enriched compared to the sun [2-4]. Moreover, the enrichment factor is uneven, varying from 2 to 6, i.e. the inter-elemental abundances are non-solar [5,6]! One missing piece of Jupiter’s formation puzzle is the oxygen elemental abundance (O/H), however. Oxygen is sequestered in water in Jupiter, and the Galileo probe entered a 5-micron hotspot, the “Sahara Desert” of Jupiter, which was dry. The determination of water is critical to the models of the origin and evolution of Jupiter as water was presumably the original carrier of the heavy elements that formed the core. If enriched by a similar factor as the other heavy elements, water could comprise one-half of the mass of Jupiter’s primordial core, or greater. Juno will measure and map water in Jupiter’s troposphere by passive microwave remote sensing in 2016. A comparison of the elemental abundances in Saturn with those in Jupiter is essential for constraining the formation models of the gas giant planets. However, remote sensing observations of Saturn from the Cassini orbiter have determined only one element, carbon, since remote sensing is not suited to measure the other heavy elements. A probe is required [6-8]. Finally, the models of the formation of the giant planets would be incomplete without similar heavy element data of the icy giant planets, Uranus and Neptune [9]. Only carbon is constrained in these planets, and the data have high uncertainty. The ice/gas ratio in

51 these planets is 90-95% compared to 3-10% for the gas giants. Whether or not the icy giant planets followed a similar path of accretion as the gas giant planets can be understood only after the determination of a full suite of their elemental composition [5,9]. The NRC Planetary Decadal Survey (2013-2023) opens a path forward for entry probes into Saturn and Uranus. I will discuss also how within the available resources, both missions will be able to determine the elemental composition that is key to the models of formation of the giant planets in particular, the solar system in general, and, by implication, the extrasolar planetary systems.

Bibliography

Author’s own publications can be downloaded from www.umich.edu/~atreya. 1. Formation of the Giant Planets, H. Mizuno, Prog. Theor. Phys. 64, 544, 1980. 2. The Composition of the Jovian Atmosphere as Determined by the Galileo Probe Mass Spectrometer, H.B. Niemann, S.K. Atreya, et al., J. Geophys. Res. 103, 22831, 1998. 3. Comparison of the Atmospheres of Jupiter and Saturn: Deep Atmospheric Composition, Cloud Structure, Vertical Mixing, and Origin, S.K. Atreya, M.H. Wong, T.C. Owen, P.R. Mahaffy, H.B. Niemann, I. de Pater, Th. Encrenaz, and P. Drossart, Planet Space Sci. 47, 1243, 1999. 4. Composition of the Atmosphere of Jupiter - An Update, and Implications for the Extrasolar Giant Planets, S.K. Atreya, P.R. Mahaffy, H.B. Niemann, M.H. Wong, T.C. Owen, Planet Space Sci., 51(2), 105, 2003. 5. Coupled Chemistry and Clouds of the Giant Planets – A Case for Multiprobes, S.K. Atreya, A.S. Wong, Space Sci. Rev. 116, Nos. 1-2, pp 121-136, 2005. 6. Saturn Probes: Why, Where, How? S. K. Atreya, Proceedings of the International Planetary Probe Workshop IPPW-4, 2007, http://www.mrc.uidaho.edu/~atkinson/IPPW- 4/Session_4/Papers/4_6ATREYA.pdf. 7. Multiprobe Exploration of the Giant Planets – Shallow Probes, S. K. Atreya, S. Bolton, T. Guillot, T. C. Owen, International Planetary Probe Workshop IPPW-3 Proceedings, ESA Special Publication WPP263, 2006. 8. Saturn Exploration Beyond Cassini-Huygens, T. Guillot, S.K. Atreya, S. Charnoz, M. Dougherty, P. Read, in Saturn From Cassini-Huygens (M. K. Dougherty et al., eds.), Chapter 23, pp 745-761, 2009, DOI 10.1007/978-1-4020-9217-6_23, Springer Dordrecht, New York. 9. Clouds of Neptune and Uranus, S. K. Atreya and A. S. Wong, NASA Proceedings of Planetary Probes Workshop NASA/CP-2004-213456 (E. Venkatapathy, et al., eds.), pp 107-110, 2004.

52

2012 DECADAL SURVEY GIANT PLANET ENTRY PROBE SCIENCE

Thomas R. Spilker(1), David H. Atkinson(2)

(1)Jet Propulsion Laboratory, California Inst. of Tech., MS 301-170S, 4800 Oak Grove Drive, Pasadena, CA 91109, USA, [email protected] (2)Univ. of Idaho, Dept. of Electrical & Computer Eng., Moscow, ID 83844, USA, [email protected]

ABSTRACT

On March 7, 2011 the US National Research Council released the draft report of the results of its 2012 Planetary Science Decadal Survey (PSDS). Atmospheric entry probes to the giant planets were well-represented in the PSDS, with addition of a Saturn Probe mission to the list of recommended NASA New Frontiers Program missions, inclusion of an entry probe along with an orbiter in a new Uranus mission concept, and recommendations for technology development leading to a mission with a Neptune orbiter and entry probe for the following decade (2023-2032). This follows closely the recommendations of the 55-author PSDS white paper, “Entry Probe Mission to the Giant Planets.” Studies conducted for the PSDS suggest that under the usual New Frontiers Program approach for budget reserves, it might be possible to add Tier 2 science investigations and instrumentation to a Saturn entry probe, enhancing the science return beyond the Tier 1 objectives. The Uranus mission would most likely be a small flagship-class mission. Studies indicate that for the most probable entry geometries for a Uranus probe mission conducted in the PSDS time frame, an entry system designed for a Saturn probe mission could also be used for a Uranus probe, without excessive margin in the design. Significant overlap of the Tier 1 science objectives at Saturn and Uranus also provides an opportunity for use of common instruments. A descent module (including instrumentation) for Uranus would be well suited for a Neptune mission, though the Neptune system presents some unique issues for the entry system. A prograde entry at Neptune would have an atmosphere-relative entry speed significantly slower than that of a Uranus probe, so the Uranus entry system might be overdesigned for Neptune, while a retrograde entry would have an atmosphere-relative entry speed significantly faster than that of a prograde Saturn probe, requiring a higher-performance entry system. As discussed in the giant planet entry probe white paper, this series of entry probe missions would complete the initial in situ exploration of all four of our solar system’s giant planets, thereby allowing comparison of gas giants (Jupiter and Saturn) to ice giants (Uranus and Neptune) and comparisons within those classifications These comparisons are expected to yield significant progress in understanding formation processes and time scales for the giant planets, and for the solar system as a whole. Such understanding would help to understand the formation of other planetary systems, and the tremendous

53 variation in the architectures of exoplanetary systems now being discovered. This presentation will summarize the science described in the PSDS for these future mission concepts, and some of the instrumentation options for implementing them. It will also discuss the programmatic environment for each mission concept, describing decisions and events that might lead to their implementation as flight projects.

54

OUTER PLANET DOPPLER WIND MEASUREMENTS

D.H. Atkinson(1), S.W. Asmar(2), T.R. Spilker(2)

(1)University of Idaho ([email protected]) (2)Jet Propulsion Laboratory, California Institute of Technology

ABSTRACT

The atmospheres of the giant planets represent time capsules dating to the epoch of solar system formation. For a complete and self consistent understanding of the giant planets, an integrated knowledge of the structure of the atmosphere is needed, including composition, clouds, energy structure, and dynamics. In particular, atmospheric dynamics - winds, waves, convection, and turbulence - are responsible for horizontal and vertical mixing of atmospheric constituents including noble gases and volatiles and their respective isotopes, and upwelling of disequilibrium species providing diagnostics of deep atmosphere compositions and chemistries. Winds and waves are essential to understanding the meteorology including the structure, location, and life cycle of clouds, and momentum transfer and overall energy structure of the atmosphere. The altitude profile of the winds places valuable constraints on the location of solar energy deposition, which affects cloud structure and the static stability of the atmosphere, and can provide an indication of the relative importance to the atmospheric energy structure of solar energy relative to internal energy sources. Some measurements of the composition, cloud structure, and dynamics of the upper atmosphere can be obtained from remote sensing. However, to measure beneath the clouds requires in situ sampling from an atmospheric entry probe, from which the dynamics of the atmosphere can be inferred by utilizing Doppler techniques to track the probe motions throughout descent. Although Doppler wind methodologies depend strongly on the target – whether a large, rapidly rotating giant planet or a smaller, more slowly rotating terrestrial planet (including Titan), the overarching principles are the same in either case. Accurate reconstructions of the probe entry and descent profile, including location, altitude, and descent speed, and the assumption of predominantly zonal (east-west) winds are used to extract the relatively small signature of probe motions resulting from atmospheric dynamics, reflected as Doppler residuals in the probe radio link frequency profile. From the residuals, the vertical profile of zonal winds is retrieved utilizing an iterative inversion algorithm that accounts for the integrated effect of the winds on the probe descent longitude. Analysis of the probe radio link frequency residuals can also provide evidence of atmospheric waves and turbulence, as well as probe microdynamics including spin and pendulum motion. Doppler wind measurements require ultrastable oscillators (USO) in both the probe transmitter and the receiver. Two USO types have Doppler wind flight heritage – crystal oscillators flown on the Galileo probe mission to Jupiter, and atomic USOs using

55 rubidium gas cells on the Huygens (Titan) probe. Key USO characteristics include the warm-up time and power profile, the time profile of stability during warm-up, long-term stability, and phase noise. Other requirements for Doppler wind measurements on mission design are relatively minor. This paper will provide an overview of Doppler wind methodologies used on the two outer solar system probe missions to date, and will present a preliminary discussion of considerations and requirements for future giant planet Doppler wind measurements.

56

TITAN AERIAL EXPLORER

Jonathan I Lunine1, Christophe Sotin2

University of Rome1, Jet Propulsion Laboratory, California Institute of Technology2

ABSTRACT

We propose the Titan Aerial Explorer (TAE) mission that would deploy and operate a superpressure helium balloon in the lower atmosphere of Saturn’s giant moon, Titan. The Cassini- Huygens mission revealed Titan to be a body with an active hydrological cycle involving methane, ethane and a variety of other organic molecules. It found methane oozing from the surface at the Huygens landing site, and liquid organics residing in vast near-polar deposits whose extent rivals or exceeds the great lakes and seas on Earth. Cassini observed methane clouds forming as convective storms in the summertime south, as ghostly echoes superimposed on methane seas sheathed in late winter darkness, and as unexpectedly vast outbursts in the mid- latitudes as the Sun crossed the equator of Titan at equinox. The geologic history of the surface remains a mystery after six years of Cassini data and will continue to be a mystery through the end of the Cassini mission. The variety of surface features and atmospheric phenomena seen only at moderate and low resolution by the orbiter tease us, because we know from nature of the one site visited in situ by the Huygens probe that hidden among the dunes and channels, the mountains and lake shores, is a complex history of climate change and chemical evolution tied to methane and its prodigious variety of organic products. We seek to understand this history by deploying at Titan the one type of vehicle that combines the mobility and coverage of the orbiter with the capability for high resolution and in situ observations demonstrated by the Huygens lander, and does so in an aerodynamically stable and low-risk fashion—an aerostat (balloon plus gondola). TAE would utilize a helium-filled super-pressure (or “pressurized”) balloon, rather than a hot air (montgolfière) design, as in many previous studies. The great advantage of the pressurized balloon is the maturity of its inflation and deployment scheme. Its disadvantage is relative sensitivity to the presence of small holes that can reduce dramatically the mission lifetime. The threshold science mission is achieved after a 3-month long navigation halfway around Titan, while the goal is a complete circumnavigation (6 months). TAE science is organized around two themes, which emphasize the special nature of Titan and at the same time its important connections to studies of other planets and the Earth. These are (1) The presence of an atmosphere and liquid volatile “hydrologic” cycle, which implies climate evolution through time and (2) organic chemistry, which is pervasive through its atmosphere, surface, and probably interior. Therefore the first science goal is to explore how Titan functions as a system in the context of the complex

57 interplay of the geology, hydrology, meteorology, and aeronomy present there. Goal 2 is to understand the nature of Titan’s organic chemistry in the atmosphere and on its surface. These in turn lead to a set of primary science objectives for a balloon-borne system, which can then, through the science investigations that devolve from them, be addressed through a set of measurement objectives.

SESSION 4 - EDL Technology Development

59

Going Beyond Rigid Aeroshells:

Enabling Venus In-Situ Science Missions with Deployables

Ethiraj Venkatapathy1, Todd White2, Gary Allen3, and Dinesh Prabhu4

NASA Ames Research Center, Moffett Field, CA USA

ABSTRACT

Missions to Venus and Jupiter are amongst the most challenging of all in situ science missions employing entry probes because of the severity of environments encountered. Although NASA’s decadal survey for Planetary Science, is eagerly awaited by the planetary science and EDL (Entry, Descent and Landing) communities, it is yet to be released as of this abstract submission and Venus could be a high priority destination. If in situ science missions to Venus are recommended in the decadal survey, then it imperative to understand the challenges and limitations of the “conventional” aeroshell architectures and offer solutions for the near and longer term exploration of Venus. Past missions (both US and USSR) to Venus have relied on traditional rigid aeroshell architectures. The entry conditions are ~5 kW/cm2 peak heat-flux, peak pressures of 5 to 10 atmospheres, and deceleration loads of 250 to 450 g’s necessitating the use of high-density and high-performance carbon-phenolic (CP) ablative thermal protection system (TPS) for the aeroshell. However, the capability for manufacturing base for heritage carbon-phenolic has significantly eroded over the last decade, and four white papers advocating revival of this manufacturing capability were submitted to the Planetary Science Decadal Survey. In the short term, reviving or developing a process for the manufacture of alternate (i.e., non-heritage) carbon-phenolic is essential because the availability of this material enables science missions to the Outer Planets, high-speed sample return missions to earth, as well as missions to Venus. However, in the long term, maintaining a material manufacturing line when NASA is effectively the only consumer of such material is probably not very cost effective. Furthermore, availability of proven high-density ablators still does not address the high decelerations at extreme entry conditions. Testing capabilities, especially arc jet capabilities, to qualify and flight certify materials is yet another challenge. Recent works by Venkatapathy, Laub, Hartman, Arnold, Wright, and Allen [1,2], outline the technical approach needed for development, testing, and qualification of ablating TPS materials, especially carbon-phenolic (including alternates to heritage CP). In addition to the high heating environment that dictates the need for very high

60 density (1.4 g/cc) ablating TPS material, the deceleration loads encountered in ballistic entries into Venusian atmosphere are so high (250-450 g’s) that it is necessary to design not only the TPS and the underlying structure, also the entire entry system including the instruments to be robust enough to withstand the entry g-load. Testing and qualifying the probe and instruments at such deceleration loads is a considerable challenge. Furthermore, the delicate instruments used for scientific measurements may require extensive efforts to make them robust and making the mission expensive. As a result, missions that need to fit in accost class such as New Frontier and Discovery may forego doing the science. Although there is the possibility of a potential compromise between robustness and performance of the instruments, such a compromise will always invariably be at the cost of additional mass to the overall system, and perhaps a significant reduction in the value of the mission. As a case in point, Venus in situ science missions proposed in the past, such as balloon and lander missions, have been forced to severely reduce the duration of the missions because the high deceleration loads were beyond those that could be withstood by advanced (but delicate) radioisotope power systems (such as ASRG) or the RTG powered Sterling cycle refrigeration system [3,4], The recent reorganization at NASA and the creation of the Office of Chief Technologist (OCT) at NASA has opened up avenues to “think outside the box” and develop new technologies to meet the challenges of entry, descent and landing in any planetary atmosphere. The proposed paper will showcase results of recent conceptual studies focused on low ballistic coefficient deployable entry technologies/architectures for Venus, and make the case for going beyond rigid aeroshell architectures for future in situ science missions. These architectures do result in benign entry environments, perhaps similar to environments associated with typical Mars entries. The low ballistic coefficient architectures and the associated low deceleration loads open up the mission design space for EDL systems, may allow these missions to include sensitive and powerful science instruments, and allow for ASRG or other RTG based power systems that would allow for longer duration science. . The use of deployable entry system architectures requires critical new technologies including flexible TPS. Research efforts led by NASA Langley Research Center on inflatable concepts, and efforts led by NASA Ames Research Center on deployable concepts hold great promise for in situ science missions to Venus. Both these innovative concepts – inflatable and deployable – were originally proposed for landing large mass at Mars (primarily focused on human missions), and thus have crosscutting nature to be attractive for development for other planetary destinations as well. 1 Chief Technologist, Entry Systems and Technology Division, NASA ARC 2 Research Scientist, ERC, Inc, Aerothermodynamics Branch, NASA ARC 3 Senior Research Scientist, ERC, Inc, Aerothermodynamics Branch, NASA ARC 4 Senior Research Scientist, ERC, Inc, Aerothermodynamics Branch, NASA ARC

61 References

1. Venkatapathy, E.; Laub, B.; Hartman, G.J.; Arnold, J.O.; Wright, M.J., Allen, G., “Thermal protection system development, testing, and qualification for atmospheric probes and sample return missions,” Advances in Space Research. Vol. 44, no. 1, pp. 138-150. 1 July 2009. 2. Venkatapathy, E.; Laub, B.; Hartman, G.J.; Arnold, J.O.; Wright, M.J., Allen, G., “Selection and Certification of TPS: Constraints and Considerations for Venus Missions” IPPW-6 Atlanta Ga. 23-27 June 2008. 3. G. Landis and K. Mellott, "Venus Surface Power and Cooling System Design," Acta Astronautica, Vol 61, No. 11-12, 995-1001 (Dec. 2007). Presented as paper IAC-04- R.2.06, 55th International Astronautical Federation Congress, Vancouver BC, Oct. 4- 8 2004. 4. Bullock, M. A., Senske, D. A., Balint, T. S., Campbell, B. A., Chassefiere, E., Colaprete, A., Cutts, J. A., Gorevan, S., Grinspoon, D. H., Hall, J., Hartford, W., Hashimoto, G. L., Head, J. W., Hunter, G., Johnson, N., Kiefer, W. S., Kolawa, E. A., Kremic, T., Kwok, J., Limaye, S. S., Mackwell, S. J., Marov, M. Y., Ocampo, A., Schubert, G., Stofan, E. R., Svedhem, H., Titov, D. V., Treiman, A. H., 2008. NASA's Venus science and technology definition team: A flagship mission to Venus. B.A.A.S. 40, 32.08 ( http://www.lpi.usra.edu/vexag/venusSTDT/)

62

A COMPARISON OF INFLATABLE AND SEMI-RIGID DEPLOYABLE AERODYNAMIC DECELERATORS FOR

FUTURE AEROCAPTURE AND ENTRY MISSIONS

Reuben R. Rohrschneider, Jim Masciarelli, and Kevin L. Miller

Ball Aerospace & Technologies Corp., [rrohrsch,jmasciar,klmiller]@ball.com

ABSTRACT

With the successful flight of Inflatable Reentry Vehicle Experiment-II, the concept of using fabric-based aerodynamic decelerators has been demonstrated. This flight was a ballistic entry from a sub-orbital velocity. Now, with the imminent launch of the Mars Science Laboratory and its guided lifting aeroshell, the bar has been raised for all future aerodynamic decelerator systems. Once this technology has been successfully demonstrated, future science missions will demand precision landing from any entry system, including future deployable aerodynamic decelerators. This paper will compare and contrast the system performance and capabilities of the two main classes of deployable aerodynamic decelerator, the inflatable and the semi-rigid deployable, with the goal of showing equal or greater performance to the existing state-of-the-art rigid aerodynamic decelerator. The current state-of-the-art for Mars entry system uses a rigid aeroshell for hypersonic deceleration, a disk-gap-band parachute deployed near Mach 2.0, and either airbags or chemical propulsion for terminal descent. This entry architecture is representative of all the successful United States Mars missions flown to date, with a maximum entry mass less than 1000 kg (590 kg payload) and landed altitude of -1.4 km referenced to the Mars Orbital Laser Altimeter (MOLA). With the successful flight of the Mars Science Laboratory (MSL) the envelope will be extended to nearly 3000 kg entry mass (800 kg payload) and +2.0 km MOLA landing altitude. Braun and Manningi show that this is very near the limit of the current landing architecture, and that new technology will be needed for larger missions. One option for this new technology is the deployable aerodynamic decelerator, which prior studies have shown offers substantial mass advantages to rigid systems at Mars and other destinations with an atmosphere for both entry and aerocaptureii,iii. Furthermore, deployable systems promise a much broader range of landing altitudes and entry masses that support human exploration. There are multiple deployable aerodynamic decelerator concepts that can be divided into two primary classes: inflatables, and semi-rigid deployables. The two main classes of deployable aerodynamic decelerator have been compared for both aerocapture and entry at Mars using ballistic trajectories, and their entry system mass fractions were shown to be within 2%iv. With such similar mass performance, other metrics such as precision

63 landing capability, resistance to micrometeoroids, and operational flexibility should be considered when planning future technology investments in aerodynamic decelerators. This paper will draw from the results of the High Mass Mars Entry Systems study, the Aerocapture GN&C study, and other previously unpublished work performed at Ball over the past 3 years. The result shows that the inflatable and semi-rigid deployable configurations are quite closely matched. Given the comparable overall desirability of these systems, future studies should include both concepts to minimize risk while developing the next generation of aerodynamic decelerator systems.

i Braun, R.D., and Manning, R.M., “Mars Exploration Entry, Descent, and Landing Challenges,” Journal of Spacecraft and Rockets, Vol. 44, No. 2, pp. 310-323, 2007. ii Miller, K.L., et al, “Trailing Ballute Aerocapture – Concept and Feasibility Assessment,” AIAA Paper 2003-4655, 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Huntsville, AL, July 2003. iii Zang, T.A., et al, “Overview of the NASA Entry, Descent and Landing Systems Analysis Study,” AIAA Paper 2010-8649, AIAA Space 2010 Conference and Exposition, Anaheim, CA, Aug. 30-Sep. 2, 2010. iv Rohrschneider, R.R., “High Mass Mars Entry System Final Report,” Unpublished final report of contract NNL08AA34C, 2010.

64

EXOMARS 2016 – GNC APPROACH FOR ENTRY DESCENT AND LANDING DEMONSTRATOR

S. Portigliotti1, P.Martella1, M.Capuano1, O.Bayle2, T.Blancquaert2

Thales Alenia Space1, [email protected]

[email protected] [email protected] European Space Agency2, [email protected], [email protected]

ABSTRACT

This paper gives an overview of the most significant results, related to the Guidance Navigation and Control system design of the ExoMars Entry Descent and Landing Demonstrator Module (EDM) for the Exomars 2016 mission. Although the technologies to descend on a planet with a capsule are well known and experienced, landing remains a critical point for whichever exploration mission. Several solutions for landing technologies have been used in past missions, from the active braking with throttleable or pulsed rocket engines and impact attenuations legs (Viking, Phoenix), pulsed raking and un-vented airbags (Pathfinder, MER) of pure impact attenuation with un-vented airbags (Beagle-2). The new JPL-NASA missions use active control with throttleable engines and direct delivery to surface of rovers with the sky-crane concept. ExoMars Descent Module relies on the new technology of crushable structures for terminal impact attenuation that requires a precise control in the final instants, to be able to drop the lander at the specified altitude and with (nominally) null velocity and displacement versus the local vertical. Terminal braking is performed on Pulse Width Modulation of three clusters of three 400N engines, located directly on the Surface platform. For the ExoMars mission success it will be necessary that every GNC task will be perfectly achieved: the Entry Point recognition, the parachutes deployment trigger, the engagement of relative terrain navigation with hybridization of the inertial navigation with direct measurements via Radar Doppler Altimeter (RDA), the engagement and control of the terminal descent phase, the terminal drop of the Surface Platform to the surface of Mars. Looking in particular at the landing phase the ExoMars GNC has been designed trying to highlight some specific drivers: 1) Modular organization of the algorithm blocks based on functional roles (reference definition, state estimation and control action dispatching) and on affected axes (descent-vertical dynamics and attitude-horizontal ones), 2) Clear identification of the interconnections among the modules, 3) Definition of rules, simple as much as possible, to maintain continuously under control the evolution of each module dynamics and to force by construction adequate separation of the dynamics in the interconnected loops. For each of the three "G", "N" and "C" it is possible to identify a

65 critical aspect. The guidance must be able to compensate the engaging inaccuracy due mainly to radar Doppler scale factor errors and has to be designed in such manner to avoid jittering profile in the attitude reference generation, despite it is fed by noisy measurements. The navigation has the complex role to guarantee sufficiently filtered state estimation but, in the same time, high promptness peculiarly in the initial instants of the controlled phase that require fast attitude control. The control has to guarantee that the command is dispatched in the most effective way among the thrusters. The priority must be given to the attitude control in such a manner to achieve as soon as possible the alignment of the capsule to a direction opposite to the relative velocity (g-turn) also when starting from large attitude errors. Once the capsule has been aligned, accrued errors versus the descent profile can be recovered, ensuring, in the end, the fulfillment of both translational and rotational requirements. Last but not the least, a Backshell Avoidance Manoeuvre (BAM) has to be implemented in the cases where weak horizontal winds may induce the risk that the separated back- cover under parachutes may fall back onto the Surface Platform. The key aspect for a project like the ExoMars EDM GNC is the verification of the robust performances. The control must work in presence of strongly variable initial conditions, radar Doppler and actuators generates high level of inaccuracies to be carefully managed Furthermore even in presence of a modular design there are several points in which the loops are interconnected. Thus a big effort has to be spent for this task to have an analytical assessment of the design reliability, to be later confirmed by the execution of a large number of Montecarlo analyses.

66

THE MARS SCIENCE LABORATORY ENTRY DESCENT AND LANDING MODE COMMANDER

Paul Brugarolas, Kim Gostelow, A. Miguel San Martin, Fred Serricchio, and Gurkipal Singh

Jet Propulsion Laboratory California Institute of Technology Email:

[email protected]

ABSTRACT The Mars Science Laboratory (MSL) is the next NASA rover mission to Mars. It will be launched in November of 2011 and arrive in Mars in August of 2012. Its Entry Descent and Landing (EDL) phase is one of the most critical phases of the mission. It uses a highly complex system to land the vehicle safely within the desired landing region. The EDL system has three main components:

i. a timeline engine to prepare and coordinate all the events, ii. a Navigation Mode Commander to manage the estimation of the vehicle position

and orientation from the Descent Inertial Measurements Units and the Terrain Descent Sensor (radar), and

iii. an EDL Mode Commander to reconfigure the vehicle and guide-and-control the vehicle to a safe landing.

This paper will describe this last component. The EDL Mode Commander is the executive that orchestrates the hardware reconfigurations (balance mass ejections, heatshield and backshell jettisons, parachute opening) and the Guidance Navigation & Control functions (position and attitude estimation, entry guidance, RCS attitude control until powered descent starts, powered descent guidance, powered descent position and attitude control). We will describe the EDL modes of operation, the vehicle reconfigurations, the GN&C functions performed at each mode, and the navigational and temporal triggers used to transition between modes. Acknowledgement: The research described in this paper was carried out at the Jet Propulsion Laboratory, California Institute of Technology, under a contract with the National Aeronautics and Space Administration.

67

SUPERSONIC RETRO-PROPULSION FLIGHT TEST CONCEPTS

Ethan Post(1), Artem Dyakonov(2), Ashley Korzun(3), Ian Dupzyk(4), Jeremy Shidner(5), Arturo Casillas(6), Karl Edquist(7)

Jet Propulsion Laboratory(1), Email: [email protected] NASA Langley Research Center(2), Email: [email protected] Georgia Institute of Technology(3), Email:

[email protected] NASA Ames Research Center(4), Email: [email protected] NASA Langley Research Center(5), Email: [email protected] Jet Propulsion

Laboratory(6), Email: [email protected] NASA Langley Research Center(7), Email: [email protected]

ABSTRACT

Supersonic retro-propulsion (SRP) is an advanced entry, descent, and landing (EDL) supersonic decelerator technology that, if developed, could significantly increase landed mass capabilities at Mars. In the development of future mission concepts, NASA has recognized the need for advanced EDL systems, and the Agency has begun targeted funding for SRP technology development. SRP has been assessed to currently be at Technology Readiness Level (TRL) 2, “Technology concept and/or application formulated”. A roadmap has been developed for the maturation of SRP to TRL 6, at which point SRP is likely considered to be sufficiently mature for incorporation into a flight project. Wind tunnel testing, systems analysis, and computational fluid dynamics simulation efforts are under way. The work contained herein represents a focused effort to define Earth-based SRP flight testing concepts. These concepts compliment ground testing and analytical efforts and will play a critical role in the maturation of SRP to a viable flight system. Two sub-scale flight test concepts, both designed for launch on sounding rockets, are considered in detail for potential proof-of-concept testing of the SRP technology. The flight test is intended to demonstrate successful operation, from initiation through nominal operation, of a “hot” SRP system at conditions that replicate the relevant physics of the aerodynamic-propulsive interactions expected in flight. Major subsystem components sufficient to close a preliminary design are defined for each flight test concept, including: mechanical, propulsion, instrumentation, telecommunications, avionics, and power. Commercial, off-the-shelf components are utilized as much as possible in both concepts. Trajectory designs and analyses are performed to understand and optimize test conditions and vehicle parameters including thrust profile and initiation altitude. The analysis and design approach used to develop these flight test concepts are discussed in detail in this paper. Following definition of a set of flight test objectives and a set of mission-level requirements, preliminary trajectory analyses were completed. These

68 analyses assumed a sounding rocket platform due to a relative low cost and complexity as compared to alternative flight testing platforms. The results of these analyses indicated that a flight test vehicle capable of meeting mission-level requirements could be designed for launch on a sounding rocket, leading to the development of two point designs. These point designs provide more accurate mass distribution and thrust profile estimates than those used in the preliminary trajectory analyses. The mass estimates were found to be within payload mass limits of a Terrier-Improved Orion sounding rocket. Further analyses are planned that will advance the most favorable concepts to a higher maturity level in preparation for a proposal as a part of a flight test program. Both sounding rocket-based flight test concepts were found to represent viable options for SRP flight tests in that they: (1) demonstrate an SRP proof-of-concept in a flight environment, (2) replicate relevant SRP physics using a minimally integrated system, (3) collect data during flight within acceptable uncertainties to satisfy relevant TRL 5 achievement criteria, (4) demonstrate the ability to design, package, integrate, and test SRP subsystems, and (5) become a stepping stone to the more complex flight tests that will follow and reduce the associated risks.

69

MAXIMUM ATTAINABLE DRAG LIMITS FOR ATMOSPHERIC ENTRY VIA SUPERSONIC

RETROPROPULSION

No ̈el M. Bakhtian1, Michael J. Aftosmis2

NASA Ames Research Center1, Stanford University2

ABSTRACT

With manned missions on the horizon for Mars exploration, the ability to decelerate high-‐mass systems upon arrival at a planet’s surface has become a research priority. Supersonic retropropulsion (SRP), the application of jets facing into the freestream, is currently being studied as a candidate enabling technology. A model describing a significant drag augmentation mechanism based on shock manipulation was recently introduced; this proposed method of augmenting decelerative forces without directly relying on escalating thrust offers a deceleration technique independent of fuel use. This work proposes preliminary quantification of the benefits offered by this method of SRP-‐based flow control. We present an analytical method yielding estimates of maximum drag coefficients attainable through shock manipulation via SRP jets, establishing the feasibility of flow control via SRP as a Mars EDL technology. The analytical study examines the benefit of maintaining stagnation pressure through cascading oblique shocks as compared to a single strong normal shock. Comparisons with both computational and experimental data of blunt body flows validates the analytic method and shock physics assumptions. A family of CD−Mach curves and corresponding tables are generated for various shock structures. We then consider real gas effects, analyzing the consequence of varying specific heat ratio, γ = {1.2 − 1.4}, and apply an effective γ value to produce a Mars-‐specific set of CD−Mach curves. A theoretical maximum drag coefficient for realizable SRP shock structures is proposed at the conclusion of this study. By engineering SRP systems optimized for drag augmentation rather than raw thrust, fuel savings allow increased payload thus maximizing landable mass. This paper examines the feasibility of SRP-‐based flow control for high-‐mass planetary EDL by quantifying the drag afforded via this technique.

