DOCSLIB.ORG

1 Design of an Integral Thermal Protection System

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
1 Design of an Integral Thermal Protection System DESIGN OF AN INTEGRAL THERMAL PROTECTION SYSTEM FOR FUTURE SPACE VEHICLES By SATISH KUMAR BAPANAPALLI A DISSERTATION PRESENTED TO THE GRADUATE SCHOOL OF THE UNIVERSITY OF FLORIDA IN PARTIAL FULFILLMENT OF THE REQUIREMENTS FOR THE DEGREE OF DOCTOR OF PHILOSOPHY UNIVERSITY OF FLORIDA 2007 1 © 2007 Satish Kumar Bapanapalli 2 To my loving wife Debamitra, my parents Nagasurya and Adinarayana Bapanapalli, brother Gopi Krishna and sister Lavanya 3 ACKNOWLEDGMENTS I would like to express my sincere gratitude to my advisor and mentor Dr. Bhavani Sankar for his constant support (financial and otherwise) and motivation throughout my PhD studies. He allowed me to work with freedom, was always supportive of my ideas and provided constant motivation for my research work, which helped me grow into a mature and confident researcher under his tutelage. I also thank my committee co-chair Dr. Rafi Haftka for his invaluable inputs and guidance, which have been instrumental for my research work. I am also grateful to him for getting my interest into the field of Structural Optimization, which I hope would be a huge part of all my future research endeavors. I am also thankful to Dr. Max Blosser (NASA Langley) for his crucial inputs in my research work, which kept us on track with the expectations of NASA. I sincerely thank my dissertation committee members Dr. Ashok Kumar and Dr. Gary Consolazio for evaluating my research work and my candidature for the PhD degree. I also would like to acknowledge Dr. Peter Ifju and Dr. Nam-Ho Kim for their useful inputs and comments. I am thankful to NASA, for their financial support through the CUIP Project, and the program manager Claudia Meyer. Thanks also go to all the past and current members of Center of Advanced Composites: Ryan, Jongyoon, Huadong, Oscar, Jianlong, Thi, Sujith and Ben, and other graduate students in the department: Christian, Tushar, Vijay, and Palani. 4 TABLE OF CONTENTS page ACKNOWLEDGMENTS ...............................................................................................................4 LIST OF TABLES...........................................................................................................................7 LIST OF FIGURES .........................................................................................................................9 ABSTRACT...................................................................................................................................11 CHAPTER 1 INTRODUCTION AND OBJECTIVES .............................................................................13 1.1 Introduction.......................................................................................................................13 1.2 Requirements of a Thermal Protection System ................................................................16 1.3 Objectives .........................................................................................................................19 1.4 Approach...........................................................................................................................20 1.4.1 The Optimization Problem .....................................................................................20 1.4.2 Geometry and Design Variables.............................................................................21 1.4.3 Analysis: Finite Element Methods .........................................................................21 1.4.4 Loads and Boundary Conditions ............................................................................22 1.4.5 Formulation of Constraints.....................................................................................22 1.4.6 Optimized Designs .................................................................................................23 2 BACKGROUND..................................................................................................................24 2.1 Approaches to Thermal Protection ...................................................................................24 2.1.1 Active TPS..............................................................................................................24 2.1.2 Semi-passive TPS...................................................................................................25 2.1.3 Passive TPS ............................................................................................................25 2.2 TPS and NASA.................................................................................................................27 2.3 Need for a load-bearing TPS: ITPS..................................................................................36 2.4 ITPS Design Challenges...................................................................................................38 2.5 Choice of Constraints .......................................................................................................41 2.6 Background on Corrugated-Core and Truss-Core Sandwich Structures..........................43 2.7 Background on Multi-Disciplinary Optimization and Response Surface Approximation Techniques.................................................................................................45 3 FINITE ELEMENT MODELS AND ANALYSES.............................................................49 3.1 Finite Element Analysis for Heat Transfer.......................................................................49 3.1.1 Incident Heat Flux and Radiation Equilibrium Temperature.................................49 3.1.2 Loads, Boundary Conditions and Assumptions .....................................................52 3.1.3 One-Dimensional and Two-Dimensional FE Models ............................................55 