70

ROTARY WING DECELERATOR USE ON TITAN

Ted Steiner1, Larry Young2

Massachusetts Institute of Technology1, Email: [email protected] NASA Ames Research

Center2, Email: [email protected]

ABSTRACT

The ongoing Cassini mission to Saturn is considered one of the most successful international collaborations in the history of space exploration. The mission included the Huygens probe, which landed on the surface of Saturn’s largest moon Titan in 2005, generating a huge amount of scientific interest in further exploration of Titan. Huygens brought its power source with it in the form of batteries, which limited its operational lifetime to about six hours, nearly half of which was spent in atmospheric descent. Huygens’ success, combined with other recent findings, provide justification for a return mission to study Titan’s atmosphere and surface. A vehicle for such a return mission would greatly benefit from a descent system that can provide landing site selection, low-velocity touchdown, and power generation capabilities, while providing a platform for atmospheric research. A comparison of various atmospheric deceleration technologies based on their potentials for providing heading control, a soft landing, and power generation during descent, shows rotary wing decelerators (RWDs) to be of significant merit for applications on Titan. RWD systems use autorotating wings to slow down a vehicle in atmospheric descent. During the majority of the descent, the rotary wing spins freely at high rpm to store up energy. When the vehicle approaches the surface, the RWD system performs a “cyclic flare” maneuver, using the stored energy to generate the lift necessary for a soft landing. Similar systems are implemented on terrestrial helicopters as a safety mechanism. Vehicle heading control is achieved by adding fully articulated blades to the system, such as are used on most modern helicopters, rather than only the collective pitch control required for landing. Titan’s dense and highly extended atmosphere make it an ideal location for RWD applications. Because the entry vehicle will spend several hours transiting hundreds of kilometers during descent, a generator attached to the autorotating wing could generate significant power while also keeping the rotation speed within a safe operating range. Preliminary calculations show that for average descent velocities of 4 to 8 m/s and probe masses of 300 to 800 kg, power generation levels of 1 to 4 kW may be feasible. Achieving similar power levels using radioisotope thermoelectric generators (RTGs), which are most commonly proposed for missions to Titan, would involve power system

71 masses on the order of 500 kg or more. We propose a preliminary design for a rotary wing decelerator system for landing on Titan. Initial blade sizes, material selections, and power systems are presented. Additionally, we provide analysis of aerodynamic performance, landing speeds, allowable probe masses, and predicted power output. We also discuss the feasibility of extending such a system for applications on Earth, Venus, and Mars.

72

SMALL PROBE REENTRY INVESTIGATION FOR TPS ENGINEERING (SPRITE) (IPPW-8)

Daniel M. Empey1, Kristina A. Skokova2, Parul Agrawal2, Gregory T. Swanson2, Dinesh K. Prabhu2, Keith H. Peterson2 and Ethiraj

Venkatapathy3

Sierra Lobo, Inc. [email protected] 1, ERC, Inc. [email protected] 2, NASA Ames Research Center3, Ethiraj.Venkatapathy-

[email protected]

ABSTRACT

At the 7th International Planetary Probe Workshop a paper [1] was presented proposing a unique strategy for Thermal Protection Systems (TPS) testing designed to strengthen the ground to flight traceability of TPS qualification programs. The concept presented was to develop an inexpensive small scale test platform, i.e., small probes that are fully instrumented, that can be tested both on the ground (in arc-jet facilities) and in flight. This proposed paper presents the results of a small focused project that addresses this concept, showing how such small probes were designed and then tested at full scale in an arc jet. This is a paradigm shift from the traditional stagnation point testing to one of “test what you fly”, not only enabling traceability from ground to flight, but also enabling the assessment of practices/margins policies used in the design of the TPS of large scale entry vehicles such as Orion or MSL. This effort, called SPRITE (Small Probe Reentry Investigation for TPS Engineering) has demonstrated the feasibility of ground testing flight-sized (35 cm diameter) reentry bodies with two very successful tests of full-sized instrumented proof-of-concept articles in the NASA Ames Research Center Aerodynamic Heating Facility (AHF). The objectives of this effort were to design, manufacture and test the article, develop a flight-like data acquisition system, demonstrate data gathering capability, application of design tools and assessment of their fidelity. The SPRITE probe (a 90° included-angle sphere-cone, with a conical after-body) was designed to represent a vehicle that could be both an arc-jet test model as well as an actual reentry body. The probe was instrumented with TPS instrumentation plugs of the same design used on the MSL heat shield as well a number of back-face and internal thermocouples. Strain gages were also mounted on the TPS-protected aluminum structure in an attempt to determine thermo-structural response. Data from the sensors was collected by an internal data acquisition system as well as by the arc-jet facility. While it was not the intent of these tests to represent a specific mission, the models were tested at a heat flux (approximately 170 watts/cm2) representative of a reentry through the Martian

73 atmosphere. This paper will present an overview of the SPRITE project including: overall design of the probe, thermal analysis of the probe, design of an internal Data Acquisition System (DAS), Computational Fluid Dynamic (CFD) simulation of the test conditions, thermal-structural analysis, and results of the arc-jet tests. Acknowledgments: The present work was supported by the Entry Systems and Technology Division, NASA Ames Research Center and Contract No. NNA09DB39C to Jacobs Technology, Inc. References: [1] Howard, Austin R., Prabhu, Dinesh, K., Venkatapathy, Ethiraj, and Arnold, James, O.: “Small Probes as Flight Test Beds for Thermal Protection Materials” Proceedings of the 7th International Planetary Probe Workshop, Barcelona, Spain, 2010.

74

THE DEVELOPMENT OF A CO2 TEST CAPABILITY IN THE NASA JSC ARCJET FOR MARS REENTRY

SIMULATION

Steven V. Del Papa1, Leonard Suess2, Brian Shafer3

NASA JSC1, [email protected] ESCG2, [email protected], ESCG3, [email protected]

ABSTRACT

The Atmospheric Reentry Materials and Structures Evaluation Facility (ARMSEF) located at NASA Johnson Space Center is used for simulating the extreme environment experienced upon reentry for the development and certification of thermal protection systems (TPS). The facility supports a large variety of programs and was heavily leveraged for the certification and operational support of the TPS for the Orbiter and, more recently, the development of the heat shield for CEV. This paper will provide more detail into the heritage of the facility. Unique attributes of the facility include a modular aerodynamically stabilized arc heater and independently controlled O2 and N2 for the test gases. When combining the O2 and N2 in a 23:77 mass ratio respectively the Earth’s atmosphere is accurately simulated and via modification of this ratio the investigation of the effects of atomic oxygen on a material’s response is possible. In the summer of 2010 a development effort was started to add CO2 as a third independently controlled test gas such that, when combined with N2, opens up the possibility of accurately simulating a Martian reentry environment. This paper will discuss the test facility, especially the arc heater, in more detail. Initial testing involved relatively low concentrations of CO2 combined with N2 for the primary purpose of gathering data to answer two pressing safety concerns. The first being the rate of production of carbon monoxide (CO) within the ejector vacuum system. The main concern was that CO can be flammable and possibly explosive at high enough concentrations and pressures. The hazard control during the development phase involved the use of injecting N2 inside the test chamber diffuser to dilute and reduce the concentration of any and all CO present. A residual gas analyzer (RGA) was used to determine the relative amount of CO in the exhaust gas and provide a conversion rate of CO2 to CO. This paper will discuss in more detail the results of the RGA data and the calculated conversion rate. The second safety concern addressed is the possible formation of hydrogen cyanide (HCN) and cyanide (CN). HCN would primarily be present in the cooling water while the

75 CN would most probably condense onto the interior surfaces of the test chamber. Water samples and wipes of the test chamber surfaces were analyzed by an industrial hygienist for the presence of HCN and CN. His paper will discuss these results in more detail. Throughout this development effort measurements of the CO2:N2 flowfield were made with heat flux and pressure probes and with laser induced fluorescence (LIF) of the atomic oxygen. This paper will discuss these results.

SESSION 5 - Science Instrumentation

77

PAYLOAD OPTIONS FOR FUTURE ENTRY PROBE MISSIONS

Thomas R. Spilker

Jet Propulsion Laboratory, California Inst. of Tech, Email:

[email protected]

ABSTRACT Atmospheric entry probes can potentially address a wide range of science objectives that involve measurements by a wide range of instruments. Rarely is a mission budget unconstrained so science teams and mission designers can simply include every instrument that might be useful. Instead, careful investigation and instrumentation choices must be made to ensure a sufficient science return to justify a mission, while staying within finite project resource limits. Such decisions involve balancing many different resources on the spacecraft and within the project, and also the priorities of the science objectives that could be addressed. The priorities of science objectives, and thus the investigations and instrumentation needed to address them, vary greatly from destination to destination. For example, probes into the atmospheres of the giant planets place a premium on the origin of the solar system and the giant planets, with the dynamics and chemistry of such deep atmospheres at a somewhat lower priority; at Titan, there is more emphasis on current organic chemistry and the evolution of complex organic molecules, from their initial production high in the upper atmosphere to their eventual deposition on Titan’s surface. This presentation will summarize investigation options, and the instrumentation options for implementing them, at various potential atmospheric entry probe destinations in the solar system, with the exception of Titan lander and balloon instruments that another paper in this session will cover. Special attention will be devoted to destinations given high priority by the 2012 Planetary Science Decadal Survey (PSDS 2012), whose preliminary results are scheduled for release in early March 2011.

78

TITAN LAKE PROBE: SCIENCE REQUIREMENTS AND INSTRUMENTATION

J. Hunter Waite1, Tim Brockwell1, John Elliott2, Patricia Beauchamp2

Southwest Research Institute1, NASA Jet Propulsion Laboratory2

ABSTRACT

The scientific objectives of a Titan Lake Probe mission are: 1) to understand the formation and evolution of Titan and its atmosphere through measurement of the composition of the target lake (e.g., Kraken Mare), with particular emphasis on the isotopic composition of dissolved minor species and on dissolved noble gases, 2) to study the lake-atmosphere interaction in order to determine the role of Titan’s lakes in the methane cycle, 3) to investigate the target lake as a laboratory for both pre-biotic organic chemistry in both water (or ammonia-enriched water) solutions and non-water solvents, and 4) to determine if Titan has an interior ocean by measuring tidal changes in the level of the lake over the course of Titan’s sixteen- day orbit. The starting point for this study is the joint NASA ESA TSSM mission. Using this as a starting point we have revisited the scientific requirements and expanded them to include the possibility of a lake floater and a submersible. The driving requirements for the mission are: 1) the need to land on and explore the lake at depth while adequately communicating the data back to Earth via either direct to Earth or relay communications, 2) thermal design that allows sustained (>32 days) sampling of the 94K lake environment, and 3) a mass spectrometer inlet system that allows sampling of gas, liquid, and solids from the 94K environment. The primary payload is an analytical chemistry laboratory that includes an inlet system for sampling gas, liquid, and solids in and above the lake feeding two capable mass spectrometers that determine the organic and isotopic composition of the sampled materials. The instrumentation also includes a meteorological package that can measure the rate of gas exchange between the lake and the atmosphere, and a lake physical characteristics package that includes pressure and temperature measurements as well as sonar.

79

INSTRUMENTS FOR IN SITU TITAN MISSIONS

Patricia M Beauchamp1, Jonathan Lunine2

Athena Coustenis LESIA3, Peter Willis1, George Cody4, Kim R. Reh1

Jet Propulsion Laboratory-‐California Institute of Technology1 [email protected], [email protected], [email protected] , Dipartimento di Fisica, Università degli

Studi di Roma2, [email protected], Observatoire de Paris-‐Meudon3, [email protected], Geophysical Laboratory Carnegie Institute4, [email protected]

ABSTRACT Other than Earth, Titan is the only world in our solar system known to have standing liquids and an active “hydrologic cycle” with clouds, rains, lakes and streams. Cassini-Huygens has provided spectacular data and has provided us with a glimpse of the mysterious surface of Titan (Lebreton et al. 2009). However the mission will leave us with many questions that require future missions to answer. These include determining the composition of the surface and the geographic distribution of various organic constituents. The dense atmosphere and hydrocarbon lakes on Titan’s surface can be explored with airborne platforms and landed probes, but the key aspect ensuring the success of future investigations is the conceptualization and design of instruments that are small enough to fit on such platforms, and yet be sophisticated enough to conduct the kinds of detailed chemical (including isotopic), physical, and structural analyses needed to understand the history and cycling of the organic materials. In addition, they must be capable of operating at cryogenic temperatures while maintaining the integrity of the sample throughout the analytic process. Illuminating accurate chemistries also requires that the instruments and tools are not simultaneously biasing the measurements due to localized temperature increases. While the requirements for these techniques are well understood, their implementation in an extremely low temperature environment with limited mass, power and volume is acutely challenging. Over the last few years there have been a number of mission studies that involve either landing in a lake on Titan or circumnavigating Titan in a balloon (Coustenis et al. 2009; Titan Saturn System Mission Final and Joint Reports). Science teams have identified investigations on these platforms that require instruments to have high resolution and high sensitivity but be lightweight and low-power to minimize mass which can also reduce mission cost (Coustenis et al. 2011). The need for high resolution and sensitivity follows from an examination of the Cassini-Huygens data and understanding what is required to interpret the complex chemistry occurring in the atmosphere and on the surface. Novel instruments are required to determine environmental conditions at the surface, such as humidity and winds as well as probe the physical properties of the lakes. Advances in the technologies required for sampling the high latitude lakes - cryogenic

80 sample acquisition and sample handling - are also essential, as are techniques for sampling cryogenic aerosols and dune materials. Because of the plentiful supply of organic material and the environmental differences, in situ instruments developed for Mars are not suitable for Titan in situ missions. New instrument paradigms must be adopted for long term operation at 94K. Developing components operable in the extreme conditions found on Titan’s surface can simplify the design of the landed element or balloon platforms and reduce operational complexities. This presentation will discuss some of the instrument and sampling systems needed for these scientifically challenging investigations and point out some of the technologies which can enable new concepts for flight instruments to study the physical properties and surface chemistry of Titan.

Reference: 1. Lebreton, J-P., Coustenis, A., Lunine, J., Raulin, F., Owen, T., Strobel, D., 2009. Results from the Huygens probe on Titan. Astron. & Astrophys. Rev. 17, 149-179. 2. Coustenis, A., and 157 co-authors, 2009. TandEM: Titan and Enceladus mission. Experimental Astronomy 23, 893-946. 3. TSSM Final Report, 3 November 2008, NASA Task Order NMO710851 4. TSSM NASA/ESA Joint Summary Report, 15 November 2008, NASA Task Order NMO710851 5. Coustenis, A., Atkinson, D., Balint, T., Beauchamp, P., Atreya, S., Lebreton, J-P., Lunine, J., Matson, D., Erd, Ch., Reh, K., Spilker, T., Elliott, J., Hall, J., Strange, N., 2011. Atmospheric planetary probes and balloons in the solar system. J. Aerospace Engineering 225, 154-180.

81

SPACECRAFT-TO-SPACECRAFT RADIO LINKS INSTRUMENTATION FOR PLANETARY GRAVITY,

ATMOSPHERIC AND SURFACE SCIENCES

Sami W. Asmar

Jet Propulsion Laboratory, California Institute of Technology

ABSTRACT Tradition Radio Science techniques utilizing microwave links between spacecraft and ground stations have successfully lead to numerous discoveries. However, limitations on the received Signal-Noise-Ratio or geometrical coverage necessitate alternate observation configurations and new instrumentation. Spacecraft-to-spacecraft observations have significant SNR advantage over the traditional technique and can yield considerably improved geometrical coverage. These observations have been rarely carried out before because a special receiver is required onboard the spacecraft. One type of such open-loop receiver has been utilized on GRACE and will be utilized on GRAIL for precision measurements of the gravitational fields of the Earth and the Moon, respectively. A Different receiver type is onboard the New Horizons mission for an uplink occulation of Pluto’s atmosphere. Yet another prototype instrument onboard the Mars Reconnaissance Orbiter and has been used to demonstrate spacecraft-to-spacecraft radio science experiments with the Odyssey spacecraft. A new digital open-loop receiver specifically designed to meet the requirements of an occultation experiment has been prototyped for flight on the Europa Jupiter System Missions to the Jovian system, i.e., a Europa orbiter and a Ganymede orbiter. This instrument can be used to achieve multiple scientific including occultations of the atmosphere and ionosphere of Jupiter, occultations of the tenuous atmospheres and ionospheres of the Jovian satellites, occultations of the tenuous Jovian rings, and bistatic scattering from surfaces of the satellites. This paper will discuss the functional instrumentation under development as well as the potential achievable scientific investigations.

82

THE MARS MICROPHONE 2016 EXPERIMENT

D. Mimoun1, Jean-Pierre Lebreton2, and the Mars Microphone 2016

team3

Université de Toulouse, ISAE/SUPAERO 10, [email protected], ESA/ESTEC + EUROPLANET3 See http://bit.ly/MM2016 for the complete team list)

ABSTRACT

The Mars Microphone is a very simple and exciting experiment proposed in the frame of the ExoMars 2016 EDM Payload. Its primary objective is to achieve a world premiere during the short life of the ExoMars EDL payload: retrieve sounds from Mars. While built in Europe, with students involvement, the Mars Microphone will also strongly rely on the heritage of the previous Mars Microphone experiments, led by Berkeley SSL and the Planetary Society for the Phoenix, Mars Polar lander and NetLander missions. This experiment will therefore feature a unique combination of outreach, educational initiative and scientific objectives, particularly suited for the EDM payload context. Experiment configurations and scientific objectives The stringent resource constraints lead us to propose 3 possible configurations for the Microphone which will eventually depend on the possible on–board resources allocation.

Sound environment on the Martian surface A thorough synthesis of the expected sound environment for the Mars microphone was given by (William, 2001) for the Mars Polar Lander Microphone. Sound behaviour at the Martian surface is expected to be very similar to the Earth stratosphere, with an average atmospheric pressure between 6 and 8 mbar and a mean temperature about 240 K. In absence of in-situ measurement, main expected attenuation sources are classical and molecular absorption, but also the effect of the carbon dioxide viscosity. As a consequence of this, a strong attenuation is foreseen: most sounds in the human ear sensitivity window will not propagate over more than some dozen of meters. However,

83 the situation improves in the lower frequencies, and infrasounds, either related to dust devils or to other sources are expected to propagate over kilometre ranges. Expected signals Therefore, expected signals are due to the interaction between the lander structure and the Martian wind. Aeolian tones will be related to the main size of the lander and to the size of the lander elements exposed to wind (Curle, 1955). Noise level will be mainly related to atmospheric turbulence next to the lander (William, 2001). As we expect a sandy environment in the vicinity of the lander, the noise of the particles against the lander structure or directly against the microphone (depending on the wind direction) could also be monitored. A random activation of the microphone will therefore most likely bring back wind and saltation related noises. In addition, several less probable phenomena could also be witnessed, especially if the EDM operational scenario allows operating an automatic “switch on” triggered by a event of some intensity: dust devils, thunderstorms, asteroid impacts. Dust devils are known to generate both infrasounds (detectable over long ranges) and high frequency sounds (short ranges) Arnold et al (1976) reported dust devil activity in the audible range [2000 Hz] for Earth dust devils. The microphone has therefore a good chance of capturing such sounds, and, in its stereo version, to provide data on its trajectory. Melnik and Parrot (1998), as well as Mills (1977) also stated that dust storms could lead to lightning through cloud dust charging: an acoustic counterpart (thunder) may be detectable. Experiment Description and Heritage The design is on-purpose very simple, and following the previous design, relies primarily on a COTS component. It offers the required functionality together with the required low power consumption (150 mW), and a sufficient reliability for short life duration. In its baseline configuration, the Mars Microphone weights 50 g, and is composed of an electronics board enclosed in a 50x50x20 mm aluminum box. The microphone component is accommodated “outside”. A simple serial bus interfaces with the internal payload unit. The proposed design has its heritage in previous Mars Microphone implementations, first on-board Mars polar Lander, and then on-board Phoenix: same microphone elements, same class of COTS components. Preliminary Team description, Student Involvement- Outreach The Mars Microphone 2016 team includes a wide panel of scientists and engineers, interested in both science and outreach. Outreach will be coordinated with the Planetary Society and Europlanet. Following the successful example of Cassini-Huygens, we will put a large emphasis on outreach activities. The strong design heritage of previous versions will also allow us to have a student involvement in the development and in the tests of the Mars Microphone 2016. [1] Ryan, J. and R. Lucich Possible dust devils, vortices on Mars. J. Geophys. Res. 88, 1983 [2] http://marsrovers.jpl.nasa.gov/gallery/press/spirit/20050819a.html, retrieved 01/2011 [3] Williams, JP Acoustic Environment of the Martian Surface, JGR, vol 106, 2001 [4] Curle, N The influence of solid boundaries upon aerodynamic sound, Proc.R. Soc.London.Ser A 231 505- 514, 1955

84 [5] Melnik, Parrot Electrostatic discharge in Martian dust storms, JGR 103, 1998 [6] Arnold, R. T., H. E. Bass, and L. N. Bolen, Acoustic spectral analysis of three tornadoes, J. Acoust. Soc. Am., 60, 584-593, 1976. [7] Mills, A. A., Dust cloud and frictional generation of glow discharges on Mars, Nature, 268, 614, 1977 [8] Vérant, J. L., Exomars Capsule Aerodynamics Analysis, 10th AIAA/ASME Joint Thermophysics and Heat Transfer Conference, Chicago (IL), July 2010 [9] Gnoffo, P. A., Prediction and Validation of Mars Pathfinder Hypersonic Aerodynamic Data Base, AIAA paper 98-2445, 7th AIAA/ASME Joint Thermophysics and Heat Transfer Conference, Albuquerque (NM), June 15-18, 1998

85

LIDAR INSTRUMENT FOR GLOBAL MEASUREMENT OF MARS ATMOSPHERE

Farzin Amzajerdian1, George Busch2, Norman Barnes1, Robert Tolson3, and Diego Pierrottet2

NASA Langley Research Center1, Email: [email protected], Coherent Applications, Inc.2, Email:[email protected]

National Institute of Aerospace3, Email: [email protected]

ABSTRACT Future NASA missions to Mars must focus on either obtaining high value scientific data or serve as precursor in preparation for human exploration. Laser-based instruments (lidars) can play a key role in achieving both of these objectives. Lasers offer clear advantages over passive and active radio-wave measurements. These advantages include excellent spatial resolution, nonreliance on natural light sources, and the ability to aim and scan. This paper proposes a multi-functional lidar instrument providing all critical atmospheric data while meeting the stringent mass and power constraints of a Mars mission. We refer to this lidar as "coherent Doppler/DIAL lidar" since it combines the attributes of a “coherent Doppler lidar” with those of a “Differential Absorption Lidar”. As an orbiting instrument, this lidar will provide global measurements of atmospheric winds, density, and aerosol with a high degree of precision and spatial resolution. The wind velocity is measured using the Doppler frequency shift of laser light scattered from suspended aerosols transported by the winds. The atmospheric density is determined from measurements of the concentration of CO2, which constitutes about 97% of Mars atmosphere. The concentration is determined by measuring the ratio of transmitted intensities of two different wavelengths emitted by the lidar, corresponding to high and low CO2 transmission. The CO2 concentration is profiled along the entire path of the laser beam. The aerosol concentration is simply derived from the intensity of the returned signals. Presently, individual lidar systems measuring Earth atmospheric winds and CO2 exists in the form of ground and airborne-based scientific instruments. Lidar aerosol measurement is more mature as several instruments have been successfully deployed to Earth orbit since 1994. The multifunctional lidar being proposed combines the functions of each individual sensor into a single device resulting in a more robust instrument with fewer components, and thus greater reliability, as well as reduced mass, volume, and power compared with multiple systems to handle each function. This lidar takes advantage of the fact that the aerosol concentration in Mars atmosphere is almost 2 orders of magnitude greater than that of the Earth thus permitting smaller lasers and smaller transmitter/receiver telescope apertures. The lower laser pulse energy required for Mars

86 allows for use of a novel highly-efficient near-infrared laser currently under development at NASA LaRC. This laser has much improved performance characteristics compared with current laser technologies available. That simplifies the instrument thermal management design and significantly reduces the overall payload mass and power consumption. This paper will describe the lidar instrument concept and its potential for providing global measurements of Mars atmospheric parameters. The instruments trades and limitations will also be discussed. Finally, the current state of the technology will be presented, along with a plan for advancing its readiness towards deployment in Mars orbit.

87

MARTIAN SONIC ANEMOMETER

Don Banfield

Cornell Astronomy, Email: [email protected]

ABSTRACT We have developed a 3-D sonic anemometer for use on Mars that exceeds the performance of previous martian wind gauges by at least an order of magnitude in sensitivity as well as sampling rate. This improvement in performance is important because it opens up new avenues of research at Mars, in the interaction of the surface and atmosphere. Fast response and sensitive wind measurements allow the direct measurement of the turbulent eddies in the martian atmospheric boundary layer. By correlating these turbulent motions with their associated vertical wind, temperature, humidity or other trace gas perturbations, we can directly measure the exchange of momentum, heat, water or other tracers with the surface. Additionally, if discrete sources of biogenic effluents are found at Mars, similar approaches may be useful to guide a rover to the precise source of the effluent plumes using a technique known as plume tracing, akin to how lobsters (among other animal predators) hunt their prey. Sonic anemometers are the gold standard for similar boundary layer studies on Earth, but terrestrial instruments are non-functional in the extremely low density atmosphere at Mars. Our instrument uses novel micro- machined capacitive transducers that more efficiently couple with the low acoustic impedance martian atmosphere, thus retaining as much acoustic signal strength on transmission under martian conditions as possible. These transducers have been specifically optimized for use on Mars including durability to the extreme temperatures, optimization for lower spacecraft power availability and miniaturization. The remainder of the instrument is a mix of analog and digital electronics to produce the acoustic signals and then process them to yield wind speeds and temperatures. The signal processing involves sophisticated algorithms borrowed from the field of RADAR to extract as much information content as possible from the signals. We are currently finishing development on this instrument using PIDD funding, and are testing it in a thermal vacuum chamber at Ball Aerospace. We are also in the process of preparing for a stratospheric balloon flight (which mimics martian surface conditions to a high degree) to raise our instrument’s TRL to between 5-6. Our performance goals are 3-D wind measurements with sensitivity down to better than 10 cm/s, an accuracy of ~10 cm/s, and with a sampling rate of 20 Hz or more. In a full flight configuration, the instrument is expected to draw ~2W while operating and 0W when quiescent, with an instantaneous startup. It should total about 1kg in mass and stow into a volume of about 1500 cm3. We are eager to include this instrument on any and ALL future landed or aerial missions to Mars. We are quite confident that it will open up exciting new avenues of research at Mars. It may also prove to be a valuable instrument at Titan.

88

THE CHEMCAM INSTRUMENT FOR THE 2011 MARS SCIENCE LABORATORY MISSION: SYSTEM

REQUIREMENTS AND PERFORMANCE

R. Perez1, B.L. Barraclough2, S.C. Bender2, A. Cousin3, A. Cros3, N. Le Roch4, S. Maurice3, A. Paillet1, L. Pares3, Y. Parot3, M. Saccoccio1, R.C.

Wiens2

Centre National d’Etudes Spatiales1, Email: [email protected] (2)Los Alamos National Laboratory2, IRAP, CNRS3, ALTEN Sud Ouest4,

ABSTRACT

The ChemCam experiment is one of ten onboard the Mars Science Laboratory (MSL) rover “Curiosity”, currently scheduled for launch in late 2011. The instrument is a combination of a Laser-Induced Breakdown Spectrometer (LIBS) and a Remote Micro-Imager (RMI) camera. The LIBS subsystem will provide remote sensing (up to ~7m range) data on the composition and elemental abundances of rocks and soils via active interrogation by a high-power laser. It is also possible to obtain passive spectra of targets using the LIBS subsystem and natural illumination. The RMI subsystem provides high-resolution images of the target regions interrogated by the LIBS laser, and will be used to provide geologic context for the LIBS data. This is the first use of a LIBS system in space. ChemCam is physically divided into two separate units: the Mast Unit (MU) and the Body Unit (BU). The MU is located at the top of the rover mast, ~2 meters above ground level, and consists of an optical telescope, a Nd/KGW laser, the RMI camera and supporting electronics. The MU is provided by the French part of the ChemCam team, IRAP laboratory, supported by CNES, the French Space Agency. The Body Unit consists of an optical demultiplexer, three independent spectrometers with CCD detectors, the experiment controller, and supporting electronics, and is located inside the body of the rover. The BU is provided by Los Alamos National Laboratory. The MU and BU are interconnected via fiber optic and electrical cables, both contributed by Jet Propulsion Laboratory. In order to excite small areas of geologic targets to temperatures high enough to radiate photons that can be analyzed by the LIBS subsystem, laser power densities of > 1 GW/cm2 at the sample are needed and, for ChemCam, these densities need to be achieved over distances ranging from 1-7 m from the rover mast. Other laser requirements for successful LIBS analyses include laser spot diameters of 200 – 600 µm over the given range, pulse energy at the sample > 13 mJ, pulse durations of ~5 ns and

89 beam quality of M2 < 3. The ChemCam 110 mm diameter telescope must perform three distinct functions: it must direct and focus the intense laser output (λ=1067 nm) on targets over the required range, it must efficiently collect the photons (λ range = 240–870 nm) emitted by the plasma cloud generated at the sample by the laser and transmit (efficiency 15-40%) this light to the remotely-located spectrometers, and it must act as specialized “telephoto” lens for the RMI subsystem. The RMI camera itself must provide < 100 µrad resolution to enable adequate imaging of the interrogated samples. The entire optical subsystem must be capable of auto-focusing very precisely over the required operational range. The optical demultiplexer subsystem of the BU must efficiently divide the LIBS photons collected by the telescope into three optical bands (UV = 240-340 nm, VIS = 385-465 nm and VNIR = 475-870 nm) and feed these photons to the three spectrometers that are optimized for their respective wavebands. The spectrometers are required to achieve optical resolutions of 0.2, 0.2, and 0.65nm (FWHM) for UV, VIS and VNIR respectively and the wavelength drift with temperature should be < 0.1 pixel/C. Extensive testing at the subsystem, integrated-instrument and integrated-system level show that all performance requirements are met over distance, temperature, etc. and that ChemCam is fully capable of achieving its science goals when it lands on the surface of Mars in August, 2012. This talk will detail the performance requirements that need to be met for successful ChemCam operation, the testing that has been performed to date to insure these requirements are being met and the overall instrument performance that can be expected on the surface of the Red Planet.

90

MEADS CALIBRATION AND MSL TRAJECTORY RECONSTRUCTION

Mark Schoenenberger1, Chris Karlgaard2, Michelle Munk1

NASA-Langley Research Center1, Email: [email protected],

[email protected], Analytical Mechanics Associates2, Email: [email protected]

ABSTRACT The Mars Science Laboratory (MSL), launching in late 2011, will be outfitted with a pressure measurement system (MEADS, the Mars Entry Atmospheric Data System), which consists of 7 ports in a cross configuration, much like air data systems used at Earth. This will be the first time such a system has been flown at Mars, and the amount and quality of data return is expected to be orders of magnitude better than that from any previous Mars mission. The MEADS objectives are to measure forebody pressure distribution, to determine angles of attack and sideslip, and to separate atmospheric density and winds from the vehicle aerodynamics. Four years in the making, the MEADS hardware is now complete and delivered to MSL. The pre-delivery calibration of the end-to-end system has resulted in hardware that meets its performance requirements. However, the ability to meet the final objectives will be influenced by a wide range of vehicle and environmental uncertainties. The Reconstruction team, planned to function beyond MSL entry and into 2013, will use a combination of ground testing and computational methods to understand and cross- correlate all of the relevant data from the day of entry. This presentation will focus on the pressure system calibration methods and results, as well as the current and planned work to ready the engineers to receive the flight data.

91

OPTICAL EMISSION SPECTROSCOPIC EXPERIMENTS FOR IN-FLIGHT ENTRY RESEARCH

Sebastian Lein1, Georg Herdrich1, Monika Auweter-Kurtz2 and Stefanos Fasoulas1

Institut für Raumfahrtsysteme (IRS), Universität Stuttgart1, Email: [email protected]

stuttgart.de, [email protected], [email protected], German Aerospace Academy ASA2, Email: [email protected]

ABSTRACT In the proposed paper, based on the emission spectroscopic payload RESPECT [1], developed for the European re-entry vehicle EXPERT [2], the applicability and benefits of emission spectroscopic payloads as part of the scientific instrumentation of re-entry vehicles will be discussed. Moreover, possible further development stages, enhancing the operational range and/or improving the scientific output of the instrument will be presented. The instrumentation of re-entry vehicles with emission spectroscopic payloads is motivated by the significant interaction of the plasma state of post shock regime and boundary layer with the thermal and mechanical loads on the heat shield surface. Especially for re-entry missions in CO2 dominated atmospheres, as well as re-entry missions to the giant planets, the radiative heat flux contributes significantly to the total heat flux on the TPS surface. Information on the plasma state can be obtained by emission spectroscopic measurements. Although various numerical codes have been developed to simulate these conditions, the experimental data which can be used to verify the numerical simulations are still poor. Thus, in-flight measurements are most valuable to increase the reliability of the current data base and therewith the design base for future missions. In the past years the payload RESPECT was developed at the Institut für Raumfahrtsysteme (IRS) to serve this purpose. Development, assembly and qualification of the payload for application on the European re-entry mission EXPERT were successfully completed and currently the flight model of the re-entry capsule is assembled. Thus, the focus of the payload related activities changes towards the preparation of the flight data analysis and the expected scientific output. In order to judge the scientific output expected from the application of RESPECT on EXPERT, numerically simulated spectra were generated. These spectra have been calculated based on flow field simulations of several trajectory points, using the URANUS code [3], and superimposed radiation simulations using the plasma radiation data base PARADE [4]. In order to generate simulated spectrometer data sets the numerical radiation data was convoluted with the optical properties of the payload gained

92 from laboratory experiments to characterize the instrument [5]. The expected results comprise among others the identification of the radiating plasma species including possible erosion products originating from the heat shield material. In addition, the occurrence of active oxidation of the ceramic heat shield can be traced on basis of the detected erosion products. Beyond that, from the numerical rebuilding of the measured spectra, the locally resolved plasma composition including particle densities and excitation temperatures shall be determined. The data is expected to allow for an accuracy analysis of the current simulation tools and the improvement of the employed chemistry and radiation models. In this paper, based on the experience gathered in the development of the RESPECT sensor system an outlook on possible further development stages will be given. This includes the enhancement of the operational range as well as design improvements to maximize the scientific output. The payload was developed for the re-entry of EXPERT into Earth’s atmosphere but is also suitable for other planets. The limits of the sensor system with respect to the operation in other atmospheres and possibly required design modifications will be discussed. In addition, design improvements for future emission spectroscopic payloads, such as the application of other spectrometer types, and their impact on the measurement data and scientific output will be presented. References 1.Lein, S., Seinbeck, A., Preci, A., Fertig, M., Herdrich, G., Röser, H.-P., Auweter-Kurtz, M., Final Design and Performance Parameters of the Payloads PYREX, PHLUX and RESPECT on EXPERT, Transactions of the Japan Society for Aeronautical and Space Sciences, Aerospace Technology Japan, Vol. 8, (2010) pp.Tm_41-Tm_47 . 2.Muylaert, J., Walpot, L., Ottens, H. and Cipollini, F., Aerothermo- dynamic Reentry Flight Experiments Expert, In Flight Experiments for Hypersonic Vehicle Development, Educational Notes RTO-EN-AVT-130., Neuilly-sur-Seine, France, 2007, pp. 13-1 – 13-34. 3.Fertig, M. and Herdrich, G., The Advanced URANUS Navier-Stokes Code for the Simulation of Nonequilibrium Re- entry Flows, Proceedings of the 26th. International Symposium on Space Technology and Science, Hamamatsu, Japan, 2008. 4.Winter, M., Pfeiffer, B., Fertig, M. and Auweter-Kurtz, M., Extension of PARADE to CO2 Plasmas and Comparison with Experimental Data in High Spectral Resolution for Air and CO2 Species, Proceedings of the 1st. International Workshop on Radiation of High Temperature Gases in Atmospheric Entry, Lissabon, Portugal, 2003, (ESA SP-533, December 2003). 5.Lein, S., Preci, A., Fertig, M., Herdrich, G. and Auweter-Kurtz, M.: Optical Design and Layout of the In-Flight Spectrometer System RESPECT on EXPERT, AIAA-2009-7263, 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference, Bremen, Germany, October 19-22, 2009.

SESSION 6A - New Technologies

94

PEDALS: EVOLVED DESIGN OF EDL ARCHITECTURES

Ed Chester, João Graciano

Aevo GmbH, Friedrichshafenerstr. Email: [email protected]

ABSTRACT Following a number of studies in the use of meta-heuristic methods for space applications, we present a simple development of a hybrid genetic algorithm that can ‘design’ EDL architectures and implicitly perform trade-off studies. To focus the application of such a tool, a secondary objective was to implement a prototype system that could possibly identify concepts for high-altitude landing on Mars, each of which would then need further study using more conventional analysis. Any EDL architecture representable within our study is captured as a list of parameters that includes objective inputs (e.g. landed system mass, location), physical measurements (e.g. parachute diameter), times (e.g. of deployment, release, etc.), ordered sequences, and derived parameters (e.g. instantaneous descent rate). Given the limited availability of the POST tools, this short research activity is a case study in the effectiveness of simple low-cost tools, and establishing whether evolutionary algorithms have any place in the mission designer’s toolkit. The resulting PEDALS tool concept is a direct coupling of a trajectory simulator with a customised genetic algorithm, implemented in Perl and F90. The trajectory simulator is known to be a coarse approximation, and does not consider variable bank angle or 6DOF effects. The modular design of the tool has however allowed the incorporation of MOLA data for altitude targeting, and Mars-GRAM for more realistic atmospheric flight simulation. The results are compared with prior EDL architectures and with the NASA Entry, Descent and Landing Systems Analysis (EDL-SA) 2008 study as presented at IPPW-7, which are adopted as a family of reference designs. The presentation concludes with recommendations for future steps in this type of work.

95

CHALLENGES OF THE INSTRUMENTATION FOR HIGH SPEED ENTRY VEHICLES

Ali Gülhan, Frank Siebe, Thomas Thiele German Aerospace Center (DLR), Supersonic and Hypersonic Technology Department, Linder Höhe, Email: [email protected], [email protected], [email protected]

ABSTRACT

In order to achieve a low cost access to space while keeping the reliability at a high level, the design cycle duration has to be reduced and the service and refurbishment of the space vehicle have to be simplified. Design tools like CFD or structural codes are to be improved with respect to physical modelling using accurate data from ground testing or flight experiments. Although ground testing facilities still provide the main validation data and allow a better understanding of the physics, a complete duplication of the flight conditions is mostly not possible. Flight experiments are the only way to obtain validation data for design or prediction tools under real conditions. On the other hand during flight experiments only coupled information can be gained and therefore a parametric study is not possible. Therefore a further use of ground testing facilities and CFD simulations for post flight analysis is essential to interpret the flight data correctly. For vehicles using an ablation material for the TPS the instrumentation is more difficult. One of the key problems is the strong contour change of such materials resulting from thermal expansion and recession. In addition the phase change inside the material leads to a significant modification of the material properties and makes the determination of the thermal properties of the structure more difficult. Ablation products in gas, liquid and solid form enhance this problem. These phenomena dominate the behaviour of the capsule front surface, which is exposed to very high aerothermal loads. On the rear surface the convective heating is low but difficult to estimate. This is a result of the shortcomings of numerical tools. In addition for some atmospheres, like the Martian atmosphere, the radiative heating on the base could reach the same level as the convective heating. To measure these phenomena in flight experiments a dedicated sensor has to be designed. The COMARS sensor of DLR has been developed to have a combined measurement of pressure, temperature, heat flux rate and radiation at the base of the capsule during a Martian entry (Figure 1).

96

97

SYSTEM DEVELOPMENT FOR MARS ENTRY IN-SITU RESOURCE UTILIZATION

Svetozar Popovic1, Robert W. Moses2, Leposava Vuskovic3

(1)Old Dominion University, 4600 Elkhorn Avenue, OCNPS, 306, Norfolk, VA 23529, USA, Email: [email protected]

(2)NASA Langley, 1 N Dryden St. (MS489), Hampton, VA 23681, USA, Email: [email protected]

(3)Old Dominion University, 4600 Elkhorn Avenue, OCNPS, 306, Norfolk, VA 23529, USA, Email: [email protected]

ABSTRACT NASA's Game Changing Development Program will require new power systems and in situ resource utilization (ISRU) technologies. Fortunately, there is an abundance of basic research that could serve as a basis for such a development. This work is motivated by the major contradiction in the concepts for power systems for Mars exploration. While the surface exploration relies on very limited solar power resources that have reduced the range of applicable solutions, the space vehicle itself has huge amount of power stored in the form of its orbiting kinetic energy. This energy is currently not utilized, but rather left dissipated through heat transfer and radiation. This unused potential is not to be neglected, since it may offer the opportunities for the use the entry plasma as a powerful resource that still remains to be utilized. In this paper we report the on the state of effort to characterize Mars entry plasma as a potential work fluid for on–board power generating systems, and a chemical reactor medium for oxygen generation. The use of Martian entry plasma can be augmented using the concept of regenerative aerobraking1 may offer a revolutionary approach for in situ power generation and oxygen harvesting during the exploration missions. The on-board power conversion system concept is based on a network of lightweight magnetohydrodynamic power generators developed in NASA LaRC and at ODU2. The system technology would capture energy and oxygen from the plasma field that occurs naturally during hypersonic entry using well understood principles of magnetohydrodynamics and oxygen filtration. This innovative approach generates resources upon arrival at the operational site, and thus greatly differs from the traditional approach of taking everything you need with you from Earth. Fundamental analysis, computational fluid dynamics, and some testing of experimental hardware have established the basic feasibility of generating power during a Mars entry. This system is an example how regenerative aerobraking may be applied to support human and robotic missions at Mars. The system consist of several subsystems that would address the Oxygen production and storage, utilize MHD cooling of thermal shield, provide power by MHD conversion for fluid cooling subsystem and heat redistribution to a resistive load to the rear of the

98 spacecraft. Detailed description of these systems is the subject of present talk. Oxygen production, separation and storage is based on two alternative solutions for separation, (a) YSZ Solid Oxide Electrolysis Cell, or (b) Silver membrane extraction. Both solutions have been tested and validated. An inflatable container is being developed for oxygen storage and the analysis of its performance will be given. Power conversion systems will be based on the planar MHD power conversion unit, presently operating using light-weight rare earth permanent magnets, but a concept using light weight electromagnets has also been developed. MHD cooling effect was confirmed in CFD simulations and an optimum distribution of magnets is evaluated. Additional fluid cooling system is based on two alternative approaches using (a) gas or (b) liquid metal as a cooling fluid. Electric circuit for MHD power redistribution and effective heat transfer to the cold region of the vehicle is described and analyzed as a separate subsystem. 1 Moses, R.W., “Regenerative Aerobraking,” Space Technology and Applications International Forum (STAIF) 2005, Paper No. 57, 13-17 February 2005, Albuquerque, New Mexico. 2 Vuskovic, L., and Popovic, S., “Magnetohydrodynamic Power Generator,”Summary of Research Report for ODURF Project #133931, March 2004.

99

TERMINAL DESCENT AND LANDING SYSTEM ARCHITECTURES FOR A MARS PRECISION LANDER

Lisa Peacocke1, Marie-Claire Perkinson1, Jaime Reed1, Tobias Lutz2, Marco Wolf2, Joerg Boltz2

Astrium Ltd1, Email: [email protected] Astrium GmbH, Airbus-Allee2 Email: [email protected]

ABSTRACT

A Mars Precision Lander mission is currently being studied under ESA contract. A landing accuracy of better than 10 km is required, with a goal of 7.5km, which is significantly better than past Mars missions. A potential mission scenario considered for the precision lander is the landing of a Sample Fetch Rover. This rover would retrieve the sample cache obtained by NASA’s Mars Astrobiology Explorer-Cacher (MAX-C) rover and place it in the Mars ascent vehicle within the overall Mars Sample Return mission architecture. A precise landing is non-trivial, and requires a highly accurate guided entry and likely a powered descent phase with potential hazard avoidance. The critical terminal descent and landing phases will safely deliver the fetch rover to the Martian surface, and have been studied in detail in the first phase of the MPL contract. To protect the fetch rover, a maximum surface impact velocity of 1.5 m/s is specified, and the system must be able to land in an area with 99% areal density of 60 cm rocks and 22.5° surface slopes. A maximum horizontal wind velocity of 20 m/s and maximum vertical wind velocity of 5 m/s must be dealt with. A minimum of two egress paths must be available to the rover, and egress must be autonomous and highly reliable and robust. A variety of architectures for the terminal descent and landing are possible and have been investigated, including:

• Legged landers • Airbags – vented or unvented • Crushable structures • Dropship (Skycrane-type) • Shell lander (Beagle-2 type) • Parafoil/aerobot with control platform

The surface rocks and slopes strongly drive the architecture design. A hazard avoidance system is one option, otherwise the system must be able to land safely in the worst case scenario – a combination of a 60 cm rock and 22.5° slope. Self- righting mechanisms, such as jointed or extendible legs, are an option for legged landers in this case. Vented airbags are preferable to unvented airbags due to the precision landing requirement – unvented airbags can bounce for a large distance before stopping, potentially violating the precision landing requirement. The safe egress of the rover is highly interlinked with the terminal descent and landing architecture, and would require complex ramps, roll-out platforms or cranes with a legged lander or airbag system. A Dropship would enable a

100 simple egress via a winch and cables, and would avoid the thruster plume and back pressure issues associated with Viking-type landers. This paper will summarise the terminal descent and landing architecture concepts and trade-offs investigated in the first half of the Mars Precision Lander contract. The advantages and disadvantages of each will be outlined, particularly in regard to a precision landing, and the preferred concepts will be identified.

101

Overview of Hypersonic Inflatable Aerodynamic Decelerator Large Article Ground Test Campaign

Alan M. Cassell, Gregory T. Swanson, R. Keith Johnson, Stephen J. Hughes, F. McNeil Cheatwood

ERC Inc. NASA Ames Research Center, NASA Langley Research Center

ABSTRACT

Hypersonic inflatable aerodynamic decelerators (HIADs) offer considerable advantages over rigid aeroshell technology for human and robotic missions requiring atmospheric entry. Most noteworthy are the considerable system mass and volume fraction savings over conventional rigid aeroshells. In addition, inflatable aeroshells Currently, HIADs are being considered for returning payloads from low earth orbit and landing heavy payloads on the surface of Mars. The Inflatable Re-entry Vehicle Experiment (IRVE) has successfully demonstrated various aspects of HIAD technologies including exo-atmospheric inflation, inflatable structure performance, thermal protection system performance, aerodynamic stability and structural integrity under aerodynamic pressure. IRVE-II, flown in 2009, a 3.0 meter diameter, 60 degree half-angle sphere cone enabled the validation of a number of design tools and approaches for inflatable decelerator technology. Scaling HIADs to the diameters relevant to the aforementioned entry missions (>10 meter diameter) presents unique challenges for validating the performance and design of such systems. There are many unquantified risks to the utilization of such large structures, such as control authority, fluid structure interactions, dynamic stability and system complexity. Understanding, developing and validating larger diameter HIAD designs will require an extensive ground testing campaign. The National Full-Scale Aerodynamics Complex (NFAC) at NASA Ames Research Center is a unique facility primarily used for determining aerodynamic characteristics of large-scale and full-scale rotorcraft and powered-lift V/STOL aircraft, as well as testing of wind turbines, parachutes, trucks, and other non-traditional types of testing. The facility is composed of two large test sections and a common, six-fan drive system. The 40-by-80 foot wind tunnel circuit is capable of providing test velocities up to 300 knots. This paper discusses the objectives, planning and challenges in testing large diameter (up to 8.5 m) HIADs in the NFAC 40 x 80 foot test section. An overview of the design reference mission, key driving requirements, structural analysis, instrumentation development and flexible aeroshell structural model validation approach will be presented. In addition, failure mode testing approaches will be presented to build further confidence in developing HIAD technology for infusion into near term flight demonstration missions.

102

Figure 1 – 6 meter diameter HIAD test article concept placed in the 40 x 80 foot test section of the NFAC.

103

LOW-DENSITY SUPERSONIC DECELERATOR SYSTEM

Mark Adler, Chuck Player, Juan Cruz, Ian Clark, Adam Steltzner, and Tom Rivellini

ABSTRACT

The heady days of sticking a rocket under your gizmo just to see if it works are coming back. In the summer of 1972, the Viking Project conducted four high-altitude tests in Earth's atmosphere of the supersonic parachute design to be used for landing on Mars. We've been stuck with that design ever since. In the spring and summer of 2013, a series of four balloon-launched, rocket-propelled tests of the next generation of supersonic decelerators will culminate the development program of new descent technologies to be used on future Mars landers. A Mach 4 inflatable decelerator and a Mach 2 ringsail parachute will team to create a low-density supersonic decelerator system for high ballistic coefficient entry systems at Mars.

104

CO2 PROPULSION FOR A MARS SURFACE HOPPER

Christopher Perry & Robert L. Ash

Mechanical and Aerospace Engineering Dept., Old Dominion University Norfolk, VA

ABSTRACT Spirit and Opportunity (MER) rovers operating currently on the surface of Mars have returned valuable data about the liquid water history of the planet, but are very slow and cannot traverse difficult terrain. The lack of mobility over difficult terrain limits surface exploration to features that lie on relatively flat surfaces. By exploiting Mars’ reduced gravity, a system that can fly over difficult terrain can provide the ability to achieve science objectives over a much wider range of potential targets. A hopper vehicle that can utilize ballistic flight trajectories over rugged terrain and reach otherwise inaccessible destinations can greatly expand mission capabilities. By heating frozen CO2, condensed from the Martian atmosphere, a supercritical fluid can be created at modest temperatures within a pressure vessel and subsequently used as rocket propellant. A supersonic carbon dioxide rocket motor, operating in blow-down mode, has been designed, built, and tested for the purpose of evaluating this type of hopper propulsion system. The operation of the propulsion system on such a CO2 hopper is as follows: Carbon dioxide is extracted from the atmosphere and stored under pressure in a tank; when enough CO2 has been collected, the system is pressurized by heating the tank, and when a valve is opened the pressurized CO2 rocket is activated. Such a system can augment the operation of a surface rover to allow access to areas that are currently inaccessible, and do so in a package that is small and simple in operation. The ability to refuel anywhere on the surface combined with the simple operation of the rocket will allow the system to operate for an indefinite time and investigate many interesting locations. Efforts at Old Dominion University are currently underway to characterize the performance of a small-scale rocket that utilizes supercritical CO2 as a propellant for application on the surface of Mars. The apparatus used for testing consists of a pressure vessel to store the compressed CO2, a valve to initiate the flow of CO2, and a supersonic nozzle to accelerate the gas to Mach 2. The pressure vessel is filled with dry ice and sealed. The dry ice is then heated until the CO2 reaches supercritical conditions. At which point the valve is opened, discharging the CO2 to atmosphere, producing thrust. Stagnation pressure, temperature, thrust, and mass flow rate histories have been obtained so that specific impulse performance can be determined. Based on initial measurements it is possible to increase the time interval over which supersonic nozzle thrust levels can be sustained by heating the gas before it reaches the supersonic nozzle. Estimates of specific impulse are in the range of 100 to 130 seconds, while producing about 30 N of thrust through a nozzle with a 2 mm throat diameter. Since systems operating on Mars can take advantage of the low ambient pressure to produce supersonic flows with lower CO2 pressures, enhanced performance for Mars surface probes can be anticipated.

105

COMPUTATIONAL STUDY OF ROUGHNESS-INDUCED TRANSITION

Seokkwan Yoon, Michael D. Barnhardt and Emre Sozer

NASA Ames Research Center

ABSTRACT Laminar-to-turbulent transition in hypersonic flows may increase heat transfer rates significantly. Detached Eddy and Direct Numerical Simulations have been performed to study the complex physics of transition triggered by an isolated roughness in hypersonic flows. Both frozen and reacting flow solutions have been obtained for a hemisphere with a disk-like surface roughness element and compared to experiments done in a ballistic range at Mach 12 in air. The effects of high-enthalpy chemical reactions on roughness-induced transition will be investigated. Also, simulations will be performed for CO2 gas to study the effect of a Mars-like atmosphere on heating augmentation.

I. Introduction The design of hypersonic vehicles is challenging in several critical technology areas. The severe heating environment encountered during hypersonic flight dictates the shape of the vehicle. Boundary-layer transition at hypersonic speeds poses an especially significant challenge. Prediction and control of boundary layer transition in hypersonic flows are of crucial importance for the design of planetary entry vehicles as well as two-stage-to-orbit airbreathing access-to-space systems. Since turbulent heat transfer rates can be significantly higher than laminar heating rates, reductions in the weight of thermal protection systems can be realized with an improved understanding of the physics of transition from laminar to turbulent flow. The hypersonic heating environment, coupled with the emphasis on reusability, creates additional severe technology challenges for materials, material coatings, and structures that not only carry aerodynamic loads but also repeatedly sustain high thermal loads requiring long-life and durability while minimizing weight. The gap-filler incident of the Space Shuttle mission STS-114 in 2005 was a potent reminder of the importance of accurate prediction of roughness-induced boundary layer transition and subsequent increase in surface heating1. Direct Numerical Simulation (DNS) solves the Navier-Stokes equations by resolving a wide range of spatial and temporal scales of turbulence. Since DNS requires a number of grid points to resolve the Kolmogorov dissipative scales, it is not feasible for high Reynolds number flow simulations even with today’s most powerful supercomputers. Large Eddy Simulation (LES) requires less computational resources than DNS by modeling small eddies using sub-grid scale models while still resolving large eddies. However, even with this improvement, the grid requirements for high Reynolds number LES calculations are still impractical. Implicit Large Eddy Simulation (ILES) is a turbulent flow simulation method without a sub-grid scale model but not a fully resolved

106 DNS. Since the high cost of computation for LES comes from the near-wall region, hybrid models like Detached Eddy Simulation (DES) have been developed which alleviate the difficulty by using a Reynolds-Averaged Navier-Stokes (RANS) model in the boundary layer, while behaving like Smagorinsky’s LES model away from the wall. Among several applications, DES has been used to study high-speed reentry base flows with favorable results.2 A study has been performed to determine the feasibility of using computational fluid dynamics as a tool for predicting hypersonic boundary layer transition to turbulence and the resulting increase in heat transfer. Of particular interest is whether DES can be used to overcome the scaling problems associated with DNS and LES of boundary layers. Numerical simulations for a boundary layer trip oriented at 45 degrees to the flow inside a Mach 10 wind tunnel have indicated that DES can predict perfect-gas flow transition.3 Also, it has been shown that DES and ILES results are comparable when the grid is fine enough to resolve some of the small length scales.3 Recently high-enthalpy flow transition experiments have been conducted in the NASA Ames Ballistic Range for blunt bodies with isolated roughness elements.4The objective of the present paper is to simulate selected cases of the Ballistic Range experiments to validate the CFD code against the test data for a hemisphere with an isolated roughness element.

II. Preliminary Results The high-enthalpy experiments4 were performed in the Hypervelocity Free Flight Aerodynamic Facility, part of the Ballistic Range complex at NASA’s Ames Research Center. The Ballistic Range employs a two-stage light- gas gun to launch individual models on trajectories through a controlled-atmosphere test section. The largest gun has an inner diameter of 38.1 mm (1.5 in), and the test section is approximately 1 m across and 23 m long, measured from the first optical measurement station to the last. The models are in flight for an additional 10 m from the exit of the gun barrel to the first optical measurement station, during which time the launch sabot separates from the model and is trapped in the receiver tank. There are 16 optical measurement stations, spaced 1.524 m (5 ft) apart, along the length of the test section. Each station is equipped with orthogonal-viewing parallel-light shadowgraph cameras and high-speed timers for recording the flight trajectories. The hemispheres were made from commercially available titanium alloy ball bearings with a diameter of 2.86 cm, which were cut in half using an electrical discharge machining wire. The arithmetical average surface roughness was 0.2 µm, giving an aerodynamically smooth surface finish. Isolated, disk-like surface roughness elements were created by drilling holes perpendicular to the model surface at parametrically varied locations of 10o, 20o and 30o of arc length from the stagnation point, then press fitting cylindrical silicon carbide pins of diameter 762 µm into each hole, leaving exposed heights that were systematically varied to cover a wide range of Rekk values. Four such pins were located on each model, all at the same arc length from the stagnation point, and separated by 90o circumferentially. Roughness element heights were measured using greatly magnified silhouette images generated with an optical comparator. Figure 1 shows a shadowgraph picture of a model in hypersonic free flight.4

107 Our first test case is the flow in air over a high trip element (58 µm) located at 20o. The freestream pressure level is 0.175 atm, and the freestream temperature of the quiescent test gas is at 294.16K. Model/sabot packages are launched from a two-stage light gas gun at a nominal muzzle velocity of 4.22 km/s, yielding a freestream velocity of 4.08 km/s at a mid-range location. The corresponding nominal freestream Mach number is approximately 12. Figure 2 shows a thermal image of the shot. The wake appears to be shorter than the actual because of the camera angle. Results have been obtained on an unstructured grid that consists of approximately 40 million cells covering a quarter hemisphere. First, the effects of chemical reactions on hypersonic flows were investigated using the modified Steger-Warming flux vector splitting scheme. At hypersonic speeds, the perfect-gas assumption is no longer valid because molecular species dissociate due to the high temperatures resulting from aerodynamic heating. Vibrational and electronic excitation, dissociation, and ionization processes absorb energy, and hence result in lower temperatures than in a perfect gas. The decrease in temperature accompanies a rise in density, which in turn causes a thinner shock layer. A reacting flow solution in Fig. 3a, compared to a frozen flow in Fig. 3b, clearly shows that the bow shock is closer to the body and hence temperatures are lower in the shock layer. Since the bow shock is closer to the edge of the boundary layer, the transition is further affected by the production of an entropy layer.

III. References 1Yoon, S., Gnoffo, P.A., White, J.A., and Thomas, J.L.,“Computational Challenges in Hypersonic Flow Simulations,” AIAA Paper 2007-4265, June 2007. 2Barnhardt, M.D. and Candler, G.V., “Detached Eddy Simulation of the Reentry-F Flight Experiment,” AIAA Paper 2008-625, Jan. 2008. 3Yoon, S., Barnhardt, M.D. and Candler, G.V., “Simulations of High-Speed Flow over an Isolated Roughness,” AIAA Paper 2010-1573, Jan. 2010. 4Reda, D. C., Wilder, M. C., and Prabhu, D.K., “Transition Experiments on Blunt Bodies with Isolated Rough ness Elements in Hypersonic Free Flight,” AIAA Paper 2010-0367, 48th Aerospace Sciences Meeting, Jan. 2010.

• Research Scientist, NASA Advanced Supercomputing Division • † Research Scientist, ERC, Inc. • ‡ Research Scientist, ERC, Inc.

108

109

THREE DIMENSIONAL RADIATION IN MARTIAN ATMOSPHERE

Daniil Andrienko1,2, Sergey Surzhikov2

Moscow institute of Physics and Technology1, [email protected] Institute for Problems in Mechanics2, [email protected]

ABSTRACT

It has been long time understood that there is a strong necessity in accurate and time efficient method of radiative heat transfer prediction during planetary entry. These requirements are defined by strong radiative-gasdynamic interaction which takes place between hypersonic inlet flow and thermal protection capsule of descent space vehicle. An experience of past space probe missions and present-day calculations show that during capsule entering in Martian atmosphere it experiences heating by strong CO2 and CO bands. On the other hand, the radiative part of complex radiative-gasdynamic solvers can take significant (up to 90%) of total time. The order of temperature entering velocity is such to influence on gasdynamic parameters of inlet flow. The trajectory of probe is considered to be under non zero angle of attack, so the flow field is essentially three dimensional. This paper presents a high time efficient computational platform for three dimensional spectral radiation transfer calculation. The governing system of equations is solved by program code NERAT-3D. The radiative heat transfer is described by the P1-approximation of spherical harmonics method. The P1- approximation is an accurate and powerful method for radiation calculation in strong absorbing media. Considering the extremely strong absorption in infrared and UV part of spectrum in Martian atmosphere, the P1-approximation seems to be good enough to describe radiative heat transfer. The media is considered to be absorbing and emitting. The multigroup spectral model (100 spectral groups) is chosen to describe optical properties of Martian atmosphere. 10 species model (C, N, O, C2, N2, O2, CN, CO, NO, CO2) and 37 reactions are used to describe chemical properties of inlet flow. The computational platform performs with multiblock structured and unstructured grids. This fact allows calculating radiative heating of complex shape bodies. Two types of three dimensional capsules are used: spherical body with radius 66 cm and the body, similar to Pathfinder shape with radius 120 cm. Such blunt cone at front shield and

110 truncated cone at back shield shape is typical for capsule of new ESA mission EXOMARS. Radiative heating parameters are obtained for point of MSRO trajectory where thermal protection system experiences maximum of radiative heating. The parameters of inlet flow are =2.462x100 erg/cm3, = 1.01x10-8 g/cm3, = 129 K, =7.49x105 cm/s. These parameters correspond to the 42nd second of MSRO flight. The altitude is 50 km approximately. The spectral and integral heating calculation along the whole body surface is presented. The verification against tangent slab approximation and ray-tracing method is demonstrated. The volumetric radiative heat release in the whole computational domain is also obtained. The adequate accuracy of P1-approximation is demonstrated. The efficient strategy of P1-approximation method in case of small optical thickness is proposed. This step of optimization allows dramatically decrease the time consuming factor of P1-approximation comparing with ray-tracing method. The summarizing table of P1-approximation, tangent slab approximation and ray- tracing method time efficiency presented. In the conclusion, some general recommendations for efficient coupling radiative and gasdynamic solvers are suggested.

Session 6B - Aeroassist, Experimental Missions and EDL Mission Design

112

OVERVIEW OF THE NASA ENTRY, DESCENT AND

LANDING SYSTEMS ANALYSIS EXPLORATION FEED-FORWARD STUDY

Alicia D. Cianciolo1, Thomas A. Zang1, Ronald R. Sostaric2, M. Kathy Mcguire3

NASA Langley Research Center1 e-mail: [email protected],

[email protected], NASA Johnson Space Center2 e-mail: [email protected], NASA Ames Research Center3 e-mail:

[email protected]

ABSTRACT NASA senior management commissioned the Entry, Descent and Landing Systems Analysis (EDL-SA) Study in 2008 to identify and roadmap the Entry, Descent and Landing (EDL) technology investments that the agency needed to successfully land large payloads at Mars for both robotic and human-scale missions. Year 1 of the study focused on technologies required for Exploration-class missions to land payloads of 20 to 50 t. Inflatable decelerators, rigid aeroshell and supersonic retro-propulsion emerged as the top candidate technologies. In Year 2 of the study, low TRL technologies identified in Year 1, inflatable aeroshells and supersonic retropropulsion, were combined to create a demonstration precursor robotic mission. This part of the EDL-SA Year 2 effort, called Exploration Feed Forward (EFF), took much of the systems analysis simulation and component model development from Year 1 to the next level of detail. A main objective of the study was to determine the maximum payload mass (to Mars touchdown) capability of a Delta IV-H launch vehicle, given the spacecraft launch mass constraint of 7.2 t and assuming the 2024 Mars opportunity. The simulation results, using the latest component mass models, indicated that a direct entry system could deliver approximately 3.5 t to 0 km above the Mars Orbiter Laser Altimeter (MOLA) areoid. A second objective was to characterize the performance required of the supersonic retro-propulsion system. The study, which assumed four engines with a specific impulse of 338s and a system thrust-to-weight of 3.7 Mars g’s, yielded descent engine initiation between Mach 1.4 and 1.8 at an altitude between 3 and 8 km. A third major objective was to use the high fidelity entry simulation to characterize an Autonomous Landing and Hazard Avoidance Technology (ALHAT) like sensor suite for Mars. Initial performance range results were obtained for terrain relative navigation, hazard detection and avoidance, velocimeter and altimeter sensor systems. This paper summarizes the analysis performed to meet the EFF objectives, the study results and recommendations for future investment.

113

AEROFAST: MARTIAN AEROCAPTURE FOR FUTURE SPACE TRANSPORTATION – MISSION OVERVIEW

T. Salmon*1, F. Bonnefond1, J-M. Bouilly1, P. Augros2, T. Lutz3

EADS Astrium Space Transportation1, EADS Astrium Space Transportation2, EADS Astrium Space Transportation GmbH- Airbus-Allee3

ABSTRACT AEROFAST is a Mars aero-capture feasibility demonstration performed by twelve European companies leaded by AST-ST as prime, and funded under seventh framework programme of the European Commission. This study planned over 2.5 years will end in June 2011. An aero-capture is a flight manoeuvre that takes place at very high speeds within a planet’s atmosphere that provides a change in velocity using aerodynamic forces (in contrast to propulsive thrust) for orbit insertion. This aero-breaking technology becomes really attractive with respect to propulsion technology when the delta-V necessary for orbit insertion becomes greater than 1 km/s, which is the case for most of the future solar system exploration missions. Aero-capture is a very challenging system level technology where compromises have to be found between individual disciplines such as system analysis and integrated vehicle design, aerodynamics, aero-thermal environments, thermal protection systems (TPS), guidance, navigation and control (GNC), instrumentation... all these disciplines needing to be integrated and optimized as a whole to meet the mission specific requirements. Currently, Technology Readiness Level (TRL) of aero-capture technology in Europe is assessed at TRL2 to 3 whereas a TRL6 is mandatory to envisage the aero-capture technology for operational missions while mitigating development risks. The AEROFAST study fits with this goal, being dedicated to increase the TRL level of aero-capture technology up to TRL4 through a complete mission study of a Martian aerocapture. The objectives of AEROFAST project are: - OBJ1: Define a project of aero-capture demonstration. - OBJ2: Make a significant progress in space transportation by increasing the TRL of the

planetary relative navigation and the aerocapture algorithm up to 5. - OBJ3: Build a breadboard to test in real time the pre-aerocapture and aerocapture GNC

algorithms, - OBJ4: Demonstrate/prototype the thermal protection system for such a mission ---- - OBJ5: Define on-board instrumentation for aero-capture phase recovery.

114 The proposed paper will be dedicated to present an overview of the mission and to point out the improvements and results gained at the end of the study wrt challenging topics: - A description of the overall mission architecture will be proposed including the pre aero-capture phase (Earth to Mars transfer), aero-capture phase and post aero-capture phase (transfer to orbit). - The spacecraft design based on a composite architecture made of several modules will be depicted, with aero shape, aero thermal behaviour and budgets justified. - A specific emphasis will be put on the GNC concerns, algorithms validation implemented within a simulator for NRT and RT test being a key factor for success. - Meantime, in order to improve robustness wrt mass & centring concerns during critical aero-capture phase, results of innovative TPS improvements and testing will be presented.