3.1.4 Temperature vs. Reentry Time and Temperature Distribution ..............................58 5 3.1.5 Obtaining Temperature Data from the FE Analysis...............................................61 3.2 Finite Element Buckling Analysis....................................................................................62 3.3 Stress and Deflection Analysis .........................................................................................68 4 RESPONSE SURFACE APPROXIMATIONS AND OPTIMIZATION PROCEDURE ......................................................................................................................69 4.1 Response Surface Approximations...................................................................................69 4.1.1 Response Surface Approximations for Maximum Bottom Face Sheet Temperature .................................................................................................................70 4.1.2 Response Surface Approximations for Buckling ...................................................70 4.1.3 Response Surface Approximations for Stress and Deflection................................71 4.2 Procedure for Generation of Response Surfaces Approximations ...................................71 4.3 Optimization Procedure....................................................................................................75 5 ITPS DESIGNS....................................................................................................................76 5.1 Selection of Loads, Boundary Conditions and Other Input Parameters...........................76 5.2 Corrugated-Core Designs .................................................................................................80 5.2.1 Accuracy of Response Surface Approximations....................................................83 5.2.2 Optimized Corrugated-Core Panel Designs ...........................................................86 5.2.3 Buckling Eigen Values, Deflections and Stresses at Different Reentry Times......90 5.2.4 Optimized Designs with Changed Boundary Conditions.......................................97 5.3 Truss-Core Structures for ITPS ........................................................................................98 6 EFFECT OF INPUT PARAMETERS ON THE ITPS DESIGN.......................................104 6.1 Sensitivity of ITPS Designs to Heat Transfer Parameters..............................................105 6.1.1 Changing the Emissivity of the Top Surface of ITPS ..........................................105 6.1.2 Allowing Heat Loss from the Bottom Face Sheet................................................106 6.1.3 Increasing the initial temperature of the structure................................................107 6.2 Sensitivity of ITPS Designs to Loads and Boundary Conditions...................................109 6.2.1 Effect of Boundary Conditions.............................................................................109 6.2.2 Increasing the Pressure Load on the Top Surface ................................................109 6.2.3 Increasing the In-Plane Load................................................................................110 7 CONCLUSIONS AND FUTURE WORK ........................................................................111 7.1 Conclusions.....................................................................................................................111 7.2 Future Work....................................................................................................................113 APPENDIX: MATERIAL PROPERTIES USED FOR THE RESEARCH WORK ..................116 LIST OF REFERENCES.............................................................................................................120
Load more

Recommended publications
  • Aerothermodynamic Analysis of a Mars Sample Return Earth-Entry Vehicle" (2018)
    Old Dominion University ODU Digital Commons Mechanical & Aerospace Engineering Theses & Dissertations Mechanical & Aerospace Engineering Summer 2018 Aerothermodynamic Analysis of a Mars Sample Return Earth- Entry Vehicle Daniel A. Boyd Old Dominion University, [email protected] Follow this and additional works at: https://digitalcommons.odu.edu/mae_etds Part of the Aerodynamics and Fluid Mechanics Commons, Space Vehicles Commons, and the Thermodynamics Commons Recommended Citation Boyd, Daniel A.. "Aerothermodynamic Analysis of a Mars Sample Return Earth-Entry Vehicle" (2018). Master of Science (MS), Thesis, Mechanical & Aerospace Engineering, Old Dominion University, DOI: 10.25777/xhmz-ax21 https://digitalcommons.odu.edu/mae_etds/43 This Thesis is brought to you for free and open access by the Mechanical & Aerospace Engineering at ODU Digital Commons. It has been accepted for inclusion in Mechanical & Aerospace Engineering Theses & Dissertations by an authorized administrator of ODU Digital Commons. For more information, please contact [email protected]. AEROTHERMODYNAMIC ANALYSIS OF A MARS SAMPLE RETURN EARTH-ENTRY VEHICLE by Daniel A. Boyd B.S. May 2008, Virginia Military Institute M.A. August 2015, Webster University A Thesis Submitted to the Faculty of Old Dominion University in Partial Fulfillment of the Requirements for the Degree of MASTER OF SCIENCE AEROSPACE ENGINEERING OLD DOMINION UNIVERSITY August 2018 Approved by: __________________________ Robert L. Ash (Director) __________________________ Oktay Baysal (Member) __________________________ Jamshid A. Samareh (Member) __________________________ Shizhi Qian (Member) ABSTRACT AEROTHERMODYNAMIC ANALYSIS OF A MARS SAMPLE RETURN EARTH-ENTRY VEHICLE Daniel A. Boyd Old Dominion University, 2018 Director: Dr. Robert L. Ash Because of the severe quarantine constraints that must be imposed on any returned extraterrestrial samples, the Mars sample return Earth-entry vehicle must remain intact through sample recovery.