115

MISSION ANALYSIS AND FLIGHT MECHANICS OF EARTH EXPERIMENTAL MISSIONS

Rodrigo Haya-Ramos1, Davide Bonetti1, Cristina Parigini1, Jorge Serna1, Gabriele de Zaiacomo1, Federico Massobrio2

DEIMOS Space S.L.U.1, Email: [email protected], Thales Alenia Space2, Email: [email protected]

ABSTRACT

In the European context, several experimental missions have been planned to improve the knowledge of hypersonic systems. The general aim is to increase the safety of the future re-entry or planetary probe missions and optimize designs by reducing margins. There are several needs for hypersonic experimentation. On one side, from the point of view of subsystems it is of interest to increase the TRL of critical EDL technologies, like TPS or GNC by demonstration with scaled vehicles in representative environments. On the other, there is also need to validate the tools used for design in several disciplines, in particular aerodynamics and aerothermodynamics, where scarce experimental data are available. Finally, the design of the experimental vehicle is also a demonstration of system design and operations. From the 3 levels of experimentation, in flight research, experimental demonstrators or full scale vehicle, the ones of interest in view of planetary probe missions are the first two. In flight research vehicles are test bed for basic research. The Expert and RadFlight capsules fall within this type, while IXV and BLAST belong to the area of experimental demonstration with subscale vehicles. The objective of the "RadFlight" Re-Entry Flight Experiment is to reduce the large margins considered today in the design of TPS for high speed science exploration sample return missions by improving our knowledge on radiation process, radiation / ablation coupling and occurrence of transition from laminar to turbulent boundary layer. It is a re- edition of the Fire II experiments. The RadFlight capsule is ballistic and falls within the 50 kg class. The Intermediate eXperimental Vehicle (IXV) is a re-entry demonstrator whose objective is to tackle the basic European needs for re-entry from LEO. The vehicle is a 2 Tons class lifting body with ceramic and ablative TPS materials performing a controlled re-entry. The Beyond LEO Advanced Subscale Test (BLAST) is a System Design Experience aimed to enhance the European system design capability and provide in-flight data.

116 BLAST is a high-speed demonstrator whose main objectives are the in-flight experimentation of TPS systems, the design and successful operation of a GNC system for a lifting re-entry configuration in skip entry and the collection of detailed information related to the flowfield during the superorbital entry phase. This paper focuses on the RadFlight, IXV and BLAST Mission Analysis and Flight Mechanics. The mission analysis and Flight Mechanics of an experimental mission plays a key role to assess the feasibility of the mission and to advance the expected benefits before entering in detailed definition phases. Experimental mission are usually very constrained by low cost and hence optimization, scaling and other simplifications cannot prevent the user from getting the intended knowledge. It is them important to understand the bounds and limitations and to analyze the compatibility between the required experimentation and the available resources. Advanced methods and tools are applied with the general aim of incorporating as much requirements as possible with increasing level of fidelity in order to reduce the design iteration loops. The Mission Analysis of such experimental vehicles has a double challenge: first, to identify a feasible design space where all of subsystems of the demonstrator can be designed. Ex: the vehicle must flight within an entry corridor with adequate stability and control characteristics. On the other, to ensure the representativeness of the flight envelope with respect to the research or demonstration. Ex: a stable flight within the entry corridor out of the region of interest for the intended experimentation is safe but useless. This paper presents the mission design for each of the 3 experimental vehicles with special emphasis on the coupling between the mission, system and the experimentation objectives. All the 3 missions are suborbital: the vehicle is injected in a suborbital arc and in the case of RadFlight and Blast there is a booster element that provides the additional energy needed to reach beyond LEO velocities. At the end of the suborbital arc the vehicle re-enters into the atmosphere. IXV and BLAST perform a controlled entry, while in the case of Radflight the entry is ballistic. As a result, in the 3 cases an end to end Mission design approach is required in order to properly couple the entry phase restrictions with the ascent capabilities. The main characteristic of the RadFlight Mission is the strong coupling between all of the phases: ascent, acceleration with the booster and re-entry. The feasibility requires an end to end approach from lift-off to parachute deployment in order to identify the entry corridor in which the capsule on one side respects all of the mission and System constraints, including compatibility with the Volna launcher and on the other fulfils the experimentation objectives in terms of aerothermodynamics environment to be measured (minimum level of radiation heat flux and coupling between convective and radiative flow). The IXV vehicle is the concept in more advance state (facing CDR actually). The large design margins required by the aerothermodynamics induce a narrow corridor which is challenging for both Mission and GNC. The Mission Design process considers visibility, Flying Qualities, safety and trajectory constraints during the mission design leading to a robust trajectory design. The end-to-end trajectory optimization process from lift-off to splashdown and performance evaluation (Monte Carl) is presented as one of the key

117 mission design elements, as well as the integration of high fidelity models for the vehicle aerothermodynamics and GNC. Finally, the BLAST Mission Analysis combines the complexity of the IXV vehicle in terms of representativeness of an operational mission with restrictive safety requirements with the peculiarities of the high speed re-entry, which tightly couples ascent and re-entry phases. Within the paper, the different Mission Design approaches will be presented and the main results and status discussed which includes trajectory optimization, Flying Qualities, GNC, end-to-end Monte Carlo assessments, visibility, safety aspects and technological aspects. All these activities have been carried out in the frame of industrial activities under ESA led projects. The author wants to acknowledge Lionel Marrafa, Salvatore Mancuso and Marco Caporicci (ESA) for their support and comments.

118

HAYABUSA REENTRY: TRAJECTORY ANALYSIS AND OBSERVATION MISSION DESIGN

Alan M. Cassell1, Gary A. Allen1, Jay H. Grinstead1, Manny E. Antimisiaris2, Jim Albers3, Petrus M. Jenniskens3

NASA Ames Research Center1, NASA Dryden Flight Research Center2,

SETI Institute3

ABSTRACT

On June 13th, 2010, the Hayabusa Sample Return Capsule (SRC) successfully re-entered Earth’s atmosphere over the Woomera Prohibited Area (WPA) in southern Australia in its quest to return fragments from the asteroid 1998 SF36 “Itokawa”. The SRC entered the atmosphere at a super-orbital velocity of 12.04 km/sec (inertial), making it the second fastest human-made object to traverse the atmosphere. The NASA DC-8 airborne observatory was utilized as an instrument platform to record the luminous portion of the SRC re- entry (~60 sec) with a variety of on-board instruments to capture the ultraviolet to near-IR wavelength regime. The predicted SRC entry state information at ~200 km altitude was propagated through the atmosphere to generate aerothermodynamic and trajectory data used in the initial observation flight path design and planning. The DC-8 flight path was designed by considering safety, optimal SRC viewing geometry and aircraft capabilities in concert with the predicted SRC trajectory. Subsequent entry state vector updates provided by the Deep Space Network (DSN) team at the Jet Propulsion Laboratory (JPL) were analyzed after the planned trajectory correction maneuvers (TCMs) to further refine the DC-8 observation flight path. Primary and alternate observation flight paths were generated during the mission planning phase which required coordination with Australian authorities for pre-mission approval. The final planned observation flight path was chosen based upon trade-offs between optimal viewing requirements, ground based observer locations (to facilitate post-flight trajectory reconstruction), predicted weather in the WPA and constraints imposed by flight path filing deadlines with the Australian authorities. To facilitate SRC tracking by the instrument operators, a series of two racetrack flight path patterns were performed prior to the observation leg so the instruments could be pointed towards the region in the star background where the SRC was expected to become visible. Initial post-flight trajectory reconstruction indicates the predicted trajectory was very close to the as-flown trajectory. An overview of the design methodologies and trade-offs used in the Hayabusa reentry observation campaign along with lessons learned will be presented.

Nomenclature

119 ARC = Ames Research Center DFRC = Dryden Flight Research Center JAXA = Japanese Aerospace Exploration Agency JPL = Jet Propulsion Laboratory DSN = Deep Space Network EFPA = Entry Flight Path Angle SRC = Sample Return Capsule TCM = Trajectory Correction Maneuver TPS = Thermal Protection System WPA = Woomera Prohibited Area EDL = Entry, Descent and Landing 1 Systems Engineer, ERC Inc., Entry Systems and Vehicle Development Branch, M/S 229-1, Associate Member. 2 Aerospace Engineer, ERC Inc., Aerothermodynamics Branch, M/S 230-4, 3 Project Manager, Aerothermodynamics Branch, M/S 230-4, Associate Fellow. 4 Navigator, Entry Systems and Technology Division, IPA University of California Santa Cruz, Fellow. 5 Systems Engineer, Aerothermodynamics Branch, M/S 230-3, Associate Member. 6 Principal Investigator, Associate Fellow.

120

A SIMPLE ANALYTICAL EQUATION TO ACCURATELY CALCULATE THE ATMOSPHERIC DRAG DURING

AEROBRAKING CAMPAIGNS VALIDATION IN THE MARTIAN CASE

F. Forget, M. Capderou

Laboratoire de Meteorologie Dynamique (LMD), Universite ([email protected]).

ABSTRACT Planetary orbital missions are often designed to fly through atmospheric layers dense enough to significantly alter the spacecraft velocity during a single orbit pass. On the one hand, such a maneuver can be used to circularize the orbit and lower the periapsis using much less fuel than what would have been necessary directly using a rocket engine (aerobraking). On the other hand, lowering the orbit periapsis of a scientific probe can be useful to perform in-situ observations in the lower thermosphere and mesosphere, increase the precision of the gravity field measurements, or improve the mapping of surface properties like the crustal magnetic field (e.g. the MAVEN mission to Mars, to be launched in 2013). To accurately compute the orbital perturbation due to the atmosphere, engineers must usually couple numerical simulators of the spacecraft navigation with atmospheric model of the density and the winds in order to integrate the action of the atmospheric friction timestep after timestep. We have developed such a tool by combining the state of the art satellite orbitography model Ixion with the LMD Mars General Circulation Model through the Mars Climate Database (see Millour et al., this issue) However, on the basis of theoretical considerations and thorough validation, we have discovered that the orbital perturbation due to the atmosphere can be calculated with very high accuracy using a simple analytical equation combining the orbit parameters and the atmospheric density and scale height at a single point: periapsis. The equation is derived from the complete equations of atmospheric motion around a planet and through the atmosphere, and take advantage of the fact that if the orbit is non circular (an eccentricity larger than 0.03 is sufficient), the time spent in the dragging atmosphere is short and the spacecraft velocity relatively constant while in the atmosphere. The main uncertainty lies in the assumed atmospheric winds. If the actual value of the velocity v0′ relative to the atmosphere at perihelion (i.e. v0′ computed with respect to planetary rotation and atmosphere) is known, the expression of the atmospheric drag over one period is simply:

121 ′2 !2π1√ ∆v=kρ0v0rp µ √e H (1) with ρ0 and v0 the atmospheric density and velocity at periapsis, rp the distance between periapsis and the center of the planet, µ = GM is the central attractive constant (for Mars µ = 4.282 837 1013 m3 s−2), e the eccentricity, and H the scale height of the atmosphere at periapsis. k = B/2 with B = CdS/m the so-called ballistic coefficient derived from the probe aerodynamical data and surface. In practice, the atmospheric circulation is not easy to estimate. It typically requires a general circulation model. An approximation is to the neglect the atmospheric winds and assumes that the satellite velocity relative to the atmosphere is the velocity in the galilean frame. This is especially valid for polar orbits.In that case, the atmospheric drag over one period simply becomes: " 1+e√ ∆v=kρ0 2πµ√e H (2) Such an equation can be useful to design future missions. For instance, future aerobraking or scientific "deep dip" campaign can be optimized by choosing the best combination of season, local time, latitude or longitude for the periapsis as well as orbit inclination, excentricity, etc. To our knowledge, such equations have not been described elsewhere. Analytical development can be found in King-Hele [1964]. In this fundamental book the author studied contraction of orbits under the influence of drag, in a spherically symmetrical atmosphere then in an oblate atmosphere. He gets very complex equations, always presented in analytical form, taking into account the variation of the orbital parameters orbit after orbit. Here we just compute ∆v for each individual orbit. We will present detailed validation studies performed by comparing ∆v calculations from a state of the art complete model with our simple equations in a wide variety of cases, and show that the results are always extremely accurate.

122

AEROBRAKING PERIAPSIS CONTROL STRATEGIES

M. Sánchez*, F. Cichocki

DEIMOS Space S.L.U [email protected]

ABSTRACT Space missions are becoming noticeably more complex and they are demanding to put increasingly bigger payloads in orbit. In this context, aerobraking emerges as an enabling technology to enhance the mass ratio about celestial bodies presenting atmosphere. Being already demonstrated operationally by NASA in various Mars missions, it has been identified as an essential phase for this type of missions to allow fulfilling tight requirements on available mass about Mars. But not only Mars missions can take benefit from this technology. Other scenarios such as Venus or Titan can consider the utilisation of this technique. Essentially, the purpose of the aerobraking technique is to reduce the energy of a highly elliptical orbit, transforming it into a low-eccentricity, low-altitude orbit by a sequence of atmospheric passes. In each pass, a small part of the S/C energy is dissipated into heat by aerodynamic friction mainly on the S/C solar panels, thus enabling progressive orbit apoapsis lowering and orbit period reduction. This technique allows achieving large ∆V savings with respect to the traditional chemical orbit insertion approach, which translates into an increase of the mass margin by reducing the propellant budget allocated to the insertion phase. On the other hand, insertion strategies based on aerobraking are longer and operationally more complex and demanding. Atmospheric passes are very demanding for the S/C structure in terms of thermal stress mainly, thus feasible aerobraking corridors must be defined so as to ensure the structural safety at a given confidence level. In particular, upper boundary of the corridor prevents from producing damages on the structures while lower boundary prevents from having too long durations. The most straightforward and efficient way to define the aerobraking corridor would be to use the actual control variable; i.e. the solar arrays temperature. The main problem with this approach is the difficulty to derive a representative thermal model to allow the derivation of realistic corridor boundaries. Moreover, the number and location of temperature sensor is a complex task. To tackle the issue described above, surrogate variables are used instead. Traditionally, constant aerobraking corridor based on dynamic pressure has been used. Although it

123 gives reasonable results in terms of safety margins and aerobraking duration, other options can be studies to improve the efficiency of the aerobraking. This paper describes alternative pericentre control strategies based on single and dual surrogate variables approach (dynamic pressure, heat flux or heat load). Control corridors based on single surrogate variables are defined by the evolution of a variable under consideration with time (from constant to more complex functions). Control corridors based on multiple surrogate variables define a domain where each pericentre pass represented as a combination of those variables must be confined, either naturally or through a dedicated manoeuvre at apocentre to counteract the natural evolution and bring it back to the boundaries of the control corridor. In order to discuss the different approaches described above, comparative performance assessments will be shown for those strategies when applied to different mission scenarios (Mars, Venus and Titan).

DEIMOS Space S.L.U.

124

PLANNED FLIGHT OF THE INFLATABLE REENTRY VEHICLE EXPERIMENT 3 (IRVE-3)

Robert A Dillman, F Neil Cheatwood, Stephen J Hughes, Joseph Del Corso, Richard J Bodkin, and Aaron Olds

ABSTRACT The Inflatable Reentry Vehicle Experiment 3 (IRVE-3) is planned for launch from NASA Wallops Flight Facility in the spring of 2012. IRVE-3 is a follow-on mission to the IRVE-II flight of 2009, which successfully demonstrated exo-atmospheric inflation, reentry survivability, and flight performance of a Hypersonic Inflatable Aerodynamic Decelerator (HIAD). IRVE-3 is intended to demonstrate the performance of a HIAD with a flight-relevant TPS exposed to a peak reentry heat rate above 15 W/cm2, and to demonstrate the effect of an offset center of gravity on HIAD flight performance. This paper discusses the IRVE-3 mission scenario, reentry vehicle design, expected flight environment, predicted vehicle response, and the various sensors that will allow quantification of the flight environment and vehicle performance. The design and expected performance of the inflatable aeroshell, inflation system, and CG offset mechanism will be discussed in detail, along with plans for future development flights and eventual mission use.

125

DIMENSIONLESS PARAMETERS FOR ESTIMATING MASS OF INFLATABLE AERODYNAMIC

DECELERATORS

Jamshid A. Samareh

NASA Langley Research Center, [email protected]

ABSTRACT This paper provides an overview of a mass estimating technique for inflatable aerodynamic decelerators. The technique uses dimensional analysis to identify a set of dimensionless parameters for inflation pressure, inflation gas mass, and flexible material mass. The dimensionless parameters are similar to drag coefficient, and these parameters allow scaling of an inflatable concept with geometry sizing parameters (e.g., diameter), environmental conditions (e.g., dynamic pressure), inflation gas properties (e.g., temperature), and mass growth allowance. This technique is suitable for estimating the mass of attached (tension cone, hypercone and stacked toroid) and trailing inflatable aerodynamic decelerators. The technique relies on simple engineering approaches developed by NASA in the 1960s, 1970s, and some recent developments. The technique was recently used for NASA’s Mars Entry and Descent Landing System Analysis (EDL-SA) project. The EDL-SA results were validated with two separate sets of finite element analyses. A typical inflatable concept consists of following components: toroid(s), radial straps, gores, rigid heatshield, and thermal protection system. The last two components will not be included in the final paper. The structural concept for toroids can be either film, coated fabric, or a combination of thin bladder covered with reinforced fabric material. The latter concept will also include axial straps to counter in- plane and out-plane buckling. The dimensionless parameter for minimum inflation pressure is a critical parameter for toroid mass, and this parameter is shown to be only dependent on the geometry of inflatable concept. The mass of the inflation system is based on a user-defined mass fraction. The gores are used as gas barrier layers that also carry loads produced by the dynamic pressure. The radial straps are used to connect the toroid(s) to the rigid heatshield and are made of high performance fabric. The results indicate that the dimensionless parameter for gas mass depends on only geometry parameters and gas properties. Similarly, the dimensionless parameter for mass of flexible material depends on only geometry and material properties.

Session 7A - Advances in TPS Technology for Planetary Probe Design

127

CHALLENGES WITH THERMAL PROTECTION MATERIAL DEVELOPMENT AND IMPLEMENTATION:

LESSONS LEARNED FROM RECENT NASA EXPERIENCE

D. Ellerby

NASA Ames Research Center, Moffett Field, CA

Unlike the structural materials used in spacecraft, (composites, Ti, Al, etc…), which have terrestrial applications, thermal protection materials are solely designed for and utilized in the reentry environment and typically do not have any other terrestrial uses. This results in TPS materials that have manufacturing processes, material constituents and material architectures with no or limited commercial applications, therefore, the burden for maintaining these supply chains rests completely upon a single end user, typically a government entity such as NASA. For large programs such as the space shuttle, with relatively frequent flight rates (flights per year), it is expensive but feasible to maintain these supply chains to support a single vehicle. But for one-off probe missions, with very low flight rates (years between flights), and depending upon the destinations wildly different reentry environments and thus TPS material requirements, Mars versus Jupiter entries for example, sustainability of material is a significant risk. Add to that the high costs associated with flight certifying new TPS materials, and the result is only a few materials ever being matured sufficiently for flight. These “flight proven” materials are applied to and/or proposed for a range of missions, even though they may not be optimized in terms of mass. Given the complexities of some of these TPS materials, the uncertainties associated with changing constituents on final material performance and the high cost associated with recertification, TPS material users fall into the trap of maintaining heritage traceability, which is often the proper approach, but has its own challenges/risks. Lastly when new materials are developed the material developers may focus on material performance in the reentry environment at the coupon level and lose sight of bigger issues with implementation of the material on the final flight vehicle. Keeping material integration in mind is important when spending scarce TPS material research and development funds. In this presentation we will review some of the challenges faced with developing, implementing and sustaining TPS materials at NASA.

128

ONGOING EUROPEAN DEVELOPMENTS ON ENTRY HEATSHIELDS AND TPS MATERIALS

H. Ritter1, O. Bayle1, Y. Mignon2, P. Portela3, J-M. Bouilly2, R. Sharda4

ESA/ESTEC1, EADS Astrium2, HPS Lda.3, Lockheed Martin Insys4

ABSTRACT The paper will provide an overview on some elements of current European developments for heatshields of atmospheric entry probes and TPS materials. In particular, it will talk about the ongoing heatshield development for the European ExoMars EDL demonstrator and the ongoing development of a European low-density ablative material for extreme heat flux applications. The joint ESA-NASA ExoMars program now includes two launches. While in 2018 a NASA spacecraft is planned to deliver a rover module to the surface of Mars, in 2016 a European composite spacecraft is planned to be launched consisting of an orbiter module and an EDL demonstrator (EDM). The EDM will have an entry mass of 600kg with a heatshield diameter of 2.4m. The heatshield will be based on a cork- based ablator. It will be designed in order to withstand not only the aerothermodynamic entry loads with peak heat fluxes up to 2 MW/m2, but also to survive the possibility of a severe dust storm during entry allowing an arrival during a global dust storm season. Further, a set of entry system sensors will be integrated in the heatshield allowing to reconstruct part of the entry environment and the TPS response. Various sample return missions have been studied in recent years. Most recently the ESA Cosmic Vision program has selected a revised version of MarcoPolo as one of four candidates for a medium-class mission that is planned to launch in the period 2020-22. MarcoPolo-R is a mission to return a sample of material from a primitive near-Earth asteroid (NEA) for detailed analysis in ground-based laboratories. The Earth return from extraterrestrial bodies involves a hyperbolic trajectory resulting in atmospheric entry velocities of typically around or above 12 km/s and resulting peak heat fluxes in the order of 10-20 MW/m2 with dynamic pressure loads up to around 1000 mbar. In addition, since the Earth return capsule is subject to a “double” delta-V (to the object and back to Earth), the return capsule and its heatshield have to conform to a very stringent mass budget. This requires the availability of a highly efficient light-weight ablator material. ESA has therefore initiated a dedicated activity aiming at the development of a European lightweight ablative material for extreme heat flux applications. Initial development has been completed following two different material concepts and first plasma tests showed promising results. Refined development is currently ongoing. However, further test results will not yet be available at the workshop.

129

MEDLI Aerothermal Environment Reconstruction Efforts

Todd White

ERC, Incorporated

ABSTRACT The Mars Science Laboratory (MSL), scheduled to launch in December 2011, is equipped with an heat shield instrumentation suite. The suite, named MEDLI for MSL Entry Descent and Landing Instrumentation, includes a series of pressure ports, thermocouples, and isotherm sensors embedded in the thermal protection material. Pressure ports and transducers are part of MEADS (Mars Entry Atmospheric Data System), while thermocouple and isotherm sensors make up the MISP (Mars Integrated Sensor Plug). This paper focuses primarily on the response of the MISP plugs (T1-T7) to Martian atmospheric entry (Figure 1).

Figure 1. MEDLI sensor locations (left); Sample heating pulse and MSL trajectory (right) The science goals of the MISP ports are to verify turbulence transition, stagnation region heating, catalytic augmentation, subsurface material response, and surface recession of the ablative heat shield. However, MISP can only directly measure discrete in-depth temperatures, thus the remaining science objectives must be addressed through data-analysis and aerothermal environment reconstruction using computational fluid dynamics (CFD) and material response codes. This paper will describe the MISP science objectives and the current state of reconstruction efforts. These efforts focus on coupled CFD and material response models to simulate anticipated effects of transition and catalycity on MISP sensors, and include sensitivity studies on MSL design trajectories. Additionally, this paper will discuss arc-jet and material properties tests planned in support of the MISP reconstruction.

130

ORION FLIGHT TEST-1 THERMAL PROTECTION SYSTEM INSTRUMENTATION

T. John Kowal1

NASA Johnson Space Center1 Email: [email protected]

ABSTRACT The Orion Crew Exploration Vehicle (CEV) was originally under development to provide crew transport to the International Space Station after the retirement of the Space Shuttle, and to provide a means for the eventual return of astronauts to the Moon. With the current changes in the future direction of the United States’ human exploration programs, the focus of the Orion project has shifted to the project’s first orbital flight test, designated Orion Flight Test 1 (OFT-1). The OFT-1 is currently planned for launch in July 2013 and will demonstrate the Orion vehicle’s capability for performing missions in low Earth orbit (LEO), as well as extensibility beyond LEO for select, critical areas. Among the key flight test objectives are those related to validation of the re-entry aerodynamic and aerothermal environments, and the performance of the thermal protection system (TPS) when exposed to these environments. A specific flight test trajectory has been selected to provide a high energy entry beyond that which would be experienced during a typical low Earth orbit return, given the constraints imposed by the possible launch vehicles. This trajectory resulted from a trade study that considered the relative benefit of conflicting objectives from multiple subsystems, and sought to provide the maximum integrated benefit to the re-entry state-of-the-art. In particular, the trajectory was designed to provide: a significant, measureable radiative heat flux to the windward surface; data on boundary transition from laminar to turbulent flow; and data on catalytic heating overshoot on non-ablating TPS. In order to obtain the necessary flight test data during OFT-1, the vehicle will need to have an adequate quantity of instrumentation. A collection of instrumentation is being developed for integration in the OFT-1 TPS. In part, this instrumentation builds upon the work performed for the Mars Science Laboratory Entry, Descent and Landing Instrument (MEDLI) suite to instrument the OFT-1 ablative heat shield. The MEDLI integrated sensor plugs and pressure sensors will be adapted for compatibility with the Orion TPS design. The sensor plugs will provide in-depth temperature data to support aerothermal and TPS model correlation, and the pressure sensors will provide a flush air data system for validation of the entry and descent aerodynamic environments. In addition, a radiometer design will be matured to measure the radiative component of the reentry heating at two locations on the heat shield. For the back shell, surface thermocouple and pressure port designs will be developed and applied which build upon the heritage of the Space Shuttle Program for instrumentation of reusable surface insulation (RSI) tiles.

131 The quantity and location of the sensors has been determined to balance the needs of the reentry disciplines with the demands of the hardware development, manufacturing and integration. Measurements which provided low relative value and presented significant engineering development effort were, unfortunately, eliminated. The final TPS instrumentation has been optimized to target priority test objectives. The data obtained will serve to provide a better understanding of reentry environments for the Orion capsule design, reduce margins, and potentially reduce TPS mass or provide TPS extensibility for alternative missions.

132

FLEXIBLE ABLATORS: APPLICATIONS AND ARCJET TESTING

James O. Arnold1, Ethiraj Venkatapathy1, Robin Beck1, Kathy M. McGuire1, Dinesh K. Prabhu2 and Sergey Gorbunov3

ABSTRACT The concept of flexible ablators was developed as a “technology pull” to meet the need for a thermal protection system (TPS) that could enable large (23 meter diameter) hypersonic inflatable aerodynamic decelerators (HIADs). The initial application was incorporated in system analysis studies [1] which showed that large HIADs can be employed to mass-efficiently place payloads of order 40 metric tons (mT) on the surface of Mars with arrival masses of ~ 80 mT. Follow-on systems analysis studies [2] of potential robotic precursor missions to Mars showed that flexible ablators could also be used on smaller HIADs to efficiently place payloads on Mars in excess of 2.5 mT with an arrival mass of 7.2 mT. This use of flexible ablators enables one approach to enable robotic Mars missions with payloads exceeding that of the Mars Science Laboratory at ~ one mT, the capped mass using Viking-era technology. Recent system studies [3] of deployable heat shields using mechanical erection methods in the transformable entry system technology (TEST) show that flexible ablators are enabling for human Mars missions and for mission to Venus, both involving aerocapture and subsequent out-of orbit entry. Flexible un-deployed supersonic inflatable aerodynamic decelerators (SIADs) from hypersonic heating of Mars landers [4]. Flexible ablators can simplify the design and manufacture and reduce cost of TPS for conventional, rigid-body vehicles [5]. This follows since flexible ablators are conformal, eliminating thermal structure design issues. They are manufacture-able from 1.8 meter wide felts so the number of gap/seams in a TPS are greatly reduced as compared to conventional tiled systems such as in the rigid phenolic impregnated carbon ablator (PICA) design for Orion. This presentation will include a brief summary of how flexible ablators are made and have been tested that is covered in more detail elsewhere [6]. This presentations will focus on the range of entry vehicle heating environments where flexible ablators may be applicable, and will also discuss new arcjet testing approaches under consideration for flexible ablators for both conformal and deployable applications. 1NASA Ames Research Center, Moffett Field, CA 94035 2ERC Inc. 3Jacobs Technology, Inc. References [1] Dwyer-‐Ciancolo, Alicia, et. al, “Entry, Descent and

133 Landing System Analysis Study: Phase I Report”, NASA/TM-‐2010-‐216720, July 2010. [2] Dwyer-‐Ciancolo, Alicia, et. Al, Entry, Descent and Landing System Analysis Study: Phase II Report, Exploration Feed-‐Forward”, NASA/TM-‐2011-‐217055, February 2011. [3] Venkataphathy, E., et. al, “Transformable Entry Systems Technology”, “ , 21st AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar 23-‐26 May 2011, Trinity College, Dublin, Ireland. [4] James O. Arnold, et. al, “Thermal Protection System for Supersonic Inflatable Aerodynamic Decelerator Cover Protective Shield”, 21st AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar 23-‐26 May 2011, Trinity College, Dublin, Ireland. [5] James. O. Arnold “Affordable Thermal Protection Systems for Future Entry Vehicles: Lessons Learned from Shuttle, Orion and Ongoing Research and Development”, Commercial and Government Responsive Access to Space Technology Exchange. October 25-‐28, 2010. Moffet Field, CA. [6] Robin A. Beck, et. al, “Overview of Initial Development of Flexible Ablators for Hypersonic Inflatable Aerodynamic Decelerators” 21st AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar 23-‐26 May 2011, Trinity College, Dublin, Ireland.

134

Overview of Initial Development of Flexible Ablators for Mars EDL

Robin A.S. Beck, Susan White, James Arnold, Wenhong Fan, Mairead Stackpoole, Parul Agrawal

NASA Ames Research Center, Moffett Field, CA, USA

ABSTRACT

The Vision for the EDL Technology Development Project (EDL TDP) is to develop world class Entry, Descent, and Landing (EDL) technologies for Exploration Class Missions. The objective of the EDL Exploration Class Missions Project is the development of applicable technologies to a readiness level of TRL (Technology Readiness Level) six for specific Exploration Class Missions. The NASA Exploration roadmap calls for human exploration of Mars beginning in the decade of the 2030s, with precursor missions to the Low Earth Orbit and the Moon in preceding decades. While the technologies for LEO and Lunar return to Earth are reasonably mature and are under further development within NASA, the necessary technologies for landing astronauts and exploration class payloads (> 40 metric ton) on the surface of Mars do not exist today. The only proven EDL architecture for Mars entry is based on Viking heritage, with extensions for Mars Science Laboratory (MSL). However, this architecture is fundamentally limited to landed masses of about 2 metric tons, and cannot meet landed elevation and landing precision requirements for larger class exploration missions. The Design Reference Mission as defined by the Entry Descent Landing Systems Analysis for Mars Missions Requiring Large Surface Payloads document (EDLSA-001) is to deliver multiple 40 metric ton payloads to the surface of Mars in order to support human exploration, in-situ resource utilization, and large scale exploration. Previous technology roadmaps have demonstrated that the current TRL of the necessary EDL components is so low that immediate technology development is required to support this timeline. Even if the need date of the technologies were to slip, low to mid TRL technology development is still a high priority, because of the long lead times of the required elements. In both the hypersonic and supersonic stages of EDL there are only two proposed technology candidates, and at the current level of fidelity it is not known whether either will be scalable to exploration class missions. During its first year, the EDL TDP was divided into three elements: Thermal Protection Systems (TPS); Aeroshell Modeling and Tool Development (MAT); and Supersonic Retro- Propulsion (SRP). Now, in its second year, the EDL TDP has further divided the TPS element into two separate elements: Rigid TPS (R-TPS); and Flexible TPS (F-TPS) This paper will describe the steps being taken in pursuit of advanced ablative flexible TPS materials and systems with performance which support Exploration Class Systems. The Flexible TPS element will focus on developing material concepts for a 23-m

135 deployable entry system to survive dual pulse heating (peak ~120W/cm2). Because the peak heat flux exceeds 50 W/cm2, ablative materials will be required for the TPS. The Flexible TPS Element will define, develop, and model the ablative thermal material protection system concepts required to allow for the human exploration of Mars via aerocapture followed by planetary entry. For the initial project year, the focus of the deployable materials tasks was on evaluating the wide breath of possible TPS concepts selecting an initial set of concepts for thermal and structural screening and evaluation. NASA scientists developed flexible versions of known rigid ablative materials by replacing the rigid reinforcements with flexible equivalent materials and exploring with various resin compounds for impregnation into the reinforcements. In addition, organic flexible materials were also impregnated with resins and included. Evaluation criteria were developed for relevant materials comparisons and ranking. Existing materials properties were used to develop low fidelity models used to determine the design the proper screening test facilities and conditions therein and specimen geometries. Folding tests, radiation transparency tests and thermal evaluation tests in a radiant environment and an aerothermal environment were performed on each of the screening materials. Results of the screening tests were ranked Aerocapture-to-orbit and Entry according to the evaluation criteria and the first round of down-selections for further development were made. This paper will present the results and overview of the initial development and evaluation of a new class of materials: ablative flexible materials.

136

AEROFAST: DEVELOPMENT OF CORK TPS MATERIAL AND A 3D COMPARATIVE THERMAL/ABLATION

ANALYSIS OF AN APOLLO & A BICONIC SLED SHAPE FOR AN AEROCAPTURE MISSION

G. Pinaud1 & A.J. van Eekelen2

ASTRIUM-SAS1, Email: [email protected], SAMTECH2, E-mail: [email protected]

ABSTRACT An Aerocapture vehicle travelling from Earth to Mars approaches that planet on a hyperbolic interplanetary trajectory. Upon arrival, the vehicle will perform a single atmospheric pass to significantly reduce its speed, and enters into an orbit around the planet. This manoeuvre uses aerodynamic drag instead of propulsion for orbit insertion, and potentially leads to large mass (fuel) savings as well as reduced flight times (higher arrival speed). However, Aerocapture results in significant aerodynamic heating, necessitating a Thermal Protection System (TPS), as well as the use of a guidance system to assure that the spacecraft leaves the planetary atmosphere on the correct trajectory. In the frame of the seventh European Community Framework Program (FP7), the AEROFAST (AEROcapture for Future space tranSporTation) research and development project aims at preparing a demonstration of a Martian Aerocapture mission and increasing the Technology Readiness Level (TRL). One of the aims of this paper, is to present the development of an innovative cork based material and the selection process of the different formulations. The material must be able to withstand the severe front shield aerothermal environment. Numerous formulations have been investigated using a parametric combination of cork granule size, resin type/ratio, reinforcement fraction, fillers and the mixing and agglomeration processes. A basic (thermo-mechanical) characterization and qualitative analysis allowed for a first selection of the 4 most promising candidates. These candidates are being tested in the inductive plasma wind-tunnel facilities (COMETE) of ASTRIUM. These tests are performed in a stagnation point configuration, for an aerothermal environment similar to the AEROFAST aerocapture mission. In parallel, a 3D ablation and charring material model has been implemented in the finite element program SAMCEF, and successfully validated during the AEROFAST project. The numerical model consists of three sets of equations, namely the transient heat balance equation, the steady state mass balance equation and the charring equations. For

137 the charring of the material we use a multi-species Arrhenius model with the species densities as degrees of freedom. The ablation is modelled by a surface imposed and temperature dependent ablation speed, followed by an in volume mesh deformation. Two main probe aerodynamic shapes and concepts have been evaluated, namely an Apollo like shape and a biconic sled with a characteristic diameter of 4 m. A thermo-mechanical comparative analysis of the front-shield has been carried out. The space probes are made of Norcoat-Liège (a low density phenolic resin impregnated cork material) which will serve as a baseline solution. The 3D heat load history (convective and radiative), over the front-shield, is based on the maximum energy trajectory extracted from a statistical Monte Carlo GNC study for a CO2 Martian atmosphere. Due to a non uniform heat load distribution on the heat shield (non axisymmetric shape plus a 30° flight trim angle), an optimization of the TPS thickness has been performed in order to save mass. Finally, the biconic sled vehicle has been selected for its several advantages (internal volume, ease of TPS manufacturing) and its innovative features (possible adaptation to other missions). On the basis of these preliminary experiments, additional efforts will be devoted to the modelling of the thermal, swelling and ablative behaviour of the selected cork based material (developed within this project).

REFERENCES [1] H. Requiston, F. Bonnefond, Ph. Augros, J.-M. Bouilly, S. Reynaud and U. Westerholt. AEROFAST: AEROcapture for future spAce tranSporTation. In 1st EU-ESA International Conference on Human Space Exploration, number IAC-09-A3.I.3, 2009. [2] H. Requiston, P. Augros, F. Bonnefond, J.-M. Bouilly, T. Lutz, H. Scheer, AEROFAST: AEROCAPTURE FOR FUTURE SPACE TRANSPORTATION, 7th International Planetary Probe Workshop,14-18 June 2010. [3] A.J.van Eekelen, G. Pinaud, J.-M. Bouilly, AEROFAST: Thermal/Ablation analysis of the front heatshield for a Martian aerocapture mission – 7th International Planetary Probe Workshop, June 14-18, 2010.

138

MODULAR MANUFACTURING OF HONEYCOMB-REINFORCED CHARRING ABLATOR SYSTEMS FOR

THE AEROSHELLS OF LARGE EDL VEHICLES

William M. Congdon

ARA Ablatives Laboratory (ABL), Centennial, CO USA

ABSTRACT Polymer-based charring ablator heatshields are made more robust by the use of honeycomb (HC) reinforcement. For large EDL vehicles of 3.5-m and greater (and especially for the massive HMMES vehicles planned for manned exploration of Mars), production by direct HC packing on vehicle aeroshells poses numerous and significant challenges. These challenges are greatly reduced by the use of modular manufacturing where pre-packed and precision-milled ablator units are secondarily bonding to vehicle structures. Producibility is enhanced, costs are lowered, and common manufacturing risks are eliminated. This paper discusses modular manufacturing techniques for EDL heatshields developed at the Ablatives Laboratory over the past four years of technical effort. Multiple modular units were produced and evaluated and the real benefits of this manufacturing approach will be discussed.

Session 7B - Airless Body Surface Missions

140

ROBOTIC AND HUMAN SPACE EXPLORATION OF

NEAR-EARTH OBJECTS

D.D. Mazanek

NASA-Langley Research Center, Space Mission Analysis Branch, Hampton, VA USA Email: [email protected]

This presentation will provide an overview of human mission planning considerations for the exploration of Near-Earth Objects (NEOs). Topics will include a brief discussion of the characteristics of these small, airless planetary bodies, the robotic precursor information required to permit future human missions, the main requirements that will drive human mission operations, and the reasons for sending humans.

141

EUROPEAN GNC TECHNOLOGY DEVELOPMENT AND PERSPECTIVE FOR AIRLESS BODIES EXPLORATION

A. Caramagno

DEIMOS Space S.L.U, Ronda de Poniente 19, Tres Cantos (28760 - Madrid) - Spain,

email: [email protected] Exploration of solar system planets and minor bodies is an outstanding goal and challenge within ESA Science and Exploration Programmes. This lecture provides a European industrial perspective of the Guidance, Navigation and Control (GNC) developments focusing on this mission category. The available capabilities and maturity status are presented, as well as a gap analysis aimed to identify the future needs and strategies to reduce cost and time for Mission implementation. GNC technologies for mission to airless bodies can be clustered under a set of similar requirements and drivers. The absence of a relevant atmospheric density represents the major difference with respect to mission to planets whose mass and origin has allowed the preservation of an atmosphere for which an Entry, Descent and Landing system is adopted. Although there is a natural link and synergy among the two classes of missions, the presence or absence of atmosphere dictate different approaches to the design of a GNC able to autonomously and safely reduce the arrival or orbital velocity down to values compatible with landing system dynamics. Furthermore, within the airless body missions, a further distinction between mission to major and minor bodies drives the strategy for mission design. As a consequence, the chain of nominal and off-nominal GNC modes depends on the expected operations: e.g. approach for rendezvous and orbiting (with or without the addition of a landing phase), or intentional impact (e.g. NEO deflection missions), and an eventually body departure for sample return missions. From the GNC perspective, a crucial problem is achieving a design implementing heterogeneous mission phases through a robust combination of on-board functions and an optimized equipment set, whilst containing the uncertainties associated to development and verification effort. With focus on airless body missions, the lecture will presents European achievements, mature and under development GNC technology solutions, as well as the expected future application in line with current ESA Programmes. Starting from GNC requirements, a solution represents a balanced level of complexity, on-board resources demand, development risk and cost. As a result, the GNC subsystems is specified and designed including the selection and specification of the sensor suite and some dedicated actuators. The perspective covers a time-span of the past decade and a look ahead to missions under study or development. The analysis puts in evidence a remarkable level of innovation as well as the mature capabilities of the European industry, worth being considered by European decision makers and international partners.

142

Magic (Mobile Autonomous Generalized Instrument Carrier) T. van Zoest1, T.-M. Ho1, C. Lange1, L. Witte1, S. Wagenbach1, C.

Krause1, S. Ulamec1,

J. Biele1, Florian Herrmann1, Joachim Block1, and Pierre Bousquet2

1 DLR – Deutsches Zentrum f. Luft- und Raumfahrt, Germany

2 CNES – Centre National d'Études Spatiales, Toulouse, France

ABSTRACT In this talk, a medium size mobile robotic surface platform (MAGIC) with a weight of 10 kg for in-situ exploration on small bodies like Asteroids and Comets will be presented. The concept of MAGIC is based on extensive feasibility studies as well as breadboarding activities of a dedicated lander, called MASCOT, studied for the flight opportunity onboard Hayabusa- 2, a NEO sample return mission of JAXA/JSPECS (Japan Aerospace Exploration Agency/JAXA Space Exploration Center) to the Asteroid 1999JU3. As a next step, MAGIC will be a standardized lander platform for different mission scenarios and varying payload components. The lander platform design will have functionality such as mobility and autonomy particularly needed to explore the uncertain surface of a small body, such as an Asteroid. The rationale for the mobility function is to access the diversity of several surface sites. Although the lander is commandable from mission control, it requires the high degree of autonomy to execute its mission in an efficient way while acting flexible and responsive in face of the uncertain environment of the asteroid’s surface. The platform is designed to be deployed from a supporting main space craft, orbiting or hovering above the target body. Once deployed on the surface it can upright and relocate by hopping and carry its scientific payload to different sampling sites. All surface operations, including hopping, science measurements and data transmission, are conducted fully autonomously. MAGIC will provide a well balanced combination of system and functional capability, lifetime and Mission flexibility, based on nanosat technology to be integrated and qualified for demanding deep space exploration. In summary, MAGIC shall be able to deliver a wide range possible scientific instrumentation (potentially up to a limit of 3kg total mass), to study the body’s physical properties (mass, density, temperature), internal structure, surface and subsurface structure (microscopic to macroscopic scale) and its chemical composition, thus being a complement to any rendezvous or sample return missions to small bodies.

143

THE ESA LUNAR LANDER MISSION

A.Pradier*, B.Gardini, C. Philippe, B. Houdou, R. Fisackerly, J.D. Carpenter, D. De Rosa

ESA-ESTEC, The Netherlands, *E-mail: [email protected]

ESA’s Lunar Lander mission shall be launched in 2018 and shall address the primary objective of demonstrating a key capability for exploration, soft precision landing with hazard avoidance. Embarking a suite of navigation sensors, and advanced navigation and hazard avoidance algorithms integrated together within the overall GNC, the Lunar Lander mission will represent a major step in technology demonstration even before beginning operations on the surface. However once on the Moon the Lander shall focus on its second objective, the deployment and operation of a surface payload dedicated to the investigation of the lunar surface specifically in preparation for future robotic and human exploration. This mission represents a major element of ESA’s preparation to participate in the future of exploration as part of a broader international cooperation. While the precise framework of this broader cooperative effort is under consolidation, the Lunar Lander represents a clearly defined and, to a certain extent self-contained, mission opportunity. Having passed through Phase A iterations which have established a strong foundation of understanding of the key issues, the project is currently progressing through Phase B1. This key project phase shall focus on important choices to be made at mission and technology level, in order to put the following design steps on a secure basis to realise a successful landing and surface mission at the end of this decade. An important aspect of this is the identification of driving characteristics of the lunar surface environment within the candidate landing zones, in terms of surface topography, local slopes, hazardous boulders etc. This work is being carried out using the most up-to-date and accurate data ever collected on the lunar South Pole, in the form of altimetry and imagery datasets from NASA’s Lunar Reconnaissance Orbiter. In support of the mission and system definition activities of Phase B1, a dedicated stream of technology breadboarding work shall provide key inputs in terms of verification of critical performances and of important assumptions. Thus adding confidence to the overall mission design maturity. The Lunar Lander project shall arrive at the end of Phase B1 activities, in mid 2012, with a clear understanding of the implications of the lunar environment, mission constraints and technological challenges, and with a mission baseline with which to continue design work up to preliminary design review (PDR). The ESA Lunar Lander project represents an opportunity in which the fields of autonomy and robotics must be considered in an overall mission context, with technological and programmatic constraints and the particular challenges of the lunar south pole, an important environment for future exploration.

144

CAMERA-AIDED INERTIAL NAVIGATION FOR PINPOINT PLANETARY LANDING ON RUGGED

TERRAINS

Jeff Delaune1, Guy Le Besnerais1, Martial Sanfourche3, Aurélien Plyer4, Jean-Loup Farges5, Clément Bourdarias6, Thomas Voirin7 and Alain

Piquereau8

(1), (5), (8) ONERA, 2 avenue Édouard Belin, 31000 Toulouse (France), Email:

[email protected], [email protected] (2), (3), (4) ONERA, Chemin de la Hunière et des Joncherettes, 91120 Palaiseau

(France), Email: [email protected], [email protected], [email protected] (6)Astrium ST, 66 Route de Verneuil, 78130 Les Mureaux

(France), Email: [email protected] (7) ESA-ESTEC, Postbus 299, 2200 AG Noordwijk (The Netherlands), Email: [email protected]

ABSTRACT

This paper tackles the pinpoint navigation challenge for autonomous planetary landers using a single camera and an Inertial Measurement Unit (IMU) to reach a 100-meter position error requirement at touchdown. Inertial navigation schemes embedded in previous missions suffer from position and attitude (pose) error growth due to integration of IMU acceleration and angular rate measurement errors. When looking down at a terrain in sunlight conditions, image data provided by the camera allow for identifying surface features from reference maps. These absolute landmarks prevent the error growth. At the same time, image features are tracked at a higher rate through the image sequence to make the absolute landmark matching step more robust. Feature tracking will eventually limit error drift at low altitude when the map resolution is too poor to be useful. IMU data allow high-bandwidth, low-delay state estimation in any environment condition and is capable of solving the scale problem associated to camera measurements. Inertial and optical data fusion is implemented through extended Kalman filtering which tightly integrates image feature points measurements to IMU-based state propagation. Tight integration of visual and inertial measurements within the navigation filter allows for working in degraded conditions with a few features only and thus is more robust than vision-only solutions. Image measurements provided by a camera are bidimensional by definition. Many times in literature, a planar terrain is assumed to avoid computationally-costful methods dealing with highly-3D areas. Though, such terrains will be encountered in future planetary exploration missions, for instance to the mountainous lunar south pole. Two aspects that are the most challenging over rugged terrains are matching absolute landmarks with the on-board map and estimating depth of relative features to predict their images coordinates in the correcting part of the filter.

145 The first contribution of this work is the definition of an absolute landmark matching process that uses landmark constellations, designed as an extension of Landstel [1], to work over any type of terrain, from flat to hilly. Another contribution is its integration within a full-state navigation filter able to cope with computer requirements associated with space exploration missions. We compute a landmark constellation at each point of interest selected in an orbital image from an image intensity criterion. 3D coordinates of the point and its neighbors are extracted from a digital elevation model in order to account for terrain relief. The constellation itself is based on the angular and distance distributions of its neighbors, stored in a signature vector. Not only are the constellation signatures robust to terrain topography and illumination changes, but they can adapt to all surface features while maintaining low-memory requirements. Once a landmark is matched with the map, its true 3D coordinates are used to build an absolute image measurement in the filter. Relative image measurements coming from other features are relying on the estimation of the 3D position of image features to predict their image coordinates and subsequently update the filter. We compare the performance for two state-of-the-art tight fusion schemes: Simultaneous Localization And Mapping filters (SLAM), and Sliding- Windows filters (SW). The SLAM approach estimates the spacecraft pose parameters along with 3D positions of the image feature points in the state vector of the filter. Visual measurements can be processed without delay but SLAM has a high computational cost associated to a large state vector. For many descent trajectories, points are only crossing the camera field of view for a limited period of time. We thus process SLAM feature points in a limited temporal window from the last image backwards and discard them when they leave the field of view to decrease the computational cost. The other approach is a SW filter [2]. Unlike SLAM, it only estimates spacecraft pose parameters, but keeps previous camera poses in the state vector for a limited and sliding temporal window in order to process measurements. These measurements are built by triangulating 3D positions of a point from the first and last image of the sequence where it appears. Measurements are thus delayed and 3D reconstruction of points is less accurate than SLAM, but the computational cost is lower. SLAM-based and SW-based version of our navigation system were implemented in an orbit-to-touchdown lunar descent and landing simulator coupled with an image generator. Results are presented, discussed and compared with a special focus on the performance over mountainous areas and filtering issues. Future indoor and experimental validation test benches are presented. [1] Pham, B. V.; Lacroix, S.; Devy, M.; Drieux, M. & Philippe, C., Visual Landmark Constellation matching for spacecraft pinpoint landing, AIAA Guidance, Navigation and Control, 2009 [2]Mourikis, A. I.; Trawny, N.; Roumeliotis, S. I.; Johnson, A. E.; Ansar, A. & Matthies, L., Vision-Aided Inertial Navigation for Spacecraft Entry, Descent, and Landing, IEEE Transactions on Robotics, 2009, 25, 264-280

146

MARCO POLO-R: AN ASTEROID SAMPLE RETURN MISSION

Mark Adler, Andy Cheng, Tom Randolph, and Rob Maddock

ABSTRACT MarcoPolo-R is an asteroid sample return mission, which has been selected by ESA for Assessment Phase study as a medium-class mission following the 2010 Cosmic Vision announcement. MarcoPolo-R is proposed as an ESA collaboration with NASA. It will rendezvous with a primitive Near-Earth Asteroid (NEA), scientifically characterize it at multiple scales, and return a unique sample to Earth unaltered by the atmospheric entry process or terrestrial weathering. The proposed baseline mission scenario of MarcoPolo-R to the primary target NEA 1996 FG3 is as follows: a single primary spacecraft provided by ESA, carrying the Earth Re-entry Capsule, sample acquisition and transfer system provided by NASA, will be launched by a Soyuz-Fregat rocket from Kourou into GTO using two space segment stages. Launch windows are identified in the 2020-2024 time frame. The re-entry system will be an application of the NASA developed, chuteless design for Mars Sample Return.

147

WHAT MOONRISE LUNAR SAMPLE RETURN CAN TEACH US ABOUT MARS SAMPLE RETURN

George Chen1, Eric Blood2

(1) Jet Propulsion Laboratory, 4800 Oak Grove Drive, Pasadena, CA, 91109, USA Email: [email protected] (2) Jet Propulsion Laboratory, 4800 Oak Grove

Drive, Pasadena, CA, 91109, USA Email: [email protected]

ABSTRACT Under consideration by NASA are two challenging robotic sample return missions: the MoonRise Lunar Sample Return, a New Frontiers proposal which could launch in 2016; and the Mars Sample Return (MSR) campaign, which could begin with the proposed Mars 2018 Sample Caching mission. Together they could represent the first wave of sample return missions envisioned from various solar system bodies. While the MoonRise and Mars 2018 mission concepts are vastly different in their details, it is a worthy exercise to take several steps back to recognize their common attributes and technology needs, especially in the disciplines of entry, descent, and landing (EDL), sample acquisition, and surface operations. Even though both proposed projects are still early in their development cycles, a number of common technologies and systems engineering disciplines are emerging as enhancing, and, in some cases, enabling for both missions. Additionally, lessons learned from the MoonRise Lunar Sample Return could also provide valuable insight for future legs of the Mars Sample Return campaign beyond the proposed Mars 2018 mission. Specifically, MoonRise experiences with Ascent, Earth Entry, Descent, and Landing (EDL), and sample recovery operations would provide useful guidance for the analogous phases of MSR. MoonRise, if selected, would return samples to Earth in 2017, the timing of which would allow the feed forward of flight system development knowledge and mission operations experience to the development of the proposed Mars Sample Return campaign. As flight systems engineers for both the proposed MoonRise and the Mars 2018 sample caching missions, the authors have a unique vantage point to the challenges of both sample return missions and will discuss how MoonRise would build the experience base for a Mars sample return. Furthermore, lessons learned from the recently completed MoonRise Mission Concept Study and the on-going Mars 2018 concept development study suggest focus areas for technology investments, which could benefit future sample return missions under study at this time.

148

FARSIDE EXPLORER: UNIQUE SCIENCE FROM A MISSION TO THE FARSIDE OF THE MOON

David Mimoun1, Mark Wieczorek2, and the Farside Explorer Team3

Universite de Toulouse, ISAE/SUPAERO1, [email protected], Instiut de Physique du Globe de Paris2, Team members: full list available at http://fars

ide.spacecampus-paris.eu/

ABSTRACT

Farside Explorer is a proposed Cosmic Vision medium-sized mission to the farside of the Moon consisting of two landers and instrumented relay satellite. The farside of the Moon is a unique scientific platform in that it is shielded from terrestrial radio-frequency interference, it recorded the primary differentiation and evolution of the Moon, and it lacks Earthshine and can be continuously monitored from the Earth-Moon L2 Langrange point. The primary scientific objectives of the Farside Explorer mission are to make the first radio-astronomy measurements from the most radio-quiet region of near-Earth space, to determine the internal structure and thermal evolution of the Moon, from crust to core, and to quantify impact hazards in near-Earth space by the measurement of impact flashes. The Farside Explorer flight system includes two identical solar-powered landers and a science/telecom relay satellite to be placed in a halo orbit about the Earth-Moon L2 Lagrange point. One lander would explore the largest and oldest recognized impact basin in the solar sysem- the South Pole-Aitken basin – and the other would investigate the primordial highlands crust. Radio astronomy and geophysical instruments would be deployed on the surface, and the relay satellite would continuously monitor the surface for impact events.

149

As a direct consequence of the scientific requirements, the proposed space segment includes two spacecraft to land on the farside of the Moon, and instrumented relay satellite, and the launcher (either a Soyuz or Ariane 5 shared commercial launch). The proposed mission concept is innovative by using a halo orbit about he Earth-Moon L2 Lagrange point (LL2) to provide a relay to the farside landers while simultaneously enabling the impact flash monitoring program. The proposed ballistic trajectory starts from GTO to go to the Earth L1 and uses the instability of the manifold next to EL1 to return to the vicinity of the moon. At lunar arrival, the intersection of the manifolds of the Earth-Sun and Earth-Moon system allows for the insertion, at a very low ∧V, of the mission elements into an Earth-Moon LL2 halo orbit.

150 The flight system uses the heritage of the Moonnext study. It is composed of two identical solar-powered landers and a science/telecom relay satellite in an LL2 halo orbit. It uses an ATV like thruster used for transfer and landing braking; bi-propellant hydrazine system provides attitude control, final descent, and landing. Continuous science operations are allowed by either RHUs or specific thermal design, based on parabolic reflectors. It has a dry mass of about 380 kg a wet mass: 1185 kg. The payload mass is 26 kg when the payload power is 200 W (day), 4 W (night). The LL2 relay satellite is based on a small-satellite bus with a wet mass of about 150 kg, including a 50 kg payload (mainly the telecom relay). Mission was foreseen to last four years, with an early launch around 2019-2020, to overlap other geophysical missions, and therefore provide a geophysical network. The Farside Explorer is supported by the radio astronomy and lunar science communities. This consortium represents 7 international Lunar Science Institutes and about 500 individuals.

151

VLBI TRACKING OF PHOBOS-GRUNT PROBE

Guifré Molera Calvés1,2, S.V. Pogrebenko3, G. Cimò3, D.A. Duev3,4, L.I. Gurvits3,5

(1) Aalto University Metsähovi Radio Observatory, Metsähovintie 114 Kylmälä, FIN-02540, Finland, E-mail: [email protected] (2) University of California, Berkeley,

Centre for Astronomy Signal Processing and Electronics Research (3) Joint Institute for VLBI in Europe, Dwingeloo, The Netherlands, E-mail: [email protected],

[email protected], [email protected], [email protected] (4) Moscow State University, Faculty of Physics, Moscow, Russia (5) Delft University of Technology, DEOS - Faculty of

Aerospace Engineering, Delft, The Netherlands

ABSTRACT The Phobos Sample Return mission, also known as Phobos-Grunt, will be launched by the Russian Federal Space Agency in November 2011 and is expected to arrive to the Martian system in May 2012. The primary focus of the robotic lander is to collect a sample of the soil from the Phobos surface and return it to Earth for laboratory analysis. After the departure of the return vehicle from Phobos, the landing module will remain operational on the Phobos surface for at least a year. Being equipped with an X-band transmitter locked to the ultra-stable oscillator, it will be used as a beacon for the Planetary Radio Interferometry and Doppler Experiment (PRIDE), which will address several key scientific objectives of the mission. In particular, PRIDE-Phobos will enable characterisation of the gravitational field and geodetic parameters of the Martian moon. The European VLBI Network (EVN) radio telescopes can offer ground support for this experiment, in collaboration with the Centre for Deep Space Communication located in Ukraine. During the last two years, as a preparatory stage for PRIDE-Phobos, several operational planetary spacecraft have been observed with the radio telescopes in Metsähovi (FI), Yebes (ES), Wettzell (DE), Onsala (SE), Matera, Medicina, Noto (IT), and Pushchino (RU). Our team has successfully conducted the Doppler and VLBI spacecraft tracking experiments with a number of deep space missions, such as the ESA’s Huygens Titan Probe [1], the Smart-1 Lunar probe [2], ESA Venus Express (VEX) and Mars Express (MEX) during the Phobos-flyby [3]. During the recent years, the PRIDE group has been developing a series of scientific software tools for measurements of the Doppler-shift of the spacecraft carrier signal and accurate estimates of the spacecraft state vectors using the VLBI phase referencing technique. Observing PRIDE sessions with the VEX spacecraft were used as a test bench to optimize the technique and reduce the lag of data processing from weeks down to several hours. Rapid results are crucial for the upcoming deep space missions in view of their potential applicability for mission operations. The accuracy of the state vectors estimates depends on several parameters, of which the most

152 important ones are the stability of the on-board oscillator and the power of the carrier signal. The SNR level of the Doppler and VLBI fringe depend on these parameters. Based on the recent experiments with the VEX and MEX spacecraft, we expect to achieve the accuracy of better than a few cm/s for the radial velocity and better than 50 m for the lateral position in the case of the Phobos-Grunt. In this paper, we report the latest results of PRIDE observations of the VEX and MEX orbiters with the EVN radio telescopes. In these experiments we achieve a milli-Hz level of radio signal spectral resolution accuracy and extract the phase of the spacecraft carrier signal with the accuracy better than 1 radian. As a scientifically attractive by-product of these observations we present characterisation of the interplanetary plasma along the signal propagation line on various spatial and temporal scales at different solar elongation angles. These carrier signal phase fluctuations are well represented by a near-Kolmogorov spectrum. Results obtained from PRIDE observations of the VEX spacecraft so far will be used as a benchmark for the future PRIDE-Phobos observations.

REFERENCES [1] Bird, M.K., Gurvits, L.I., Pogrebenko, S.V. et al., “The vertical profile of winds on Titan”, Nature, Vol. 438,8 December 2005 [2] Pogrebenko, S.V. et al., “First results of the First EVN VLBI Practice Run on the Smart‐1”, Presentation at Cassini, PSG meeting, 21‐23 June 2006, Nantes, France. [3] Molera Calvés G. et al., Venus Express spacecraft observations with EVN radio telescopes, 7th International Planetary Probe Workshop, 12-18 June 2010, Barcelona.

Session 8 - Closing

154

AUSTERITY IN THE AGE OF INNOVATION

Bethany Johns

ABSTRACT Federal budget cuts, deficit spending, raising the debt ceiling, mandatory spending, fiscal responsibility, balancing the budget – these phrases are guiding much of the debate on Capitol Hill. How does this rhetoric affect the funding for the sciences? The Administration believes that funding the sciences and education is the way to, “Out innovation, out educate, and out build,” the rest of the world and has proposed a federal budget that supports this initiative. However, the debate in Congress is about how the federal deficit impacts our global economic competitiveness and how cuts in spending are necessary for a stable government. This debate has led to a late enactment of the fiscal year 2011 federal budget. Its many continuing resolutions cause confusion on how the federal budget process usually works. In fact, there is a time line with which the process should follow. There are points along the time line when you can make an impact on the policy making process. The Decadal Surveys produced by the astrophysics, planetary science and heliophysics communities in the United States, impact policy by the community coming to a consensus and prioritizing the science it wants to accomplish within the decade. I will speak on the current events on the fiscal year 2011 budget, the 2012 federal budget, the current climate for science funding, and the impact you can make on the policy making process for science and planetary exploration. I will also talk about current progress on the funding for restart of production of plutonium-238, the fuel for planetary spacecraft.

NASA-LANGLEY RESEARCH CENTER’S ENGINEERING DIRECTORATE

Stephen P. Sandford

NASA-Langley Research Center, 5 N. Dryden St., Hampton, VA 23681 Email: [email protected]

The personnel in the Engineering Directorate at NASA-Langley Research Center are involved in a significant number of spaceflight projects, fulfilling roles spanning from concept development to flight hardware manufacturing and mission operations. This presentation will provide an in-depth look at Langley’s contributions to planetary probes, atmospheric science instruments, human spaceflight vehicles, and entry vehicle technology development. With Langley’s rich heritage as an atmospheric flight center, many of the Directorate’s roles on these modern-day missions relate to aerodynamics, aerothermodynamics, trajectory simulation, instrumentation, and flight dynamics and control. These strengths, along with expertise in structures and several unique, world-class facilities, has postured NASA Langley to be a key team member on today’s most exciting space missions.

156

POSTERS

Poster Session 2 – Probe Missions

Poster Session 2 158

STUDY OF PLANETARY ENTRY PROBES (PEP) FOR VENUS AND OUTER PLANETS: SATURN, URANUS AND

NEPTUNE

Denis Rebuffat, Peter Falkner, Jonan Larranaga, Jens Romstedt, Kelly Geelen

European Space Agency- ESTEC, Advanced Studies and Technology Preparation Division, Directorate of Science and Robotic Exploration, European Space Agency – ESTEC, Keplerlaan 1, 2201AZ Noordwijk, The Netherlands [email protected]

ABSTRACT The aim of the Planetary Entry Probe (PEP) study in ESA’s Concurrent Design Facility (CDF) was to examine entry and descent conditions for Venus, Saturn, Uranus and Neptune. Further, commonalities and dissimilarities and their relation to technological challenges between the different entry scenarios were assessed. While the Venusian atmosphere is CO2 rich with traces of N, the other three planets maintain a H and He dominated atmosphere like Jupiter. In a previous ESA CDF study, a Jupiter entry probe was designed. The respective entry conditions with respect to velocity and correlated heat load on the probe is considered as a worst case condition. Thus these design specifications were used as the starting point for the current activity and updated according the environmental specifics of the new target bodies. As a nominal scenario, the probe is released in a hyperbolic trajectory by a carrier which is subsequently used for data relay. The probe performs an entry followed by a descent phase performing scientific measurements down to the 100 bar pressure line. The parachute design and release strategy was adapted to the scientific relevance of the various atmospheric layers. Sometimes this led to free-fall phases where no parachute is used in order to balance the time spent into different layers according to their science interest, while meeting the time constraint imposed by the data relay. In the case of Venus, a piggyback launch of the probe on a mission performing a gravity assist manoeuvre at Venus was addressed as well . All entry probes showed a good degree of similarity, with a mass ranging from 254 to 326 Kg depending on the target planet. A mass of 10 kg was reserved for a generic suite of scientific payload for atmospheric research. The main differences concerning the Entry

Poster Session 2 159 and Descent System (EDS) are the Thermal Protection System (TPS) thickness (due to different heat loads) and the parachute release strategy. For each planet, a description of the probe is provided as well as the mission profile, both being justified by design, mission and environment analyses performed by the study team experts.

Poster Session 3 – Science from Probes and Penetrators


ACCOMMODATION STUDY FOR AN ANEMOMETER ON A MARTIAN LANDER

Benjamin Lenoir1, Don Banfield1

(1)Cornell Astronomy, 420 Space Sciences, Cornell University, Ithaca, NY 14853, USA, Email: [email protected]

ABSTRACT

Measuring the winds near the surface of Mars as well as their turbulent fluctuations is important to more fully understand the behavior of the boundary layer of Mars.This in turn is important to minimize the risk in landing for future exploration at Mars, but also to understand the interaction between the surface and atmosphere in terms of the transfer of heat, momentum and trace constituents including dust, water and other trace gases. Instrumentation is now becoming available for Mars that can measure not only the mean winds, but also their turbulent fluctuations and also resolving the full 3-D nature of the wind rather than just the horizontal winds (e.g., see Banfield’s abstract regarding a Martian Sonic Anemometer). With the instruments becoming available, the question is raised of how best to place such an instrument on a Martian lander or rover to yield the most undisturbed flow measurements in the presence of the lander/rover, and in the case where flow distortions can not be avoided, how to correct for these perturbations. To address this question, we used computation fluid dynamics to model the boundary layer flow at Mars, as well as the mean and turbulent flow distortions that would be realized at various positions around simplified lander/ rover structures. We first tuned our model to match the rough conditions experienced by Mars Pathfinder in terms of the range of roughness lengths and friction velocities seen, although under the assumption of neutral stability. Armed with this, we inserted into the flow a hemispheric lander with radius 1m and a half cube that just fit inside the hemisphere.We investigated the nature and correctability of the flow distortions that resulted from the flow around these simplified lander/rovers at various positions around them. We found that the exact shape of the lander/rover was not very important for ranges greater than 1.2m from the center of the sphere or cube. Presumably these results may then be extrapolated to more complex lander/rovers of similar sizes. We found that the mean flow and the turbulent characteristics of the flow (as expressed in terms of the 6 Reynolds Stresses) were least perturbed when the anemometer was placed at least 1.8m from the center of the spherical lander/rover. Additionally, if the instrument were canted 55 degrees above horizontal the flow distortions were again minimized when considering all possible azimuths for wind direction. Finally, our modeling suggests that the mean and turbulent characteristics of the perturbed flow are correctable to a high degree to yield the equivalent unperturbed

Poster Session 3 162 flow that would have resulted without the lander/rover present at all when the anemometer placement meets or exceeds this range from the lander/rover center and is placed at 55 degrees elevation. While this study used idealized lander/rovers and neutral stability conditions, we believe it is instructive in a general sense for the placement of anemometers on rovers. It is encouraging that good results were found to be possible with an instrument located only 0.8m from the edge of the lander/rover, simplifying anemometer accommodation on a realistic martian lander.