    [Show full text]
  • The Flight to Orbit
    There are numerous ways to get there— Around the Corner Th As then–US Space Command chief from rocket The Gen. Howell M. Estes III said to launch to space defense writers just before his re- tirement in August, “This is going maneuver to come along a lot quicker than we vehicles—and the think it is. ... We tend to think this stuff is way out there in the future, Air Force is but it’s right around the corner.” keeping its Flight The Air Force and NASA have divided the task of providing the options open. US government with a means of reliable, low-cost transportation to Earth orbit. The Air Force, with the largest immediate need, is heading to up the effort to revamp the Expend- to able Launch Vehicles now used to loft military and other government satellites. Called the Evolved ELV, this program is focused on derivatives of existing rockets. Competitors have been invited to redesign or value- engineer their proven boosters with new materials and technologies to provide reliable launch services at a he Air Force would like to go Orbit far lower price than today’s bench- back and forth to Earth orbit as bit mark of around $10,000 a pound to T O easily as it goes back and forth to Low Earth Orbit. The reasoning is 30,000 feet—routinely, reliably, and that an “evolved”—rather than an relatively cheaply. Such a capability all-new—vehicle will yield cost sav- goes hand in hand with being a true ings while reducing technical risk.
    [Show full text]
  • APOLLO EXPERIENCE REPORT - THERMAL PROTECTION SUBSYSTEM by Jumes E
    NASA TECHNICAL NOTE NASA TN D-7564 w= ro VI h d z c Q rn 4 z t APOLLO EXPERIENCE REPORT - THERMAL PROTECTION SUBSYSTEM by Jumes E. Puulosky und Leslie G, St. Leger Ly12d012 B. Johlzson Space Center Honst0~2, Texus 77058 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, 0. C. JANUARY 1974 ~--_. - .. 1. Report No. 2. Government Accession No. 3. Recipient's Catalog No. D-7564 4. Title and Subtitle 5. Report Date January 1974 APOLLOEXPERIENCEREPORT THERMAL PROTECTION SUBSYSTEM 6. Performing Organization Code I 7. Author(s) I 8. Performing Organization Report No. JSC S-383 James E. Pavlosky and Leslie G. St. Leger, JSC 10. Work Unit No. I 9. Performing Organization Name and Address 11. Contract or Grant No. Lyndon B. Johnson Space Center Houston, Texas 77058 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address 14. Sponsoring Agency Code National Aeronautics and SDace Administration Washington, D. C. 20546 1 15. Supplementary Notes The JSC Director waived the use of the International System of Units (SI) for this Apollo Experienc Report because, in his judgment, the use of SI units would impair the usefulness of th'e report or result in excessive cost. 16. Abstract The Apollo command module was the first manned spacecraft to be designed to enter the atmos- phere of the earth at lunar-return velocity, and the design of the thermal protection subsystem for the resulting entry environment presented a major technological challenge. Brief descrip- tions of the Apollo command module thermal design requirements and thermal protection con- figuration, and some highlights of the ground and flight testing used for design verification of the system are presented.
    [Show full text]
  • Preparation of Papers for AIAA Technical Conferences
    Thermal Protection Systems Technology Transfer from Apollo and Space Shuttle to the Orion Program Michael Stewart 1 Lockheed Martin, Kennedy Space Center, Florida, 32899 and William J. Koenig2 NASA Kennedy Space Center, Florida, 32899 and Richard F. Harris 3 NASA Kennedy Space Center, Florida, 32899 This paper describes how the Orion program is utilizing the Thermal Protection System (TPS) experience from the Apollo and Space Shuttle programs to reduce program risk and improve affordability to meet NASA’s future manned exploration missions. The Orion program successfully completed the Exploration Flight Test (EFT-1) mission in 2014 and is currently assembling, integrating, and testing the next spacecraft for the Exploration Mission (EM-1) to meet the flight test objectives of an unmanned orbital mission to the moon and return to earth in 2019. The Orion spacecraft production operations are located in the Neil Armstrong Operations and Checkout (O&C) facility at the Kennedy Space Center (KSC) providing an affordable and seamless delivery approach of vehicles directly to the launch site eliminating spacecraft transportation and additional checkout testing. Innovative vehicle design, manufacturing and test operations approaches are maturing and evolving with each Orion vehicle build to support the challenging NASA exploration mission requirements beyond Low Earth Orbit (LEO) while reducing program cost and schedule impacts. An example of Orion’s evolution is the incorporation of an improved heat shield design, assembly and testing approach to meet the higher re-entry velocities for a lunar return for the EM-1 mission. The EFT- 1 heat shield was based on the Apollo heat shield manufacturing processes and was assembled at a supplier location and then transported to KSC for final integration.