THE MARS CLIMATE DATABASE, CURRENT STATUS AND FUTURE IMPROVEMENTS

E. Millour(1), F. Forget(1), A. Spiga(1), S. Lebonnois(1), S.R. Lewis(2), L. Montabone(3), P.L. Read(3),M.A. López-Valverde(4), F. González-Galindo(4), F. Lefèvre(5), F. Montmessin(5), M.-C. Desjean(6), J.-P.

Huot(7) and the MCD/GCM development team

(1)Laboratoire de Météorologie Dynamique, IPSL, Université Pierre et Marie Curie, BP99, 4 Place Jussieu, 75005, Paris, France, Email: [email protected] (2)Department of Physics and Astronomy, The Open University, Milton Keynes, UK

(3)Atmospheric, Oceanic & Planetary Physics, University of Oxford, UK (4)Instituto de Astrofísica de Andalucía, Granada, Spain (5)Laboratoire Atmosphères, Milieux,

Observations Spatiales, IPSL, Paris, France (6)Centre National d'Etudes Spatiales, Toulouse, France (7)European Space Research and Technology Centre, European Space

Agency, Noordwijk, Netherlands

ABSTRACT What is the Mars Climate Database? The Mars Climate Database (MCD) is a database of meteorological fields derived from General Circulation Model (GCM) numerical simulations of the Martian atmosphere and validated using available observational data. The MCD includes complementary post-processing schemes such as high spatial resolution interpolation of environmental data and means of reconstructing the variability thereof. The GCM is developed at Laboratoire de Météorologie Dynamique du CNRS (Paris, France) [1,2] in collaboration with the Open University (UK), the Oxford University (UK) and the Instituto de Astrofisica de Andalucia (Spain) with support from the European Space Agency (ESA) and the Centre National d'Etudes Spatiales (CNES). The MCD is freely distributed and intended to be useful and used in the framework of engineering applications as well as in the context of scientific studies which require accurate knowledge of the state of the Martian atmosphere. Since its release in May 2008, Mars Climate Database v4.3 has been distributed to over 130 teams around the world. Current applications include entry descent and landing (EDL) studies for future missions (ExoMars, MSL), investigations of some specific Martian issues (via coupling of the MCD with homemade codes), analysis of observations (Earth-based as well as with various instruments onboard Mars Express and Mars Reconnaissance Orbiter),... The MCD may be accessed either online (in a somewhat simplified form) via an interactive server available at http://www-mars.lmd.jussieu.fr (useful for moderate needs), or from the full DVD-ROM version which includes advanced access and post-

Poster Session 3 164 processing software (contact [email protected] and/or [email protected] to obtain a free copy). Overview of MCD contents The MCD provides mean values and statistics of the main meteorological variables (atmospheric temperature, density, pressure and winds) as well as atmospheric composition (including dust and water vapor and ice content), as the GCM from which the datasets are obtained includes both chemistry [3] and full water cycle [4] models. The database extends up to ~350km, i.e. up to and including the thermosphere[5,6]. Since the influence of Extreme Ultra Violet (EUV) input from the sun is significant in the latter, 3 EUV scenarios (solar minimum, average and maximum inputs) account for the impact of the various states of the solar cycle. In order to account for and adequately represent the variability of the Martian atmosphere due to atmospheric dust distribution, the MCD includes 4 different dust scenarios which describe extreme cases (from very clear skies to global planet-wide dust storms) and a baseline scenario “MY24” for which the dust loading of the atmosphere is that obtained from assimilation of TES observations [7] in 1999-2001 (i.e. during Mars Year 24, following the calendar proposed by R.T. Clancy [8], which starts on April 11, 1955, at Martian solar longitude Ls=0°). The following values are provided in the MCD: • Atmospheric density, pressure, temperature and winds (horizontal and vertical), • Surface pressure and temperature, • CO2 ice cover, • Atmospheric turbulent kinetic energy, • Thermal and solar radiative fluxes, • Dust column opacity and mass mixing ratio, • [H2O] vapor and [H2O] ice columns and mixing ratios • [CO], [O], [O2], [N2], [CO2], [H2] and [O3] volume mixing ratios, • Air specific heat capacity, viscosity and molecular gas constant R. Validation of the MCD Climatology The MCD has been validated using available data, from TES, onboard MGS, for surface and atmospheric temperature, but also from atmospheric temperature retrieved from radio occultation using the ultra-stable oscillator onboard MGS. The assessment of the correctness of the surface pressure predictions was obtained using Viking Lander 2 measurements. The MCD includes a validation document which reports all the comparisons between MCD outputs and available datasets of measurements.

Poster Session 3 165 Left: Distributions of binned temperature differences (using bins of 1K) between MCD predictions (using different dust scenarios) and TES measurements for latitudes ranging from 50°S to 50°N. Displayed MEAN and RMS values are computed from the obtained histograms and the curves correspond to normal distributions of same MEAN and RMS. Right: Surface pressure cycle over a Martian year, as predicted by the baseline MY24 scenario at Viking Lander 2 site, with an envelope of twice its standard deviation, compared to recorded values. Towards the next version of the MCD We are currently working on a building a new and improved Mars Climate Database (version 5). One essential step towards this achievement is running an improved version of the GCM which will include all recent improvements and developments [9]:

o An improved CO2 cycle resulting from the inclusion of realistic subsurface water ice tables in the Polar Regions [10].

o Improved radiative transfer with updated radiative properties of dust, along with the implementation of the radiative effect of water ice clouds [11].

o An improved water cycle [9,11]. o An updated chemistry package [12]. o An improved representation of the non LTE (Local Thermodynamical

Equilibrium) phenomena in the thermosphere [13]. o We plan to update the thermal inertia and albedo maps used by the GCM. o We will take into account the recently derived map of surface roughness

values [14] (instead of using a fixed value of 1cm everywhere, as we have so far).

o We are also currently working on implementing the “thermal plume model” [15], a significant improvement to the current convective adjustment scheme in the GCM.

In addition to these technical improvements of the LMD GCM itself, we will include in Mars Climate Database version 5 more dust scenarios, which will include all Mars Years from MY24 to MY29 (as derived by [16]). Again some “extreme” (cold, warm, global dust storm) scenarios will also be provided to bracket reality as best as possible. We also plan to improve the MCD software with the addition of a subgridscale variability near the surface, where the nature and amplitude of this added variability would be derived from simulations using the LMD Mars Mesoscale Model [17].

References

[1] Forget F. et al. (1999) JGR, 104, E10. [2] Lewis S. R. et al. (1999) JGR, 104, E10. [3] Lefèvre F. et al. (2004) JGR, 109, CiteID E07004. [4] Montmessin F. et al. (2004) JGR, 109, E10, CiteID E10004. [5] Angelats I Coll et al. (2005) Geophys. Res. Lett., 32, 4, CiteID L04201. [6] Gonzalez-Galindo F. et al. (2005) JGR., 110, E9, CiteID E09008. [7] Montabone L. et al. (2006) 2nd Int. Workshop on Mars Atmosphere Modeling and Observations. [8] Clancy R.T. et al. (2000) JGR, 105, E4. [9] Forget et al (2011) 4th Int. Workshop on Mars Atmosphere Modeling and Observations. [10] Millour E. et al. (2009) 3rd Int. Workshop on Mars Polar Energy Balance. [11] Madeleine J.-B. et al. (2011) 4th Int. Workshop on Mars Atmosphere Modeling and Observations. [12] Lefèvre F. et al. (2011) 4th Int. Workshop on Mars Atmosphere Modeling and Observations. [13] Lopez-

Poster Session 3 166 Valverde M. A. et al. (2011) 4th Int. Workshop on Mars Atmosphere Modeling and Observations. [14] Vistowski C. et al (2011) 4th Int. Workshop on Mars Atmosphere Modeling and Observations. [15] Rio C. and Hourdin F. (2008) J. Atmos. Sci., 65, 407-425. [16] Montabone et al. (2011) 4th Int. Workshop on Mars Atmosphere Modeling and Observations. [17] Spiga A. and Forget F. (2009) JGR, 114, CiteID E02009


ARMADILLO – A DEMONSTRATION FOR LOW-COST IN-SITU INVESTIGATIONS OF THE UPPER

ATMOSPHERE OF PLANETARY BODIES

Rene Laufer(1,3), Glenn Lightsey(2), Georg Herdrich(3,1), Ralf Srama(3,4,1), Gregory Earle(5), Carsten Wiedemann(6), Ed Chester(7),

Hugh Hill(8), Troy Henderson(9), Rainer Sandau(10,11,1), Lorin Matthews(1), Truell Hyde(1)

(1)Center for Astrophysics, Space Physics and Engineering Research (CASPER), Baylor

University, Waco, Texas, USA, E-Mail: [email protected], (2)Department of Aerospace Engineering and Engineering Mechanics, University of Texas at Austin,

Texas, USA, (3)Institute of Space Systems, Universitaet Stuttgart, Germany, (4)Max-Planck-Institute for Nuclear Physics, Heidelberg, Germany, (5)Department of Physics,

University of Texas at Dallas, Texas, USA, (6)Institute of Aerospace Systems, Technische Universitaet Braunschweig, Germany, (7)AEVO GmbH, Gilching, Germany,

(8)International Space University (ISU), Strasbourg, France, (9)Aerospace and Ocean Engineering Department, Virginia Polytechnic Institute and State University (Virginia

Tech), Blacksburg, Virginia, USA (10)German Aerospace Center (DLR), Berlin, Germany, (11)International Academy of Astronautics (IAA), Paris, France

ABSTRACT

ARMADILLO (Attitude Related Maneuvers And Debris Instrument in Low Orbit) is a low Earth orbit small satellite mission under development by the Satellite Design Lab (SDL) of the University of Texas at Austin in collaboration with the Center for Astrophysics, Space Physics and Engineering Research (CASPER) of Baylor University and the Institute of Space Systems of the University of Stuttgart. The project was recently selected to participate in the University Nanosatellite Program UNP-7 to be designed and built in the 2011-2013 timeframe with the goal to target a 2014 launch opportunity. The 3-unit cubesat will demonstrate the combination of precise attitude control for nanosatellites, a cold-gas micro- propulsion system and a miniaturized dust/debris detector. The attitude control system consists of GNC computer, IMU, GPS receiver, sun sensors, magnetometer, reactions wheel, magnetorquers and low-cost optical navigation star tracker with the goal of achieving 0.1 degree 3-axis attitude control. The cold gas propulsion system is based on an Aerospace Corporation design and will provide approximately 50 m/s impulsive capacity and a delta-v resolution of around 0.1 m/s – during the ARMADILLO mission used for the end-of-mission de-orbit from low Earth orbit. The Piezo Dust Detector (PDD) is a miniaturized in-situ measurement instrument of around 0.5 kg to detect dust and debris particles of up to 1 mm size. The detector is a

Poster Session 3 168 joint development of CASPER (Baylor University) and the Institute of Space Systems (University of Stuttgart) in partnership with the Cosmic Dust Group at the Max-Planck- Institute for Nuclear Physics, Heidelberg based on the experience from preparation and tests of the Mercury Dust Monitor for the European BepiColombo mission. ARMADILLO will demonstrate the capabilities necessary for a mission to perform in-situ investigations of the upper atmosphere, e.g. of the Earth. At least two – preferred is a constellation of more than two – ARMADILLO-like spacecraft would travel piggyback with a carrier probe, separating at some point in orbit. Using its own chemical or electrical micro-propulsion system for de-orbit, the nanosatellites would lower their altitude performing in-situ plasma and dust measurements before being destroyed. The paper will present the ARMADILLO satellite and possible instrument design (e.g. the PDD and plasma instrumentation from partners such as UT Dallas and Institute of Space Systems, Stuttgart), Also the required adjustments for planetary upper atmosphere investigation missions (e.g. in the AOCS subsystem) will be addressed as well as the scientific results expected from that missions.

Poster Session 4 – EDL Technology Development


ONGOING VALIDATION OF COMPUTATIONAL FLUID DYNAMICS FOR SUPERSONIC RETRO-PROPULSION

Daniel G. Schauerhamer,* Kerry A. Trumble†, William Kleb†, Jan-Renee Carlson§, Pieter G. Buning,** Karl Edquist††, and Emre

Sozer‡‡

Daniel G. Schauerhamer*, Jacobs Technology, Houston, Texas, 77058 Kerry A. Trumble, NASA Ames Research Center, Moffett Field, California, 94035

William Kleb‡, Jan-Renee Carlson§, Pieter G. Buning** and Karl Edquist†† NASA Langley Research Center, Hampton, Virginia, 23681

Emre Sozer‡‡ ERC, Moffett Field, California, 94085

ABSTRACT

Supersonic Retro-Propulsion (SRP) is a viable means for deceleration of high mass vehicles entering into the Martian atmosphere1-6. Previous methods of deceleration are not scalable for exploration type vehicles which can potentially weigh tens of metric tons. Since ground and flight testing of SRP at entry conditions can be difficult and cost- prohibitive, the development of this enabling technology can be enhanced with the ability to predict the flow field numerically using Computational Fluid Dynamics (CFD). SRP results in a complex flow structure involving shocks, shear layers, recirculation and stagnation regions, which makes validation of the CFD methods a high priority. The validation process includes using multiple CFD codes to compare to historic and recent wind tunnel tests7. Through code-to-code and code-to-test comparisons, best practices in gridding, numerical method selection, and solution advancement are established, and validity is added to the CFD methods. With validation, CFD can be confidently applied to the actual entry problem in all its complexity8. Four CFD codes are being applied to SRP: DPLR9, FUN3D10, 11, OVERFLOW12, and US3D13. The codes all solve the Reynolds Averaged Navier-Stokes equations, but differ in implementation, grid type, and numerical methods. The focus of this paper will be on the comparison of the CFD codes to a recent wind tunnel test which was designed primarily for CFD validation. The experiment was conducted by the NASA Exploration Technology Development Program in the Langley supersonic 4’x4’ Unitary Plan Wind Tunnel in June, 201014, 15. The cases that will be presented all have a free stream Mach and Reynolds (per foot) number of 4.6 and 1.5E+06, respectively, but vary by the number due to the inherent unsteadiness of the flow fields. Qualitative comparisons of the flow structure will be made by comparing CFD to high-

Poster Session 4 171 speed test Schlieren, and quantitative comparisons will be made by comparing averaged surface pressure with pressure tap data from the tunnel. Unsteady shedding frequencies of the CFD solutions are also compared to high-frequency pressure gauges from the test. This paper will first introduce SRP, the CFD codes, and the wind tunnel test. Then it will present code-to-code and code-to-test comparisons, discuss the results including modeling strengths and weaknesses, and offer conclusions of the study. * Aerospace Engineer, Applied Aerosciences and CFD Branch, MS EG-3, [email protected]. † Research Scientist, Aerothermodynamics Branch, MS 230-2, [email protected]. ‡ Aerospace Engineer, Aerothermodynamics Branch, MS 408A, [email protected]. § Aerospace Engineer, Computational Aerosciences Branch, MS 128, [email protected]. **Aerospace Engineer, Computational Aerosciences Branch, MS 128, [email protected]. †† Aerospace Engineer, Atmospheric Flight and Entry Systems Branch, MS 489, [email protected]. ‡‡ Research Scientist, Aerothermodynamics Branch, MS 230-3, [email protected]. of nozzles (0, 1, 3, or 4 nozzles), thrust coefficient (CT = T/qA = 2, 3), angle of attack (0, 12, and 20 degrees), and roll angle (0 and 180 degrees). Time-accurate CFD simulations were conducted

References [1] Braun, R. D. and Manning, R. M., “Mars Exploration Entry, Descent, and Landing Challenges," Journal of Spacecraft and Rockets, Vol. 44, No. 2, Mar-Apr 2007. [2] Steinfeldt, B. A., Theisinger, J. E., Korzun, A. M., Clark, I. G., Grant, M. J., and Braun, R. T., “High Mass Mars Entry, Descent, and Landing Architecture Assessment," AIAA Paper 2009-6684, Sep 2009. [3] Zang, T. A., Dwyer-Dianciolo, A. M., Kinney, D. J., Howard, A. R., Chen, G. T., Ivanov, M. C., Sostaric, R. R., and Westhelle, C. H., “Overview of the NASA Entry, Descent and Landing Systems Analysis Study," AIAA Paper 2010-8649, Aug 2010. [4] Edquist, K. T., Dyakonov, A. A., Korzun, A. M., Shidner, J. D., Studak, J. W., Tigges, M. A., Kipp, D. M., Prakash, R., Trumble, K. A., and Dupzyk, I. C., “Development of Supersonic Retro-Propulsion for Future Mars Entry, Descent, and Landing Systems," AIAA Paper 2010-5046, Jun 2010. [5] Korzun, A. M. and Braun, R. D., “Performance Characterization of Supersonic Retropropulsion Technology for High-Mass Mars Entry Systems," Journal of Spacecraft and Rockets, Vol. 47, No. 5, Sep-Oct 2010. [6] Korzun, A. M., Braun, R. D., and Cruz, J. R., “Survey of Supersonic Retropropulsion Technology for Mars Entry, Descent, and Landing," Journal of Spacecraft and Rockets, Vol. 46, No. 5, Sep-Oct 2009. [7] Trumble, K. A., Schauerhamer, D. G., Kleb, W. L., Carlson, JR., Buning, P. G., Edquist, K. T., and Barnhardt, M. D., “An Initial Assessment of Navier-Stokes Codes Applied to Supersonic Retro-Propulsion,” AIAA Paper 2010-5047, June 2010. [8] Kleb, W. L., Carlson, JR., Buning, P. G., Berry, S. A., Rhode, M. N., Edquist, K. T., Schauerhamer, D. G., Trumble, K. A., Sozer, E., “Toward Supersonic Retropropulsion CFD Validation,” Accepted to 20th AIAA Thermophysics Conference, June, 2011. [9] Wright, M.W., White, T., and Mangini, N., “Data Parallel Line Relaxation (DPLR) Code User Manual Acadia – Version 4.01.1,” NASA/TM‐2009‐215388, October 2009. [10] Anderson, W.K. and Bonhaus, D.L., “An Implicit Upwind Algorithm for Computing Turbulent Flows on Unstructured Grids,” Journal of Computational Physics, Vol. 128, No. 2, 1996, pp. 391‐408.

Poster Session 4 172 [11] Anderson, W. K., Rausch, R. D., and Bonhaus, D. L., “Implicit/Multigrid Algorithm for Incompressible Turbulent Flows on Unstructured Grids,” Journal of Computational Physics, Vol. 128, No. 2, 1996, pp. 391–408. [12] Buning, P. G., Jespersen, D. C., Pulliam, T. H., Klopfer, G. H., Chan, W. M., Slotnick, J. P., Krist, S. E., and Renze, K. J., “Overflow User’s Manual,” NASA Langley Research Center, Hampton, VA, 2002. [13] Nompelis, I. N., Drayna, T., and Candler, G. V., \A Parallel Unstructured Implicit Solver for Hypersonic Reacting Flow Simulations," AIAA Paper 2005-4867, June 2005. [14] Berry, S. A., Rhode, M. N., “Supersonic Retro-Propulsion Test 1853 in NASA LaRC Unitary Plan Wind Tunnel Test Section 2,” NASA EDL-01-TR-9178, Nov. 2010. [15] Berry, S. A., Laws, C. T., Kleb, W. L., Rhode, M. N., Spells, C., Mccrea, A. C., Trumble, K. A., Schauerhamer, D. G., Oberkampf, W. L., “Supersonic Retro-Propulsion Experimental Design for Computational Fluid Dynamics Model Validation,” Accepted to 20th AIAA Thermophysics Conference, June, 2011.


ENTRY AND POWERED DESCENT GUIDANCE FOR MARS ROBOTIC PRECURSORS

Sostaric, Ronald R.; Garcia-Llama, E. Powell, R. W.

ABSTRACT Future crewed missions to Mars require improvements in landed mass capability beyond that which is possible using state-of-the-art Mars Entry, Descent, and Landing (EDL) systems. Current systems are capable of an estimated maximum of 1-1.5 metric tons (MT), while human Mars studies require 20-40 MT. A set of technologies were investigated by the EDL Systems Analysis (SA) project to assess the performance of candidate EDL architectures. A single architecture was selected for the design of a robotic precursor mission whose objective is to demonstrate these technologies. In particular, inflatable aerodynamic decelerators (IADs) and supersonic retro-propulsion (SRP) have been shown to have the greatest mass benefit and extensibility to future exploration missions. In order to evaluate these technologies and develop the mission, candidate guidance algorithms have been coded into the simulation for the purposes of studying system performance. These guidance algorithms include entry and powered descent (in addition to aerocapture, which is the subject of another paper). The performance of the algorithms for each of these phases in the presence of dispersions has been assessed using a Monte Carlo technique. The aerocapture maneuver is used to slow the vehicle from a hyperbolic orbital energy to an elliptical energy by utilizing the atmospheric drag. The mission design assumes that a period of time is spent in orbit for checkout prior to entry. A de-orbit burn is then performed to initiate the entry sequence and drive the vehicle toward the atmosphere. Once the atmospheric drag forces increase above a threshold, bank modulation is accomplished according to calculations provided by the entry guidance. Two entry guidance methods have been incorporated: a numerical predictor corrector, and the Apollo Entry Guidance. The numerical predictor corrector integrates a simplified set of the equations of motion, and iterates on a specified control parameter (e.g. bank angle command) to determine the optimum. The Apollo Entry Guidance is an analytical terminal point control method which calculates control parameter gains based on driving the final state to a pre-determined value. A description of each guidance and performance results will be presented. Following the entry phase and jettison of the heat shield, a supersonic retro-propulsion (SRP) phase is initiated. During the SRP phase, the remaining vehicle velocity is reduced using a propulsive method with thrust magnitude and thrust direction calculations provided by the guidance. The guidance can dynamically retarget the landing site real-

Poster Session 4 174 time to avoid hazards. The SRP phase culminates with safe touchdown on the Martian surface. A description of the guidance, powered descent performance, divert performance, and some further considerations for safe landing will be included.


MULTI-MISSION EARTH ENTRY VEHICLE DESIGN TRADE SPACE AND CONCEPT DEVELOPMENT STATUS

(VERSION 2.0)

Robert W. Maddock

NASA Langley Research Center Engineering Directorate Atmospheric Flight and Entry Systems Branch

ABSTRACT

The Multi-Mission Earth Entry Vehicle (MMEEV), directed as part of the In-Space Propulsion Technology (ISPT) Program, is based on the Mars Sample Return (MSR) EEV design and was first introduced at IPPW6. The MMEEV is a flexible design concept which can be optimized and/or tailored by any sample return mission, including lunar, asteroid, comet, and planetary (including Mars), to meet that mission’s specific requirements. By leveraging common design elements, this approach could significantly reduce the risk and associated cost in development of EEV technologies across all sample return missions by providing significant cross-feed and feed-forward in the areas of design and development, trade space analyses, testing, and even flight experience. This presentation describes the current status of the MMEEV concept development, with focus placed on the changes and updates made, specifically in the parametric vehicle model, since version 1.0 was completed in early 2010 (and presented at IPPW8). An overview of a MATLAB vehicle model, which includes increased fidelity in the areas of iterative sizing for payload accommodation, impact attenuation sizing based on impact velocity estimates, structural sizing based on estimates of entry loads, and increased definition of the payload itself, is presented. In addition, application of this vehicle model in both a “standard” and “MSR-like” mode is described. Validation of this model, in both geometry and mass properties, using the Pro/Engineer (ProE) software is also discussed. Engineering estimates of MMEEV vehicle and trajectory performance, generated using the NASA Langley Research Center’s Program to Optimize Simulated Trajectories (POST2) 6-DOF simulation software, across the entirety of the vehicle and mission trade space are also presented, with emphasis on comparisons with the version 1.0 results. Future plans for continued MMEEV development are also discussed. These include the next steps in development of current models, as well as the addition of new models, such as an aftbody TPS MER and thermal soak. Plans for integration of the MMEEV multi-discipline analysis models into the System Analysis of Planetary Entry, Descent and Landing (SAPE) tool, originally developed under ISPT for utilization on aerocapture mission studies, are also presented.


THERMAL SOAK ANALYSIS OF SPRITE PROBE

P. Agrawal1, Y.K. Chen2, D.K. Prabhu1 D. Empey3, E. Venkatapathy2, J. Arnold2

ERC1, NASA Ames2, Sierra Lobo, Inc.3

ABSTRACT

A concept called SPRITE (Small Probe Reentry Investigation for TPS Engineering) has been developed at NASA Ames Research Center to facilitate arc-jet testing of a fully instrumented prototype probe at flight scale [1]. Besides demonstrating the feasibility of testing a flight-scale model and the capability of an on-board data acquisition system, another objective for this project was to investigate the capability of simulations tools to predict thermal environments around the probe/test article and its interior. The present paper summarizes results of thermal analyses that were performed during the early design phase for the SPRITE project to provide input to the design team, as well as post-test analyses to obtain the temperature histories of the probe, substructure and payload and compare them against measured data. The results reported here have been obtained using a commercial finite element (FE) solver MSC.Marc [2], which supports fully transient, non-linear, coupled thermal/mechanical FE analyses. In the test design phase, several conduction and re-radiation based thermal analyses were performed for variations in design parameters. The requisite surface heat-flux distributions (used as boundary conditions for the 2D axi-symmetric FE model) were obtained using DPLR [3] for arc-heated flow fields. These results helped guide the test and design teams in – (1) selecting the material for the substructure and container box for the data acquisition system, (2) determining the exposure time for arc-jet test article, and (3) determining thermal pathways for suitably placing thermocouples in the test article. For post-test predictions of temperature histories for the probe and internal payload during the cool down process, the fidelity of the modeling was improved through the integration of a materials response code, TITAN [4], in the analyses. The temperature maps obtained from TITAN were imposed on the finite element model at the end of the heat pulse. Doing so ensured that ablation and pyrolysis during the exposure were included in the analysis. The cooling process and heat transfer from the forebody and aft TPS to the substructure and payload were then analyzed using the MARC solver. Furthermore, the heat generated by the battery (installed to power the internal data acquisition system) was also included in the model. The temperature histories predicted by the FE model were in good agreement with data obtained from thermocouples placed on the battery and metal container. The finite element analyses were able to predict the time and magnitude of the peak heat in the aluminum box and the battery within ± 5 °C. This approach will be further advanced to develop thermal soak models for Multi mission earth entry vehicles. Acknowledgments: The present work was supported by the Entry Systems and Technology Division, NASA Ames Research Center and Contract No. NNA10DE12C to

Poster Session 4 177 ERC, Incorporated. The authors acknowledge SPRITE team members K. Peterson, K. Skokova, G. Swanson and arc- jet test team for providing valuable test data. References [1] Howard, Austin R., Prabhu, Dinesh, K., Venkatapathy, Ethiraj, and Arnold, James, O.: “Small Probes as Flight Test Beds for Thermal Protection Materials” Proceedings of the 7th International Planetary Probe Workshop, Barcelona, Spain, 2009. [2] “Multidimensional Testing of Thermal Protection Materials in the Arcjet Test Facility”, Parul Agrawal, Donald T. Ellerby, Mathew R. Switzer,Thomas H. Squire, proceedings, 10th AIAA/ASME Joint Thermophysics and Heat Transfer Conference Chicago, Illinois, June 28- July1 2010. [3] Prabhu, D. K., Saunders, D. A., Oishi, T., Skokova, K. A., Santos, J., Fu, J., Terrazas-Salinas, I., Carballo, J. E., Jr., and Driver, D. M., “CFD Analysis Framework for Arc-Heated Flowfields, I: Stagnation Testing in Arc-jets at NASA ARC,” AIAA Paper 2009-2081, AIAA Thermophysics Conference, San Antonio, TX, June 2009. [4] Chen, Y.-K., and Milos, F.S., “Two-Dimensional Implicit Thermal Response and Ablation Program for Charring,” Journal of Spacecraft and Rockets, Vol. 38, No. 4, July-August 2001, pp. 473-481.


DESIGN CHOICE CONSIDERATIONS FOR VEHICLES UTILIZING SUPERSONIC RETROPROPULSION

Ashley M. Korzun(1), Ian G. Clark(2), Robert D. Braun(3)

(1)Georgia Institute of Technology, Email: [email protected], (2) Email: [email protected], (3) Email: [email protected]

ABSTRACT

The entry, descent, and landing (EDL) systems for the United States' six successful landings on Mars and the 2011 Mars Science Laboratory (MSL) rely heavily on extensions of technology developed for the Viking missions of the mid 1970s.i To achieve NASA's long-term exploration goals at Mars, including human exploration, technologies are needed that enable substantial improvements in landed mass and landing accuracy as compared to the expected performance of MSL. Supersonic deceleration has been identified as a critical deficiency in extending Viking-heritage technologies to the high mass, high ballistic coefficient systems required to achieve these goals.i,ii As the development and qualification of significantly larger supersonic parachutes is not a viable path forward to increase landed mass capability to 10+ metric tons, alternative approaches must be developed.i Supersonic retropropulsion (SRP), or the use of retropropulsive thrust while an entry vehicle is traveling at supersonic conditions, is one such alternative approach.i,ii,iii Work has been completed to define mission scales and relevant operating conditions for which SRP may be beneficial.iv As part of one study, the propulsion system was sized to simultaneously minimize the mass and volume of a generic multiple nozzle propulsion system in order to achieve a designated subsonic condition (altitude and velocity).iv In contrast, without accounting for the mass and volume of the propulsion system, NASA’s EDL Systems Analysis studyii minimized the propellant mass required for propulsive deceleration with solutions resulting in a maximum vehicle thrust-to-weight three times greater than that derived from consideration of both propulsion system mass and volume. SRP aerodynamic effects were not considered in the definition of either configuration. Continued work is establishing a minimum fidelity requirement on SRP aerodynamics models for systems analysis in support of developing a capability to evaluate and compare a number of SRP concepts against one another and also against alternative decelerator concepts. How these SRP concepts are to be derived and how much consideration should be given to SRP aerodynamics in defining the configurations remain open questions. Significant effort in the wind tunnel testing of small supersonic retropropulsion models took place in the 1960s and early 1970s, though the combined data is not of sufficient

Poster Session 4 179 breadth to draw detailed conclusions on the effects of utilizing SRP within a full-scale EDL architecture. The aerodynamic-propulsive interaction arising from SRP significantly alters the static aerodynamic characteristics of the vehicle.iii However, no work has yet attempted to develop an SRP configuration targeting an advantageous relationship between the SRP aerodynamic-propulsive interaction and the system performance of the powered descent vehicle. The planned outcome of the work to be presented is a point design of a flight-relevant SRP configuration that considers the sensitivities of parameters governing SRP aerodynamics to variation in physical quantities related to vehicle configuration and system performance. Momentum transfer within the flowfield governs the change in the surface pressure distribution of the vehicle, and accordingly, governs the integrated change in the vehicle’s static aerodynamic characteristics. Parameters governing SRP aerodynamics can be identified using both experimental trends in the literature and analytical statements of momentum transfer within the SRP flowfield. These analytical statments are specific to highly under-expanded jet flows, contact surfaces, and blunted bodies in supersonic flows. Experimental efforts have determined that the flowfield structure and the flowfield stability for SRP are highly dependent on the retropropulsion configuration and the strength of the retropropulsion exhaust flow, relative to the strength of the freestream flow. For a fixed set of freestream conditions, thrust coefficient is a force coefficient based directly on the ideal retropropulsive nozzle thrust. As a limited example of parameter identification, the expression of thrust coefficient based on ideal nozzle thrust can be translated into an expression that is dependent on the ratio the total pressure of the exhaust flow to the total pressure of the freestream flow, the freestream Mach number, the nozzle expansion ratio, composition of the freestream and exhaust flows, and the ratio of the nozzle exit area to the reference area of the vehicle. These parameters are directly related to the operating conditions, propulsion system composition, nozzle geometry, vehicle configuration, and required propulsion system performance, all of which can be considered to be design choices. Investigation into the sensitivities of such parameters to variation in physical quantities related to vehicle configuration and system performance will allow for conclusions to be drawn about the impact of design choices related to system performance on the change in the vehicle’s static aerodynamic characteristics. An initial understanding of the significance of powered descent vehicle configuration on the change in the vehicle’s static aerodynamic characteristics arising from SRP and the relationship to other vehicle performance-based metrics that traditionally determine vehicle configuration is necessary for identification of the types of configurations to be prioritized for SRP concept development. i Braun, R. D., and Manning, R. M., “Mars Exploration Entry, Decent and Landing Challenges,” Journal of Spacecraft and Rockets, Vol. 44, No. 2, 2007, pp. 310-323. ii Zang, T. A., and Tahmasebi, F., “Entry, Descent and Landing Systems Analysis Study: Phase 1 Report,” NASA TM 2010-216720, July 2010. iii Korzun, A. M., Braun, R. D., and Cruz, J. R., “Survey of Supersonic Retropropulsion Technology for Mars Entry, Descent, and Landing,” Journal of Spacecraft and Rockets,

Poster Session 4 180 Vol. 46, No. 5, 2009, pp. 929-937. iv Korzun, A. M., and Braun, R. D., “Performance Characterization of Supersonic Retropropulsion for High-Mass Mars Entry Systems,” Journal of Spacecraft and Rockets, Vol. 47, No. 5, 2010, pp. 836-848.

Poster Session 5 – Science Instrumentation


THE STUDENT RAINDROP DETECTOR (SRD): AN INSTRUMENT FOR MEASURING METHANE RAIN ON

TITAN

Allison Tucker1, Gabriel Wilson1, Hieu Truong1, Tim Kunz1, Kysen Palmer1, Colton Therrian1, Jason W. Barnes1, David H. Atkinson1

Ralph D. Lorenz2

University of Idaho1, JHU/APL2

ABSTRACT

Besides Earth, Saturn’s moon Titan is the only other place in the solar system where rain falls onto a solid surface. Although we have evidence that the rain does interact with the surface from erosion patterns, the actual rain itself has not yet been directly measured. Future in situ Titan probes could make this measurement. We have developed a demonstration instrument capable of detecting Titan’s rain. The device is based on a piezoelectric microphone, an instrument concept that has been space qualified on the European Giotto mission to comet Halley, where it listened for dust impacts. The piezoelectric detector is attached to a 10cm-square strike plate that will be exposed to the Titan sky. We will show how monitoring the piezo voltage over time allows us to identify raindrop hits and ascertain their momentum, from which we can calculate the drop’s radius given knowledge of the local atmospheric density. In our poster, we will present the instrumental design and the results of tests in both ambient and Titan-relevant environments. This work has been done by undergraduates at the University of Idaho as an Engineering Senior Design Project with the goal of developing the instrument to TRL 6 for use on the AVIATR Titan Airplane mission. We call it the Student Raindrop Detector (SRD) because it would be included in a mission proposal as a student-built element. It could also be used on Titan atmospheric probes, landers, or airships. Its application to a balloon is not immediately clear, however, given that a gondola will necessarily be in the rain shadow of the balloon itself.


PLANETARY POLARIZATION NEPHELOMETER

Don Banfield(1), Adam Saltzman(1)

(1)Cornell Astronomy, 420 Space Sciences, Cornell University, Ithaca, NY 14853, USA, Email: [email protected]

ABSTRACT

We have completed a breadboard validation for a planetary polarization nephelometer, raising this instrument from just a concept through to TRL 4 using PIDD funding. We are currently seeking further PIDD funding to continue this instrument development up through to TRL 6, culminating in its demonstration on a stratospheric balloon as a proxy for a planetary flight mission through an exotic and relatively inhospitable atmosphere. Our instrument is aimed at determining the characteristics of the aerosols that are present in essentially all planetary atmospheres. These aerosols are important to understand, not only to catalog the particles that are the visible faces of most planets, but also because they have significant impact on the climates and atmospheric dynamics of these planets, and the details of their makeup can help us to understand other questions about the planets, such as processes active on or below the surfaces. For Venus, the aerosols contain a significant fraction of the Sulfur compounds in the atmosphere. Our instrument is crucial for a full inventory of these compounds in Venus’ atmosphere and further for understanding the processes that get them there from the surface. Our instrument would also help identify the unknown blue absorber in Venus’ clouds that accounts for 25% of its powerful greenhouse. For Jupiter and the other giant planets, our instrument would definitively identify the aerosol layers that we see and the altitude levels at which winds are tracked. For Titan, we would be able to more fully understand the absorption and emission of radiation within this extended atmosphere; both processes have a large influence on the climate and dynamics of Titan’s atmosphere. The polarization nephelometer uses a novel approach to the illuminating beam to allow us to extract more information from the light scattered back from adjacent aerosols than is typical in predecessor nephelometers. In addition to measuring the intensity phase function of the light scattered off adjacent aerosols, our instrument also measures the polarization ratio phase function of the aerosols as well. The polarization ratio phase function adds significantly more information about the particle properties, removing ambiguities that intensity phase functions alone leave. We are able to extract this information from the scattered light through fast modulation of the polarization of the illuminating beam, combined with temporal analysis of the scattered light. To ruggedize the instrument for use on planetary descent probes in harsh planetary atmospheres, we use only solid state and temperature insensitive techniques to modulate the polarization of the illuminating beam. We also only expose optical fibers to the harsh environment outside of the spacecraft hull, holding all the electronics, lasers and detectors within the

Poster Session 5 184 hull of the spacecraft where thermal protection is presumably much greater. We anticipate a flight version of our instrument would require about 6W of power, 1kg of mass and 0.7L of volume. It would be a good addition to any probe sent to a planet or satellite with an atmosphere, or even to the moon to analyze lofted moon dust.


SCIENCE AND EDUCATION WITH MARS EXPRESS' VISUAL MONITORING CAMERA

(VMC)

H.S.Griebel*1, T.Ormston1, M.Denis2, J.Landeau-Constantin2, D.Scouka2,3, L.Griebel4, C.Scorza5, M.Frommelt5

1VEGA Space, 2European Space Operations Centre, ESA

3EJR-Quartz, 4Mars Society Germany, 5Haus der Astronomie, Max Planck Institute for Astronomy e-mail:

[email protected], [email protected], [email protected], [email protected], [email protected], [email protected], [email protected]

ABSTRACT

Mars Express, the European Space Agency’s mission to the Red Planet launched in 2003, carries a small camera designed to provide ‘visual telemetry’ of the separation of the Beagle-2 lander. This activity was completed in 2003. The camera, known as VMC, was reactivated by the flight control team in 2007 as part of a small student thesis project and has since evolved into a unique public outreach tool. With its small aperture, a 640x480 pixel colour CMOS sensor and a 30ox40o field of view, the camera’s optical properties are similar to those of a typical PC webcam or mobile phone camera. Its turn-around time from observation to reception of imagery on Earth is therefore extremely short, and the processing power required to generate meaningful results is low. On the basis of non-interference with science operations, the camera now frequently provides internet users with stunning, up-to-date and unaltered vistas without relying on any intermediate process. Interested users are therefore provided a unique ‘hands-on- Mars’ opportunity: they can download the latest images first-hand or browse the extensive archive to conduct their own research and even publish their results in our ESA VMC user forum. This project proved so successful that in 2010, the VMC was used for two dedicated educational projects. Two groups of school children were given the opportunity to conduct their own observations of Mars. One group was tutored by the Mars Society Germany, the other by the Astronomy teacher of a German secondary school. In addition to this, a first-of-its-kind stop-motion animation was created, showing one full orbit of Mars Express, providing a first-hand demonstration of Kepler’s laws of celestial motion and the rotation and axial tilt of Mars. Some of Mars’s most prominent surface features as well as the moon Phobos were also shown. The outstanding results of these activities inspired us to further develop this effort in cooperation with the Haus der Astronomie / Center for Astronomy Education and

Poster Session 5 186 Outreach at the Max-Planck-Institute Campus, providing a regular opportunity for educators and schools to use actual Mars observations as part of their educational outreach material. In this paper we demonstrate how the Mars Express VMC is operated without hindrance to primary science mission; how schools, scientists and the interested public have benefited from the data thus provided and how we intend to support the younger generations in their pursuit of first-hand, ‘citizen-science’ in the future.


DEVELOPMENT OF INSTRUMENTATION FOR HYPERSONIC INFLATABLE AERODYNAMIC

DECELERATOR CHARACTERIZATION

Gregory T. Swanson1,2, Alan M. Cassell 2

1Santa Clara University, Department of Electrical Engineering, 500 El Camino Real, Santa Clara, CA 95053 USA Email:[email protected]

2[ERC Incorporated], Entry Systems and Vehicle Development Branch, NASA Ames Research Center, Moffett Field, CA 94035 USA

Email:[email protected]

ABSTRACT To realize the National Aeronautics and Space Administration’s (NASA) goal of sending humans to Mars, development of technologies to facilitate the landing of heavy payloads are being explored. Current entry, descent, and landing technologies are not practical for heavy payloads due to mass and volume constraints dictated by limitations imposed by launch vehicle fairings. Therefore, technologies are now being explored to provide a mass- and volume-efficient solution for heavy payload capabilities, including Inflatable Aerodynamic Decelerators (IADs) [1]. Consideration of IADs for space applications has prompted the development of instrumentation systems for integration with flexible structures to characterize system response to flight-like environment testing. This development opportunity faces many challenges specific to inflatable structures in extreme environments, including but not limited to physical flexibility, packaging, temperature, structural integration and data acquisition [2]. In the fall of 2011, three large scale Hypersonic Inflatable Aerodynamic Decelerators (HIAD) will be tested in the National Full-Scale Aerodynamics Complex’s (NFAC) 40’ by 80’ wind tunnel at NASA Ames Research Center. The test series will characterize the performance of a 3.0 m, 6.0 m, and 8.3 m HIAD at various angles of attack and levels of inflation during flight like loading. To analyze the performance of these test articles as they undergo aerodynamic loading, an instrumentation system has been developed. This system will utilize new experimental sensing concepts, developed by the large scale HIAD instrumentation team, in addition to traditional wind tunnel sensing techniques in an effort to improve test article characterization. The instrumentation system will target HIAD pressure distribution, flexible aeroshell static and dynamic deformation, rigid hardware stress, torus inflation pressure, and flexible aeroshell structural strap loading. Pressure measurements on the rigid HIAD hardware and in the inflatable tori will be conducted using traditional systems, while MEMS pressure sensors will be integrated

Poster Session 5 188 into the flexible aeroshell in a new experimental concept. In addition, an approach to monitor static and dynamic deformation in the HIAD’s flexible aeroshell will be characterized using string potentiometers to provide a reference distance from the supporting test structure. Stress seen by the rigid hardware will be characterized by traditional strain gages. During this test series, we will also explore many developmental embedded sensing concepts for space flight test applications. Bend sensors will be incorporated into the HIAD Thermal Protection System (TPS) between structural tori presenting an indication of the deformation seen during aerodynamic loading. Flight-like configurations of the MEMS pressure sensors will also be tested accounting for on orbit and re-entry environmental constraints. Finally, MEMS accelerometers will be co-located with the string potentiometers as a proof of concept to measure deflection during flight. All developmental embedded sensing concepts will utilize flexible printed circuit board technology in order to meet the stringent launch packaging requirements of the HIAD aeroshell minimizing the risk of initiating punctures in the flexible materials. Additionally, a subset of the embedded sensing concepts will employ wireless data transfer reducing the wiring bundle mass and complexity. [1] Clark, I.G., Hutchings, A.L., Tanner, C.L., Braun, R.D., “Supersonic Inflatable Aerodynamic Decelerators for Use on Future Robotic Missions to Mars,” Journal of Spacecraft and Rockets, Vol. 46, No. 2, March 2009. [2] Brandon, E.J. et al., Structural Health Management Technologies for Inflatable / Deployable Structures: Integrated Sensing and Self- Healing, Acta Astronautica (2010), doi:10.1016/j.actaastro.2010.08.016


MARS MICROPHONE 2016: A UNIQUE OPPORTUNITY FOR STUDENT INVOLVEMENT

*A. Minier1, *W. Rapin1, D. Perez Escobar1, D. Mimoun1 and the Mars Microphone Team 2

1Institut Supérieur de l’Aéronautique et de l’Espace,*Undergraduate Students 2 see

bit.ly/MM2016 for the team description

ABSTRACT

General outline The Mars Microphone is an E/PO experiment proposed in the frame of the ExoMars Entry and Descent Module of the ExoMars Trace Gas orbiter. It aims at retrieving the first sounds ever recorded from Mars. Besides of the technical and science team, its development involves undergraduate and graduate students. The experiment will be built on the heritage of previous Mars microphone experiments led by Berkeley SSL and the Planetary Society. The main scientific objectives include basic atmospheric investigation, analysis of dust devils and wind vortexes and the capture of sounds related to atmosphere electrical activity and meteoritic impacts.

Design of the Mars Microphone The microphone relies on a simple and robust design, a small mass (about 50g) and size (50x50x20mm) and a low power budget (around 150mW). It is based on a very limited set of well-known and previously space-qualified components in order to minimize development risks. From a functional point of view, the microphone is a very simple and robust state machine, fully under the control of the CEU. Its life duration in the surface operational phase will be very short, about 4 days, thus reducing risks of failure and development constraints. Three different configurations including up to three microphones for extended science experiments are considered, depending on the possible on-board resources allocation.

Poster Session 5 190 Student Contributions The simple design and development approach of the Mars Microphone combined with the exciting perspective of its mission make it an ideal support for education at the University level. Students will (and already are!) deeply involved at all stages of the development, operations and post-flight analysis of the instrument’s mission, as show on the following diagram:

While waiting for the payload selection, we are currently setting up an experiment for ground calibration tests carried out in a pressurized Martian Simulation Chamber. This set-up will allow performing realistic measurements of the propagation of acoustic waves in the Martian atmosphere and will give us a keener understanding of the scientific phenomena to be detected on Mars by the microphone. It will also give us a preliminary validation of the signal to noise ratio. Other on-going activities include the electronics design, the design of the mechanical box, contribution to various environment test setups and reports, and participation in outreach activities. References [1]The Mars Microphone team, 2011, Mars Microphone 2016, proposal for the Exo/mars EDM [2] Nornberg P et. al. the new Danish/ESA Mars simulation wind Tunnel at Aarhus University. [3] Williams, J.-p. (2001), Acoustic environment of the Martian surface. Journal of Geophysical Research 106, 5033-5042 [4] Sparrow, V.W. (1999): Acoustics on the planet Mars: A preview. The Journal of the Acoustical Society of America, 106, 2264.

Poster Session 6A – New Technologies

Poster Session 6A 192

TDNR: A MODULAR NANO-ROVER PLATFORM FOR NETWORKED PLANETARY MISSIONS

Abraham Rademacher (1), Amardeep Singh(2), Jasvir Singh(3), Jose Cortez (4), Kavinda Wittahachchi(5), Mikhail Paremski(6), Yawo

Ezunkpe(7), Dr. Periklis Papadopoulos(8), Marcus S. Murbach(9), Bob Feretich(10)

(1) Graduate, AE, One Washington Square, San Jose, CA 95192, United States, Email:

[email protected] (2)Undergraduate,ME,One WashingtonSquare,SanJose,CA95192,UnitedStates,Email:[email protected] (3)

Undergraduate, ME, One Washington Square, San Jose, CA 95192, United States, Email: [email protected] (4) Undergraduate, AE, One Washington Square, San Jose, CA 95192, United States, Email: [email protected] (5) Undergraduate,

ME, One Washington Square, San Jose, CA 95192, United States, Email: [email protected] (6) Undergraduate, AE, One Washington Square, San Jose, CA

95192, United States, Email: [email protected] (7) Undergraduate, AE, One Washington Square, San Jose, CA 95192, United States, Email:

[email protected] (8) Advisor, Professor, MEA Department, One Washington Square, San Jose, CA 95192, United States, Email: [email protected] (8)

Advisor, Principal Investigator SOAREX Flight Project, NASA Ames Research Center Moffett Field, CA 94035- 1000, United States, Email: [email protected] (10)

Advisor, REF Research, LLC. P.O. Box 20098, San Jose, CA 95160, United States, Email: [email protected]

ABSTRACT

A 1kg nano-rover platform is designed to extend the benefits recognized in nano-satellites into the domain of planetary rovers. This nano-rover concept, hereto referred to as the Tube Deployable Nano Rover (TDNR), provides a unique solution to planetary exploration; augmenting traditional high-cost, high-risk, rover/sensorsystems by expanding the traditional networked planetary probe configuration. With the TDNR, a conventionalnetworked mission can be enhanced with a sub-network of modular, low-cost, nano-rovers. Each nano-rover has modular payload capacity, and each rover can be equipped with independent instruments to cover a range of scientific missions. The TDNR employs a unique expanding wheel design based on Hoberman geometry and a tubular shape allowing the two -wheeled rover to efficiently package within a SCRAMP (Slotted Compression RAMP) entry decent and landing probe. Integrating the TDNR within the SCRAMP leverages several years of research at NASA Ames in next-generation self-stabilizing planetary entry vehicles, which are specifically designed for companion missions. A prototype TDNR is designed and constructed using off-the-shelf

Poster Session 6A 193 hardware at San Jose State University equipped with a Beagle Board main processing board, camera and a 3 axis accelerometer for sensing and navigation, wireless N adapter and foldable antenna for command and communication, and as described in the concept, a payload bay to accommodate mission specific sensors. A prototype TDNR is constructed to identify components required, demonstrates feasibility of the concept, and serves as a development platform to further refine the concept.


ANALYSIS OF ANOMALOUS VARIATIONS IN HIGH ALTITUDE BALLOON ASCENT RATES NEAR THE

TROPOPAUSE

Walter Taresh*, Kevin Ramus, Kim Baird, Carlos Gonzalez, Gabe Wilson, Rory Riggs, George Korbel, David H. Atkinson, and the Idaho

Near Space Engineering Team

University of Idaho e-mail: [email protected]

ABSTRACT High altitude balloons provide a simple, inexpensive, and reliable means of studying planetary atmospheres. In recent balloon flights conducted by the University of Idaho’s RISE (Research Involving Student Engineers) Near Space Engineering program, the ascent rate and trajectory of the balloon path has been a major concern. Anomalous variations in ascent rate have been observed near the tropopause on recent flights. Simple models indicate that ascent speed should be essentially constant with altitude. However, near the tropopause a virtually instantaneous reduction in ascent rate of approximately 50% has been observed. Several possible phenomena to explain this effect are being studied, including changes in drag coefficient near Reynolds number of 3x105 and a temperature induced loss of buoyancy due to cooling of the lifting gas during adiabatic expansion of the balloon in the near-isothermal layer above the tropopause (temperature drag effect).


DEVELOPMENT OF AN AUTONOMOUS HIGH ALTITUDE BALLOON CUTDOWN SYSTEM

Kevin Ramus*, Kim Baird, Carlos Gonzalez, Gabe Wilson, Walter Taresh, Rory Riggs, George Korbel, David H. Atkinson, and the Idaho

Near Space Engineering Team

University of Idaho e-mail: [email protected]

ABSTRACT

Balloons are a simple and economical way to carry scientific instrumentation into the upper atmosphere and can provide a platform for atmospheric flight testing of prototype planetary mission instrumentation, reaching elevations up to and beyond 100,000 feet (30,000 m) on Earth. The University of Idaho Near Space Engineering program known as RISE (Research Involving Student Engineers) has now been launching balloons for seven years. Idaho RISE has a data acquisition system that measures atmospheric pressure and temperature as a function of altitude, and a redundant GPS tracking system that provides real time tracking of the balloon system through ascent, decent, and landing, allowing for a quick recovery of the descent package. A cutdown system has been developed by which payloads can be autonomously released based on timer, altitude, or if the balloon drifts outside of a preprogrammed latitude/longitude box. Working with NASA Ames Research Center, Idaho RISE is currently preparing for a flight of Snowflake, a miniature high- precision aerial delivery system developed by the Naval Postgraduate School and University of Alabama at Huntsville to evaluate advanced control, communication and command concepts for autonomously guided parafoil-payload systems. To date, Snowflake has been successfully deployed over 120 times from altitudes of up to 10,000 feet. The goal of the Idaho RISE Snowflake experiments is to provide a platform to deploy Snowflake at and above 30,000 feet and investigating its performance in these conditions. The launches with Idaho entail the 3rd stage of a proposed ISS sample return capability currently under development (SPQR- Small Payload Quick Return) at Ames Research Center.


THE TITAN SKY SIMULATORTM NEW LOW COST CRYOGENIC TEST FACILITY AVAILABLE.

DEVELOPED FOR TITAN BALLOONS BUT SUITABLE

FOR MANY APPLICATIONS

J. Nott

Affiliated to UC Santa Barbara, Nott Technology LLC Email: [email protected]

The Decadal Survey "Vision and Voyages for Planetary Science" recommends developing Titan balloons. The Titan Sky SimulatorTM is a low cost facility for testing balloons and other Titan hardware. An instrumented working Titan balloons has been flown at 95°K.

CONFLICTING REQUIREMENTS exist for testing a Titan balloon: the gas must circulate fast for uniform temperature yet balloons must fly in completely calm. This is

Poster Session 6A 197 achieved with a highly insulated outer chamber. Liquid nitrogen is sprayed in and gas circulated rapidly with a fan. A calm zone is achieved with an inner aluminum cylinder. This arrangement also allows rapid cool-down. It takes many hours for the insulation to reach temperature equilibrium, but this does not matter. Only the circulating gas and inner cylinder need to reach a stable temperature and this is rapidly achieved. AVAILABLE TO THE COMMUNITY An improved Simulator is under development in Santa Barbara, naturally benefiting from of everything learned. This will have significantly better insulation, be much lighter and much and easier to operate. This resource can be available to the Community in two ways, The facility in Santa Barbara is available. Or with low cost [student] labor, building versions of the Simulator will be inexpensive, perhaps ten thousand dollars for materials. Constructing a Simulator might be an excellent undergraduate team project, yet create a useful, long term research facility. Like many good ideas, the basic concept is simple . But execution involves numerous details, a few examples being insulation attachment, nitrogen injection and adapting low cost cameras for cryogenic use. Nott Technology is able to supply this knowhow. CYLINDER: Ladder right give scale.


ONE-WAY UPLINK RANGING FOR ENHANCING PLANETARY WIND MEASUREMENTS

K. Oudrhiri(1), D.H. Atkinson(2), S.W. Asmar(1), S. Bryant(1), T.R. Spilker(1)

(1)Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA 91109-

8099, USA ([email protected]) (2)University of Idaho, Moscow, ID 83844-1023 USA

ABSTRACT

Uplink One-Way Ranging techniques can be used to improve the accuracy of planetary atmospheric wind profiles measured during entry probe descent using Doppler wind techniques. Advances in Radio Science flight instrument technologies and post-processing capabilities allow for the possibility of utilizing a One- Way sequential ranging signal transmitted from Deep Space Network antennas and recorded onboard a probe-mounted Radio Science open-loop receiver with onboard post-processing algorithms to produce precision measurements of probe range and position, thereby significantly improving Doppler retrievals of atmospheric winds. The probe velocity relative to Earth is computed as the derivative of the ranging positional information and is therefore unaffected by any constant biases in the ranging data. In addition, velocities derived from ranging data will not have an error term that grows with the descent time. By providing an accurate Earth-to-probe baseline range and velocity, knowledge of the planet-centered probe descent location can be significantly improved. Additionally, probe measurement of the DSN uplink signal can provide a second projection of the horizontal winds that, when coupled with the probe-orbiter wind projection, will provide the complete horizontal wind vector. To make the measurements fully complementary, the angle between the Earth-to-probe and probe-to-orbiter baselines should be large, and to increase the sensitivity to winds in the probe local horizontal plane, the probe-orbiter and probe-Earth angles should be at a non-zero angle to the probe nadir vector. In this paper we will review opportunities for and benefits of uplink One-Way ranging for enhancing future planetary entry probe wind measurements.

Poster Session 6B - Aeroassist, Experimental Missions and EDL Mission Design

Poster Session 6B 200

SATURN SYSTEM MISSION OPPORTUNITIES USING A TITAN AEROGRAVITY ASSIST FOR ORBITAL

CAPTURE

Robert M. Booher(1) and J.E. Lyne(2)

(1)Undergraduate Student, Department of Mechanical, Aerospace, and Biomedical Engineering, 414 Dougherty Engineering Building University of Tennessee Knoxville, TN

37996-2210 USA Email: [email protected] (2)Associate Professor, Department of Mechanical, Aerospace, and Biomedical

Engineering, 414 Dougherty Engineering Building University of Tennessee Knoxville, TN 37996-2210 USA Email: [email protected]

ABSTRACT

In previous papers authored by our group, the use of an aerogravity assist (AGA) maneuver at Titan for orbital capture about Saturn was evaluated for a Cassini-class vehicle.1 These studies confirmed that the proposed maneuver is a viable alternative to the use of a traditional propulsive orbital insertion maneuver. A vector diagram of the maneuver is shown in Figure 1. In the current paper we build on this concept and discuss mission opportunities for future return voyages to the Saturn system. Specifically, four candidate mission plans were developed based on overall performance including total ΔV, flight duration, launch year, and launch energy. (Initial trajectories were taken from the European Space Agency Advanced Concepts Team website2 and then optimized through SAIC Trajectory Optimizer3.) The candidate trajectories, displayed in Table 1, range from an October 2018 departure date to a departure in December of 2024. Though likely too near in the future for an actual Saturn return, the launch window in 2018 provided the best results of any tested trajectory and is included for comparison. In the table, the “Trajectory type” column indicates the planetary encounter sequence, including gravity assisted flybys. Fig. 1 Vector diagram of AGA maneuver

Once promising mission opportunities had been identified, the four candidate trajectories

Saturn System Mission Opportunities Using a T itan Aerogravity Assist for O rbital

Capture

Robert M . Booher(1)

and J.E . Lyne(2)

(1)Undergraduate Student, Department of Mechanical, Aerospace, and Biomedical Engineering, 414 Dougherty Engineering Building

University of Tennessee Knoxville, TN 37996-2210

USA Email: [email protected]

(2)Associate Professor, Department of Mechanical, Aerospace, and Biomedical Engineering,

414 Dougherty Engineering Building University of Tennessee

Knoxville, TN 37996-2210 USA

Email: [email protected] !

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Once promising mission opportunities had been identified, the four candidate trajectories were then evaluated for their arrival conditions at Titan. The Saturn arrival declination and V were used to find

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Poster Session 6B 201 were then evaluated for their arrival conditions at Titan. The Saturn arrival declination and V∞ were used to find the probe’s velocity relative to Titan (VE,Titan) at an altitude of 1000 km, corresponding to the Titan atmospheric interface. Our coordinate system was defined such that the probe arrives from the negative y direction as Titan orbits posigrade about Saturn at the origin; the Titan-spacecraft intercept position in the Saturn-centered reference frame is designated by the angle Q. The coordinate system for the Saturn system arrival geometry is illustrated in Figure 2, and a graph of intercept position versus atmospheric entry speed for each candidate mission opportunity is shown in Figure 3.

The various mission opportunities were then examined to determine the desirable range of intercept positions that would provide Titan atmospheric entry velocities of 6 to 10 km/s. This velocity range was chosen to avoid the excessive aerothermal environment that would be associated with higher entry speeds. The approach trajectories for each candidate mission opportunity was then compared to Titan ephemeris data, and the Saturn system arrival dates were adjusted to accurately target the optimal Titan intercept positions. Using these final trajectories, the AGA maneuvers were computed using the Program to Optimize Simulated Trajectories (POST).5

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The various missions opportunities were then examined to determine the desirable range of intercept positions that would provide Titan atmospheric entry velocities of 6 to 10 km/s. This velocity range was chosen to avoid the excessive aerothermal environment that would be associated with higher entry speeds. The approach trajectories for each candidate mission opportunity was then compared to Titan ephemeris data, and the Saturn system arrival dates were adjusted to accurately target the optimal Titan intercept positions. Using these final trajectories, the AGA maneuvers were computed using the Program to Optimize Simulated Trajectories (POST).5

The final paper will present a description of the candidate mission opportunities, including details of the interplanetary trajectories and data on the Titan AGA maneuvers.

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The various missions opportunities were then examined to determine the desirable range of intercept positions that would provide Titan atmospheric entry velocities of 6 to 10 km/s. This velocity range was chosen to avoid the excessive aerothermal environment that would be associated with higher entry speeds. The approach trajectories for each candidate mission opportunity was then compared to Titan ephemeris data, and the Saturn system arrival dates were adjusted to accurately target the optimal Titan intercept positions. Using these final trajectories, the AGA maneuvers were computed using the Program to Optimize Simulated Trajectories (POST).5

The final paper will present a description of the candidate mission opportunities, including details of the interplanetary trajectories and data on the Titan AGA maneuvers.

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Poster Session 6B 202 References [1] Philip Ramsey and James Evans Lyne, “An Investigation of Titan Aerogravity Assist for Capture into Orbit About Saturn,” Journal of Spacecraft and Rockets. Vol. 43, No. 1, pp. 231-233, Feb. 2011. [2] Mission Analysis Advanced Concepts Team, European Space Agency. Nov. 2010. <http://www.esa.int/gsp/ACT/mad/op/AdvancesInGO/SaturnDatabase/SemanticSaturn.htm>. [3] SAIC Trajectory Optimizer, Science Application International Corporation. [4] Philip Ramsey and James Evans Lyne, “Enceladus Mission Architecture Using Titan Aerogravity Assist for Orbital Capture About Saturn,” Journal of Spacecraft and Rockets. Vol. 45, No. 3, pp. 635-638, Feb. 2011. [5] Brauer, G. L., Cornick, D. E., Olson, D. W., Peterson, F. M., and Stevenson, R., “Program to Optimize Simulated Trajectories (POST),’ NASA CR NAS1-18147, Sept. 1989.


AERODYNAMIC STABILITY OF BLUNTED-CONE ENTRY VEHICLES

Daniel R. Ladiges*, Eleanor C. Button, Charles R. Lilley, Nicholas S. Mackenzie, Edward Ross, And John E. Sader

Department of Mathematics and Statistics, The University of Melbourne, Victoria 3010,

Australia e-mail: [email protected]

ABSTRACT

The heat shield of a spacecraft provides protection against the extreme temperatures that result from aerodynamic heating on atmospheric entry. Additionally, the heat shield serves as an important aerodynamic component of the craft, providing the necessary drag and stability at hypersonic speeds. The shape of the heat shield used varies considerably between spacecraft, and spherical and blunted-cone geometries are often employed. The `blunted-cone' heat shield has been developed through experimental design and computational simulation. Previous work (presented at IPPW7) [1] showed that a generic heat shield shape can also be mathematically derived and gave the maximum stabilizing aerodynamic torque of all possible shapes. The derived shape displays minimum variation in torque due to minor shape changes (such as those caused by ablation), and depends only on the center-of-mass of the craft. The shape corresponds closely to the ‘blunted-cone’ design already commonly used in entry vehicles. This previous derivation was performed for both free molecular (FM) and continuum flows. Importantly, both regimes yield an identical shape. In the present work we extend the FM model to allow for finite entry speeds, i.e., the thermal motion of the gas molecules is considered. This is found to modify the shape, but at high Mach numbers these effects cause very little change to the derived shape. Analytical and numerical results of these effects will be presented. Importantly, the previous work [1] only considered the limiting cases of free molecular and continuum flows. It is thus important to investigate flows in the transition regime, which lie between these limits. To this end, Direct Simulation Monte Carlo (DSMC) [2] simulations have now been used to (i) verify the behaviour of the derived shape in the FM regime, (ii) investigate the transition region between the FM and continuum regimes, and (iii) examine the influence of thermochemical effects. Simulations have been performed for both two-dimensional (for reference) and three-dimensional geometries, both with and without the inclusion of thermochemical effects. Excellent agreement with the previous analytical predictions is obtained in the FM regime. The same shape is found for much of the transition regime. However, deviations are observed as the continuum regime is approached – possible causes of these effects will be discussed. The design of practical heat shields involves numerous competing factors, which include the expected heat load and the craft volumetric efficiency, in addition to aerodynamic

Poster Session 6B 204 stability. We thus emphasize that the presented results focus on only one component of this multi-objective problem. The presented results and simulations provide further information on the validity of the derived shape in Ref. [1]. By accounting for effects that are not analytically tractable (using computational simulations), the shape that gives maximum static stability is derived. This is obtained as a function of entry velocity, gas rarefaction and thermochemical effects. [1] E. C. Button, C. R. Lilley, N. S. Mackenzie and J. E. Sader, “Blunted-cone heat shields of atmospheric entry vehicles”, AIAA Journal, 47, 1784-1787 (2009). [2] Graeme Bird, Molecular Gas Dynamics and the Direct Simulation of Gas Flows, Oxford Engineering Science Series, 1994.


DETERMINATION OF AERODYNAMIC DAMPING COEFFICIENTS OF ENTRY VEHICLES IN TRANSONIC

REGIME

S. Paris, O. Karatekinn, A. Karitonov+, J. Ouvrard*

* Von Karman Institute for Fluid Dynamics Ch. De Waterloo, 72, B-1640 Rhode-St-Genèse, Belgique nRoyal Observatory of Belgium 3 Avenue Circulaire, 1180 Bruxelles,

Belgium + Khristianovich Institute of Theoretical and Applied Mechanics, Siberian Branch Russian Academy of sciences, Institutskaya 4/1, Novosibirsk 630090, Russia

ABSTRACT

Oscillatory motion is a dynamic phenomenon experienced by space capsules upon re-entry to the Earth’s atmosphere. This behavior needs to be well understood for specific geometries to avoid unstable flight. Proper characterization of aerodynamic damping for stability evaluation can allow drogue chute deployment at lower Mach number. The purpose of this work is to characterise the steady and unsteady aerodynamic characteristics of entry vehicles in the range of Mach numbers from low speed to supersonic. We use the EXPERT vehicle for the test case. The aerodynamic derivatives have been determined in several wind tunnels over a wide range of velocity. Experiments have been first carried out using the forced oscillation technique in sting configuration. Transversal rod axis has been used as support for the model in the transonic/ supersonic S1 wind tunnel at The von Karman Institute for Mach numbers going from 0.5 to 2. The static efforts have been measured and the dynamic behaviour has been investigated thanks to the two types of oscillations techniques, namely the free oscillations and the forced oscillations. Finally, different post-processing methods used to extract the damping in pitch parameter have been compared. A significant difference between the different sets of results shows that the support is a very important parameter which can produce significant discrepancy. The sources of uncertainty and the effect of wake flow on dynamic stability are discussed in detail in this paper for free and forced oscillation technique and for the different supports. That shows that a deep analysis of the support interference is needed to improve the quality of the results.


Figure 1 Expert model with transversal support (VKI) and sting support (SibNIA) The von Karman Institute for Fluid Dynamics

8th International Planetary Probe Workshop, 6-10 June, 2011 in Portsmouth, VA

The von Karman Institute for Fluid Dynamics Page 1 of 1

Determination of aerodynamic damping coefficients of entry vehicles in transonic regime

S. Paris!, O. Karatekinn, A. Karitonov+, J. Ouvrard*

* Von Karman Institute for Fluid Dynamics Ch. De Waterloo, 72, B-1640 Rhode-St-Genèse, Belgique

nRoyal Observatory of Belgium 3 Avenue Circulaire, 1180 Bruxelles, Belgium

+ Khristianovich Institute of Theoretical and Applied Mechanics, Siberian Branch Russian Academy of sciences, Institutskaya 4/1, Novosibirsk 630090, Russia

Oscillatory motion is a dynamic phenomenon experienced by space capsules upon re-entry to the Earth’s atmosphere. This behavior needs to be well understood for specific geometries to avoid unstable flight. Proper characterization of aerodynamic damping for stability evaluation can allow drogue chute deployment at lower Mach number. The purpose of this work is to characterise the steady and unsteady aerodynamic characteristics of entry vehicles in the range of Mach numbers from low speed to supersonic. We use the EXPERT vehicle for the test case. The aerodynamic derivatives have been determined in several wind tunnels over a wide range of velocity. Experiments have been first carried out using the forced oscillation technique in sting configuration. Transversal rod axis has been used as support for the model in the transonic/ supersonic S1 wind tunnel at The von Karman Institute for Mach numbers going from 0.5 to 2. The static efforts have been measured and the dynamic behaviour has been investigated thanks to the two types of oscillations techniques, namely the free oscillations and the forced oscillations. Finally, different post-processing methods used to extract the damping in pitch parameter have been compared. A significant difference between the different sets of results shows that the support is a very important parameter which can produce significant discrepancy. The sources of uncertainty and the effect of wake flow on dynamic stability are discussed in detail in this paper for free and forced oscillation technique and for the different supports. That shows that a deep analysis of the support interference is needed to improve the quality of the results.

Figure 1 Expert model with transversal support (VKI) and sting support (SibNIA)


STATISTICAL ENTRY, DESCENT AND LANDING PERFORMANCE RECONSTRUCTION OF THE MARS

PHOENIX LANDER

Soumyo Dutta(1), Ian G. Clark(2), Ryan P. Russell(3), Robert D. Braun(4)

(1)Daniel Guggenheim School of Aerospace Engineering, Georgia Institute of Technology, 270 Ferst Drive NW, Atlanta, GA 30332, USA, Email:

[email protected] (2) Daniel Guggenheim School of Aerospace Engineering, Georgia Institute of Technology, 270 Ferst Drive NW, Atlanta, GA 30332, USA, Email:



[email protected]

ABSTRACT The Phoenix Lander successfully landed on the surface of Mars on May 25, 2008. During the entry, descent and landing (EDL), the vehicle had instruments on-board that took sensed acceleration, angular rates and altimeter measurements. Additionally, satellites orbiting Mars during Phoenix’s entry took range measurements of the descending vehicle. This paper will demonstrate the methodology used to reconstruct the trajectory information from observations from these various sensors. The paper will also present the reconstructed flight trajectory and the atmospheric profile sensed by the vehicle during its landing sequence. Although Phoenix’s trajectory has been reconstructed in the past by NASA, this current reconstruction differs from these past efforts due to the stochastic estimation techniques used to blend the different EDL data types. The estimation algorithm used in this case will be an Extended Kalman filter (EKF), which is adept at reconstructing states and their uncertainties for a non-linear problem. The stochastic nature of the reconstruction uses the inherent uncertainty in the measurement sensor data and propagates these values to quantify the uncertainties associated with the estimated trajectory and atmospheric parameters. The results of this reconstruction can thus allow a statistical comparison of the actual trajectory and atmosphere experienced by the vehicle and what was expected. Moreover, the paper will analyze the aerodynamic performance of the vehicle through reconstruction of the aerodynamic coefficients of the vehicle throughout its trajectory. The primary dataset used in the aerodynamic reconstruction is the sensed acceleration measurements; thus, the aerodynamic and atmospheric uncertainties cannot be separated. Nevertheless, analysis of the estimated aerodynamic performance of the vehicle can be compared with predicted aerodynamic behavior, which in turn can lead to insight about

Poster Session 6B 208 crucial EDL events. One possible application of this performance analysis could be the during the parachute deployment phase, when one can compare the reconstructed aerodynamic coefficients with the predicted values. Additionally, Phoenix made atmospheric measurements shortly after it landed. Although these measurements do not exactly correspond with the timeline of the EDL events and measurements, these independent density, temperature and pressure measurements will be used to reconstruct the aerodynamic coefficients to see if the predicted behavior matched the estimated values. Since the atmospheric quantities are directly observable with this dataset, the uncertainties between the aerodynamics and atmospheric parameters can be separated. The methodology and tools used to generate the results in this paper were created during the development of an EKF- based reconstruction tool by the Space Systems Design Laboratory at the Georgia Institute of Technology. This EKF tool has been used in the past to reconstruct the trajectory of Mars vehicles, such as Pathfinder and the Mars Exploration Rovers, and has been augmented to reconstruct the trajectory and atmospheric profile of the 2012 Mars Science Laboratory. This tool development effort has been supported by a NASA Research Announcement (NRA) award.


VERTICAL STRUCTURE AND WIND SHEAR IN A SIMULATED TRITON ATMOSPHERE

Charles Miller(1), Nancy J. Chanover(1), James R. Murphy(1)

(1)New Mexico State University, Department of Astronomy, P. O. Box 30001, MSC 4500, Las Cruces, New Mexico 88003-8001 Email: [email protected], [email protected],

[email protected]

ABSTRACT Together, Neptune and its satellite Triton provide a unique opportunity for an automated orbiter to study both an ice giant planet and a possible captured Kuiper Belt Object, neither of which has been observed for long duration by an orbiting spacecraft. A Neptune orbiter would likely use Neptune’s upper atmosphere for initial aerocapture and subsequent aerobraking [1]. However, Triton also possesses an N2 atmosphere dense enough for aerobraking deceleration maneuvers on the order of those performed by the Mars Global Surveyor and Mars Odyssey. The aerobraking corridor density target of 100 kg/km3 for these missions is met by Triton’s atmosphere at an estimated height of approximately 100 km [2]. In addition to saving fuel, aerobraking provides a scientific benefit in that it directly provides information about density variations in the aerobraking corridor through analysis of changing deceleration rates as the spacecraft encounters the upper atmosphere. Planning an aerobraking trajectory in Triton’s atmosphere requires a prediction of density variation as a function of time, height, and latitude. Three-dimensional dynamic General Circulation Model (GCM) simulations provide a prediction of geographically and temporally varying atmospheric temperatures, which in turn determine changes in atmospheric scale height. Atmospheric density variations as a function of height can be derived from these globally varying scale heights. GCM simulations also provide predictions of wind speed and direction that must be considered in design of potential landing probes. These simulations are commonly used to provide the predictions needed for planning and implementing atmospheric insertion trajectories of Mars orbiters and landers. To model Triton’s atmosphere, we used a modified version of the NASA Ames Mars General Circulation Model, version 2.0. The Ames GCM incorporates several physical processes critical to modeling the atmosphere of Triton, including condensation and sublimation of the main atmospheric constituent gas as well as subsurface storage and release of heat. We altered the Ames GCM to simulate conditions found on Triton. These alterations included changing the size, rotation rate, orbital inclination, surface gravity, and distance to the Sun of the parent body to model the appropriate insolation over time. We also changed the gas properties from those of a CO2 atmosphere in the original Ames GCM to those of an N2 atmosphere, including appropriate values for latent heat, specific

Poster Session 6B 210 heat, and the proper vapor pressure-temperature relationship for N2 frosts. We chose initial albedo and emissivity values for the surface substrate and N2 frost from published values based on global thermal simulations of Triton [3]. Our simulations did not include atmospheric radiative heat transfer, but did include conduction, convection, and surface-boundary layer heating. The sub-solar latitude for our simulations was 48° S, similar to that at the Voyager 2 fly-by, which resulted in a continuously illuminated southern hemisphere and a dark northern pole. Triton is known to possess a thin N2 atmosphere with a surface pressure in the range of 10 - 20 microbars that is in vapor pressure equilibrium with N2 surface frosts [4]. We started with a global covering of 10 cm of N2 frost and an initial surface pressure of 10 microbars, and ran simulations covering 340 Triton days, or 2000 Earth days, to allow for the establishment of an equilibrium condensation flow. These simulations covered a 5 Earth year period just prior to Triton’s southern summer solstice. We experimented with changing N2 frost albedo and emissivity values and established a stable average atmospheric pressure of 18 microbars over a period of 330 Triton days. Our simulations produced a prograde polar jet that formed 3 km above the surface around both the subliming south pole and the condensing north pole. We attribute these flows to thermal winds caused by warmer air temperatures above the equator than above either pole. These thermal winds produced a wind shear most prominent at latitudes 65° S and 65° N. The prograde jet wind speed varied from 20 – 35 m/s at an altitude of 10 km from 100 to 300 Triton days. Lower altitude winds in the southern hemisphere flowed northward from the subliming pole and were deflected in a retrograde direction due to Coriolis effects, with a wind speed on the order of 5 m/s. The direction of these winds is consistent with the direction of ground plume streaks and active plumes imaged by Voyager 2 in August 1989 [5]. We are in the process of modifying the model to study global circulation patterns caused by variable surface ice patterns and a latitudinally changing frost albedo. We will also investigate the differences in wind patterns at several additional regions in Triton’s orbit including at equinox and at a sub-solar point of 25-30° S, which corresponds to a period in Triton’s orbit between 2025 and 2030, when a Neptune orbiter might conceivably arrive at Neptune. This study was funded by a NASA Earth and Space Science Fellowship through grant number NNX09AQ96H. References [1] Lockwood, Mary 2004. Neptune Aerocapture Systems Analysis. AIAA Atmospheric Flight Mechanics Conference and Exhibit. [2] Tolson, R. H. et al., 2005. Application of Accelerometer Data to Mars Odyssey Aerobraking and Atmospheric Modeling. J. Spacecraft Rockets. 42, 435-443. [3] Stanberry, J. A. et. al., 1990. Zonally averaged thermal balance and stability models for nitrogen polar caps on Triton. J. Geophys. Res. 17, 1773-1776. [4] Smith, B. A. et. al., 1989. Voyager 2 at Neptune - Imaging science results. Science. 246, 1422-1449. [5] Hansen, C. J. et. al., 1990. Surface and airborne evidence for plumes and winds on Triton. Science. 250, 421-424.

Poster Session 7A - Advances in TPS Technology for Planetary Probe Design


DEVELOPMENT OF A THERMAL PROTECTION SYSTEM MASS ESTIMATING RELATIONSHIP BASED

ON FIAT PREDICTIONS

S. Sepka1, J. O. Arnold2, E. Venkatapathy3 and K. Trumble4

ABSTRACT

Of major interest in the design of a thermal protective system (TPS) for entry into Earth’s atmosphere is the space ship’s required amount of heat shield material for safe passage. Presented here is the development of mass-estimating-relationships (MERs) used to predict the amount of TPS material to keep its back face temperature of the ablator below 250°C, which is considered to but the typical maximum temperature an epoxy can withstand when holding the TPS to its aeroshell. The MERs were developed based upon FIATi predictions at the stagnation point for a range of possible flight paths that resulted in the creation of 840 trajectories using DPLRii. Variables considered in this MER correlations included entry flight path angle, entry velocity, ballistic coefficient, heat load, peak heat flux, and maximum surface pressure. It will be shown that entry flight path angle and heat load had the greatest sensitivity to required thickness. Multiple MERs were developed using the PICAiii and one MER was developed for Carbon Phenolic [3]. Accuracy of the PICA MERs to FIAT prediction were within 13% at one standard deviation (SD), while the Carbon Phenolic MER had an accuracy of 7% at one SD. How the MERs were created, their modeling assumptions and limitations, and the applicability of these MERs will be discussed. 1 Senior. Research Scientist, ERC Corporation, Thermal Protection Materials and Systems Branch, NASA Ames Research Center, MS-234-1, Moffett Field, CA, 94035 (Tel. 650-604-3833). 2 NASA Ames IPA, Entry Systems and Technologies Division (Code TS), MS 229-3 3 Chief Technologist, Entry Systems and Technologies Division (Code TS), MS 229-3 4 Research Scientist, NASA Ames Research Center, Reacting Flow Environments Branch (Code TSA), MS 230-2, References i Chen, Y.-K., and Milos, F. S., “Fully Implicit Ablation and Thermal Analysis Program (FIAT),” Journal of Spacecraft and Rockets,Vol. 36, No. 3, pp 475-483, May–June 1999 ii Wright, M.W., White, T., and Mangini, N., Data Parallel Line Relaxation (DPLR) Code User Manual Acadia Version 4.01.1, NASA/TM-2009-215388, October 2009. iii Tran, H., Johnson, C, Rasky, D., Hui, F., Chen, Y.-K., and Hsu, M., “Phenolic Impregnated Carbon Ablators (PICA) for Discovery Class Missions,” AIAA Paper 96-1911, June 1996.


RASTAS SPEAR : RADIATION-SHAPES-THERMAL PROTECTION INVESTIGATIONS FOR HIGH SPEED

EARTH RE-ENTRY

J-M Bouilly1, A. Pisseloup1, O. Chazot2, G. Vekinis3, A. Bourgoing4, B. Chanetz5, O. Sladek6

1EADS Astrium Space Transportation - BP 20011, 33 165 Saint-Médard_en-Jalles

Cedex, France e-mail : [email protected] 2Von Karman Institute, Aeronautics & Aerospace Dept, Ch De Waterloo 72, 1640 Rh-St-Genese, Belgium

3Institute of Materials Science, NCSR "Demokritos", 15310, Aghia Paraskevi Attikis, Greece 4EADS Astrium Space Transportation, Route de Verneuil, 78133 Les Mureaux CEDEX, France 5ONERA, 8 rue des Vertugadins, 92 190 Meudon, France 6Kybertec,

s.r.o., Tovarni 1112, 537 01 Chrudim, Czec Rep.

ABSTRACT

An important step for Space Exploration activities and for a more accurate knowledge of the Earth, universe and environment is to develop the capability to send vehicles into space, which collect and return to Earth samples from solar system bodies. To return these samples, any mission will end by high-speed re-entry in Earth’s atmosphere. This requires strong technological bases and a good understanding of the environment encountered during the Earth re-entry. Investment in high speed re-entry technology development is thus appropriate today to enable future Exploration missions such as Mars Sample Return. Rastas Spear project started in September 2010, with the main objective to increase Europe’s knowledge in high speed re-entry vehicle technology to allow for planetary exploration missions in the coming decades. The research leading to these results has received funding from the European Union Seventh Framework Programme (FP7/2007-2013) under grant agreement n° 241992. The project’s main objective can be derived in sub-objectives as follows: " OBJ1: To better understand phenomena during high speed re-entry enabling more precise capsule sizing and reduced margins. " OBJ2: To identify the ground facility needs for simulation " OBJ3: To master heat shield manufacturing techniques and demonstrate heat shield capabilities. " OBJ4: To master damping at ground impact and flight mechanics and thus ensure a safe return of the samples. This study is carried out by a consortium of European companies and institutes : VKI (B), Kybertec (Cz), Demokritos (Gr), IoA (Pl), CIRA (I), CFS (CH), MSU (Ru), CNRS and ONERA (F), and coordinated by Astrium (F). The aim of this paper is to present the

Poster Session 7B 214 organisation, objectives and main actions proposed on RASTAS SPEAR project, to enlarge the basic capabilities on some specific topics such as: " Aeroshape stability " High speed aerothermal environment " Sub-system / equipment : Thermal protection, Crushable material


RESIN IMPREGNATED CARBON ABLATOR (RICA): A NEW THERMAL PROTECTION SYSTEM MATERIAL FOR HIGH-SPEED PLANETARY ENTRY VEHICLES

Jaime Esper (1), Hans-Peter Roeser (2), Georg Herdrich (2)

(1) NASA Goddard Space Flight Center, Greenbelt MD 20771, USA; Email: [email protected] (2) Institute of Space Systems, University of Stuttgart,

Pfaffenwaldring 31, 70569 Stuttgart, Germany; Email: [email protected]

ABSTRACT A new high-temperature carbon/Phenolic ablative Thermal Protection System (TPS) material was manufactured and tested at the Institute of Space Systems of the University of Stuttgart in Germany in the summer and fall of 2010. Designed for high velocity, hyperbolic entry speeds, the Resin Impregnated Carbon Ablator (RICA) material was developed as part of the principal author’s doctoral research focused on a planetary entry probe into Saturn’s moon Titan, and is the result of a collaborative endeavor under the auspices of NASA GSFC's Student Fellowship Program and the Institute of Space Systems (Institut für Raumfahrtsysteme - IRS). Heritage hyperbolic-entry speed carbon/Phenolic ablators rely on materials that are no longer in production (i.e., Galileo, Pioneer Venus); hence the development of alternatives such as RICA is necessary for future NASA planetary entry and Earth re-entry missions. RICA's performance was tested both in Methane to simulate Titan’s atmospheric composition, and in air. Several variants of the material were exposed to heat fluxes ranging from 1.4 to 14MW/m2, and durations from 7.5 minutes to 22 seconds at the Institute’s plasma wind tunnel (PWK1). The TPS' integrity was well preserved in most cases, and results show great promise for the intended applications. The flight environment for the Titan Aerobot Balloon System (TABS) entry heat shield is defined in Figure 1, which gives the convective and radiative heating rates as well as the vehicle speed from entry interface at 1000 km altitude to drogue deployment about 136 seconds later. The radiative heating rate correlation value (1903 W/cm2), or maximum computed rate, together with the convective heating rate (996 W/cm2) were used as the boundary condition for testing the manufactured TPS material. Figure 1 also provides a summary of key thermal requirements levied on the entry vehicle and TPS material.


There are several major elements involved in the creation of a successful ablative TPS material: the choice of fabric and resin formulation is only the beginning. The actual processing involved in manufacturing involves a careful choice of temperature, pressure, and time. This manufacturing process must result in a material that survives heat loads with no de-lamination or spallation. Several techniques have been developed to achieve this robustness. Variants of RICA’s material showed no delamination or spallation at intended heat flux levels, and their potential thermal protection capability was demonstrated. Three resin formulations were tested in two separate samples each manufactured under slightly different conditions. A total of six samples were eventually chosen for test at the IRS PWK1. Boundary conditions, test set-up, and results will be discussed in this paper.

Esper et. al.

RESIN IMPREGNATED CARBON ABLATOR (RICA): A NEW THERMAL PROTECTION SYSTEM MATERIAL FOR HIGH-SPEED PLANETARY ENTRY VEHICLES

Jaime Esper (1), Hans-Peter Roeser (2), Georg Herdrich (2)

(1) NASA Goddard Space Flight Center, Greenbelt MD 20771, USA; Email: [email protected] (2) Institute of Space Systems, University of Stuttgart, Pfaffenwaldring 31, 70569 Stuttgart, Germany; Email:

[email protected]

A new high-temperature carbon/Phenolic ablative Thermal Protection System (TPS) material was manufactured and tested at the Institute of Space Systems of the University of Stuttgart in Germany in the summer and fall of 2010. Designed for high velocity, hyperbolic entry speeds, the Resin Impregnated Carbon Ablator (RICA) material was developed as part of the principal author’s doctoral research focused on a planetary entry probe into Saturn’s moon Titan, and is the result of a collaborative endeavor under the auspices of NASA GSFC's Student Fellowship Program and the Institute of Space Systems (Institut für Raumfahrtsysteme - IRS). Heritage hyperbolic-entry speed carbon/Phenolic ablators rely on materials that are no longer in production (i.e., Galileo, Pioneer Venus); hence the development of alternatives such as RICA is necessary for future NASA planetary entry and Earth re-entry missions. RICA's performance was tested both in Methane to simulate Titan’s atmospheric composition, and in air. Several variants of the material were exposed to heat fluxes ranging from 1.4 to 14MW/m2, and durations from 7.5 minutes to 22 seconds at the Institute’s plasma wind tunnel (PWK1). The TPS' integrity was well preserved in most cases, and results show great promise for the intended applications.

The flight environment for the Titan Aerobot Balloon System (TABS) entry heat shield is defined in Figure 1, which gives the convective and radiative heating rates as well as the vehicle speed from entry interface at 1000 km altitude to drogue deployment about 136 seconds later. The radiative heating rate correlation value (1903 W/cm2), or maximum computed rate, together with the convective heating rate (996 W/cm2) were used as the boundary condition for testing the manufactured TPS material. Figure 1 also provides a summary of key thermal requirements levied on the entry vehicle and TPS material.

Peak specific heat input (enthalpy) 1.10 x 107 J/kg Stagnation point Integrated Convective Heat Flux 1.91 x 104 J/cm2 Stagnation point Integrated Radiative Heat Flux 4.42 x 104 J/cm2 Total Stagnation Point Integrated Heat Flux 6.33 x 104 J/cm2 Maximum Heat Shield Thickness (stagnation point) 1.648 cm

Figure 1: Flight Conditions and TPS Requirements for TABS

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PERFORMANCE CHARACTERIZATION, SENSITIVITY AND COMPARISON OF A DUAL LAYER THERMAL

PROTECTION SYSTEM

Cole D. Kazemba(1), Mary Kathleen McGuire(2), Austin Howard(3), Ian G. Clark(4), Robert D. Braun(5)

(1) Daniel Guggenheim School of Aerospace Engineering, Georgia Institute of

Technology, 270 Ferst Drive NW, Atlanta, GA 30332, USA, Email: [email protected]

(2) Systems Analysis Branch, NASA Ames Research Center, Moffett Field, CA 94035-0001 (3) Neerim Corporation, 2551 Casey Ave. Ste. B, Mountain View, CA 94086

42) Daniel Guggenheim School of Aerospace Engineering, Georgia Institute of Technology, 270 Ferst Drive NW, Atlanta, GA 30332, USA, Email: [email protected]

(5) Daniel Guggenheim School of Aerospace Engineering, Georgia Institute of Technology, 270 Ferst Drive NW, Atlanta, GA 30332, USA, Email:

[email protected]

ABSTRACT With the goal of landing high-mass cargo or crewed missions on Mars, NASA has been developing new thermal protection technologies with enhanced capability and reduced mass compared to traditional approaches. Two examples of new thermal protection system (TPS) concepts are dual layer and flexible TPS. Each of these systems introduces unique challenges along with potential performance enhancements. Traditional monolithic ablative TPS, which have been flown on every Mars robotic mission to date, use a single layer of ablative material. The new dual layer TPS concepts utilize an insulating layer of material beneath an ablative layer to save mass. A study was conducted on the dual layer system to identify sensitivities in performance to uncertainties in material properties and aerothermal environments. A performance metric which is independent of the system construction was developed in order to directly compare the abilities and benefits of the traditional, dual layer and eventually, flexible systems. Using a custom MATLAB code enveloping the Fully Implicit Ablation and Thermal Response Program (FIAT), the required TPS areal mass was calculated for several different parametric scenarios. Overall TPS areal mass was found to be most sensitive to the allowable temperature at the ablator/insulator interface and aerothermal heat transfer augmentation (attributed here to material surface roughness). From these preliminary results it was found that the dual layer TPS construction investigated could produce improvements over a traditional TPS in the specified performance metric between 14-36% (depending on the flight environments and total integrated heat load expected) with nominal material properties.


EDL HEATSHIELD EXPERIMENTS WITH DUAL-LAYER ABLATORS, ADVANCED MATERIALS AND VARIABLE

HONEYCOMBS

Jennifer N. Congdon

ARA Ablatives Laboratory (ABL) Centennial, Colorado 80112

ABSTRACT Today’s challenge is to make ablative heatshield systems lighter and more efficient for thermal protection of the large EDL vehicles baselined for manned exploration of Mars. This paper discusses interim results from a three-year ABL program to improve upon heatshield systems already available. The project is focused on three elements: 1) fabricate and test dual-layer ablator systems with a higher-density, more robust top layer over a lower-density, more insulative sublayer of the same chemistry; 2) develop and investigate new ablator constituents such as silicon-carbide microballoons and fibers to replace less-durable fillers currently in use; and 3) produce, test and evaluate honeycombs with a wide range of cell size to better understand the dependence of ablator performance on reinforcement configurations. Primary ablator performance testing has consisted of: 1) arc-jet iso-q stagnation testing using the NASA/ARC IHF tunnel; 2) arc-jet aeroshear testing with a swept-cylinder design using the IHF tunnel, and 3) concentrated solar radiation testing using the Sandia Labs Solar Tower facility. The main focus of this presentation will be summarizing test results collected to date and interpretations of sample performance from the wide array of experimental ablator samples.


LOW DENSITY FLEXIBLE CARBON PHENOLIC ABLATORS

Mairead Stackpoole1, Jeremy Thornton1, Wendy Fan1 and Parul Agrawal1, Evan Doxtad2, Robin Beck3

ERC Inc., NASA Ames Research Center, NASA Ames Research Center3, NASA Education

Associates Program2, NASA Ames Research Center

ABSTRACT Phenolic Impregnated Carbon Ablator (PICA) was the enabling TPS material for the Stardust mission where it was used as a single piece heatshield. PICA has the advantages of low density (~0.27g/cm3) coupled with efficient ablative capability at high heat fluxes. Under the Orion program, PICA was also shown to be capable of both ISS and lunar return missions; however some unresolved issues remain for its application in a tiled configuration for the Orion-specific design. In particular, the problem of developing an appropriate gap filler resulted in the Orion program selecting AVCOAT as the primary heatshield material over PICA. We are currently looking at alternative architectures to yield flexible and more conformal carbon phenolic materials with comparable densities to PICA that will address some of the design issues faced in the application of a tiled PICA heat shield. These new materials are viable TPS candidates for upcoming NASA missions and as material candidates for private sector Commercial Orbital Transportation Services (COTS). This presentation will discuss flexible alternatives to PICA and include preliminary mechanical and thermal properties as well as arc jet and LHMEL screening test results.


ROTATING ARC JET TEST MODEL: TIME-ACCURATE TRAJECTORY HEAT FLUX REPLICATION IN A

GROUND TEST ENVIRONMENT

Bernard Laub and Jay Grinstead1, Artem Dyakonov2, Ethiraj Venkatapathy1

NASA Ames Research Center1, NASA Langley Research Center2

ABSTRACT

Though arc jet testing has been the proven method employed for development testing and certification of TPS and TPS instrumentation, the operational aspects of arc jets limit testing to selected, but constant, conditions. Flight, on the other hand, produces time- varying entry conditions in which the heat flux increases, peaks, and recedes as a vehicle descends through an atmosphere. As a result, we are unable to “test as we fly.” Attempts to replicate the time-dependent aerothermal environment of atmospheric entry by varying the arc jet facility operating conditions during a test have proven to be difficult, expensive, and only partially successful. A promising alternative is to rotate the test model exposed to a constant-condition arc jet flow to yield a time-varying test condition at a point on a test article (Fig. 1). The model shape and rotation rate can be engineered so that the heat flux at a point on the model replicates the predicted profile for a particular point on a flight vehicle. This simple concept will enable, for example, calibration of the TPS sensors on the Mars Science Laboratory (MSL) aeroshell for anticipated flight environments.


During the test, the model is rotated such that the test model’s sensor will sweep through points of varying heat flux: near zero when directed away from the flow, and the maximum when the sensor is rotated to the stagnation point. Since the flight profile spans from zero to a maximum and back to zero, the angular direction and instantaneous rate at which the model is rotated will be programmed to realize a time-accurate heat flux profile that maps to the predicted profile for a chosen location on the flight vehicle. The result is a “test-as-you- fly” heat flux condition at the sensor location on the test model, yet requires no change to the facility operating condition. Although the surface pressure at the sensor location cannot follow the flight profile in tandem with the heat flux, the material response will not be significantly affected by small differences in pressure. The rotation of the cylinder will be accomplished with a programmable hydraulic

Rotating arc jet test model: Time-accurate trajectory heat flux replication in a ground test environment Bernard Laub and Jay Grinstead NASA Ames Research Center Moffett Field, CA 94035 Artem Dyakonov NASA Langley Research Center Hampton, VA 23681 Ethiraj Venkatapathy NASA Ames Research Center Moffett Field, CA 94035

Though arc jet testing has been the proven method employed for development testing and

certification of TPS and TPS instrumentation, the operational aspects of arc jets limit

testing to selected, but constant, conditions. Flight, on the other hand, produces time-

varying entry conditions in which the heat flux increases, peaks, and recedes as a vehicle

descends through an atmosphere. As a result, we are unable to “test as we fly.” Attempts

to replicate the time-dependent aerothermal environment of atmospheric entry by varying

the arc jet facility operating conditions during a test have proven to be difficult,

expensive, and only partially successful. A promising alternative is to rotate the test

model exposed to a constant-condition arc jet flow to yield a time-varying test condition

at a point on a test article (Fig. 1). The model shape and rotation rate can be engineered

so that the heat flux at a point on the model replicates the predicted profile for a particular

point on a flight vehicle. This simple concept will enable, for example, calibration of the

TPS sensors on the Mars Science Laboratory (MSL) aeroshell for anticipated flight

environments.

Figure 1: a) Schematic of rotating arc jet model concept. Embedded sensor encounters varying heat

fluxes as model is rotated. b) Trajectory heat flux profile and correlation with rotating arc jet model

positions.

During the test, the model is rotated

such that the test model’s sensor

will sweep through points of

varying heat flux: near zero when

directed away from the flow, and

the maximum when the sensor is

rotated to the stagnation point. Since

the flight profile spans from zero to

a maximum and back to zero, the

angular direction and instantaneous

rate at which the model is rotated

will be programmed to realize a

time-accurate heat flux profile that

maps to the predicted profile for a

chosen location on the flight

vehicle. The result is a “test-as-you-

fly” heat flux condition at the sensor

location on the test model, yet requires no change to the facility operating condition.

Although the surface pressure at the sensor location cannot follow the flight profile in

tandem with the heat flux, the material response will not be significantly affected by

small differences in pressure.

The rotation of the cylinder will be accomplished with a programmable hydraulic

actuator and position encoder. The transmission mechanism and encoder will be designed

to interface with the model support arm and accommodate the sensor’s instrumentation

wiring. Figure 2 shows a design concept for a cylindrical model shape.

This approach will be applied first to validation of sensor performance for the MSL Entry

Descent and Landing Instrumentation (MEDLI). We will present high fidelity rotating arc

jet model simulations and analyses of test protocols to realize time-accurate correlation to

MEDLI sensor points.

Figure 2: Rotating arc jet test model concept.

Hydraulic rotary actuator is encoded to enable

precise positional control for time-accurate heat flux

profile replication.

Poster Session 7B 222 actuator and position encoder. The transmission mechanism and encoder will be designed to interface with the model support arm and accommodate the sensor’s instrumentation wiring. Figure 2 shows a design concept for a cylindrical model shape. This approach will be applied first to validation of sensor performance for the MSL Entry Descent and Landing Instrumentation (MEDLI). We will present high fidelity rotating arc jet model simulations and analyses of test protocols to realize time-accurate correlation to MEDLI sensor points.


ADVANCED RIGID ABLATIVE TPS

Matt Gasch

NASA Ames Research Center

ABSTRACT Heritage ablative TPS materials using Viking or Pathfinder era materials are at or near their performance limits and will be inadequate for future missions. Significant advances in TPS materials technology are needed in order to enable any subsequent human exploration missions. This poster summarizes some recent progress at NASA in developing families of advanced rigid/conformable and flexible ablators that could potentially be used for thermal protection in planetary entry missions. In particular the effort focuses technologies required to land heavy (~40 metric ton) masses on Mars to facilitate future exploration plans.


MODELING OF THE MATERIAL RESPONSE OF THERMAL PROTECTION SYSTEMS IN HYPERSONIC

FLOWS

Jonathan Wiebenga(1), Iain D. Boyd(2), Alexandre Martin(2)

(1)Department of Aerospace Engineering, University of Michigan, 1320 Beal Ave. Ann Arbor, MI 48109, United States, Email: [email protected] (2)Department of Aerospace

Engineering, University of Michigan, 1320 Beal Ave. Ann Arbor, MI 48109, United States, Email: [email protected]

(3)Department of Mechanical Engineering, University of Kentucky, 261 Ralph G. Anderson Building Lexington, KY 40506, United States, Email:

[email protected]

ABSTRACT A one dimensional material response model called MOPAR has been developed to study ablation processes on hypersonic vehicles. MOPAR uses the Control Volume Finite-Element Method (CVFEM) to model the inner decomposition of an ablator, pyrolysis gas behavior, and surface ablation with wall recession. The solid and gas phase mass conservation equations are solved along with the total energy (solid and gas) conservation equation and the momentum conservation equation, which is pore- averaged to Forchheimer's Law. MOPAR has also been strongly coupled to LeMANS, a hypersonic CFD code, in order to simulate coupled flow field and material response problems. To demonstrate the coupling methodology, two test-cases are considered: the IRV-2 vehicle and the Passive Nosetip Technology (PANT) program. The IRV-2 results are compared with other published numerical results and show good agreement, and the PANT results are compared with experimental data and published numerical results.

Documents

IPPW8 Abstract Book June 1