    [Show full text]
  • Tactical Missile Structures and Materials Technology
    WILLIAM C. CAYWOOD, ROBERT M. RIVELLO, and LOUIS B. WECKESSER TACTICAL MISSILE STRUCTURES AND MATERIALS TECHNOLOGY Significant advances have occurred in missile structures and materials technology since the late 1940's. Some of the advances were the result of improvements in design concepts and material prop­ erties; others were from the development of new materials and fabrication techniques. Rapid strides in computer technology and methods of structural analysis have also played an important role. This article discusses the structural design requirements for the next generation of missiles and reviews the technological advances that will make it possible to meet those requirements. INTRODUCTION Antenna housing The Applied Physics Laboratory's experience in Cowl assembly tactical missile structural design began with the devel­ opment of the rocket-powered Terrier and the ramjet-powered Talos in the late 1940's. The air­ frames for those missiles were designed with conven­ tional aircraft materials, using slide rules, mechanical desk calculators, handbooks, and rule-of-thumb pro­ cedures. To ensure adequate safety margins, conser­ Innerbody fairing vative analytical procedures were pursued, followed Sheet magnesium Innerbody by extensive environmental testing. Simple aluminum forward cone Cast magnesium and steel structures operating at temperatures below Figure 1 - Magnesium components in Talos represented 300°F for Terrier and 600°F for Talos made up the 13 % of the weight of the primary load-carrying structure major portion of the missiles. Terrier was con­ and 10 % of the gross launch weight. structed primarily from machined castings. Talos, shown in Fig. 1 with its central air duct configura­ tion, used sheet metal construction.
    [Show full text]
  • Delta Clipper a Path to the Future
    Delta Clipper A Path to the Future By Jason Moore & Ashraf Shaikh Executive Summary Although the Space Shuttle has well served its purpose for years, in order to revitalize and advance the American space program, a new space launch vehicle is needed. A prime candidate for the new manned launch vehicle is the DC-X. The DC-X isn’t a state-of-the-art rocket that would require millions of dollars of new development. The DC-X is a space launch vehicle that has already been tested and proven. Very little remains to be done in order to complete the process of establishing the DC-X as an operational vehicle. All that’s left is the building and final testing of a full-size DC-X, followed by manufacture and distribution. The Space Shuttle, as well as expendable rockets, is very expensive to build, maintain and launch. Costing approximately half a billion dollars for each flight, NASA can only afford to do a limited number of Space Shuttle missions. Also, the Shuttle is maintenance intensive, requiring hundreds of man-hours of maintenance after each flight. The DC-X, however, is very cheap to build, easy to maintain, and much cheaper to operate. If the DC-X was used as NASA’s vehicle of choice, NASA could afford to put more payloads into orbit, and manned space missions wouldn’t be the relative rarity they are now. Since not much remains in order to complete the DC-X, a new private organization dedicated solely to the DC-X would be the ideal choice for the company that would build it.
    [Show full text]
  • Heat Flux at a Point on the Vehicle (W/Cm2)
    Entry Aerothermodynamics • Review of basic fuid parameters • Heating rate parameters • Stagnation point heating • Heating on vehicle surfaces • Slides from 2012 NASA Termal and Fluids Analysis Workshop: https://tfaws.nasa.gov/TFAWS12/Proceedings/ Aerothermodynamics%20Course.pdf © 2016 David L. Akin - All rights reserved http://spacecraft.ssl.umd.edu U N I V E R S I T Y O F Entry Aerothermodynamics MARYLAND ENAE 791 - Launch and Entry Vehicle Design1 Background 2 • The kinetic energy of an entry vehicle is dissipated by transformation into thermal energy (heat) as the entry system decelerates • The magnitude of this thermal energy is so large that if all of this energy were transferred to the entry system it would be severely damaged and likely vaporize – Harvey Allen - the blunt body concept • Only a small fraction of this thermal energy is transferred to the entry system – The thermal transfer fraction is dependant on vehicle shape, size, aerodynamic regime and velocity – Near peak heating, 1% to 5% of the total thermal energy is transferred to the entry system – Example: at the peak heating point the freestream energy transfer for 1 3 2 2 Pathfinder was q Ý Ñ = 2 r V ~ 4 , 000 W/cm but only about 110 W/cm (2.7%) was actually transferred to the surface Example 3 E V2 Energy density: = + g h m 2 o V E/m Entry (km/s) (MJ/kg) MER 5.6 16 Note that: Apollo 11.4 66 Water boils @ 2.3 MJ/kg Carbon vaporizes @ 60.5 MJ/kg Mars 14.0 98 Return Galileo 47.4 1130 In each case goh is about 1% of total Side Note: What Can We Test? 4 Missions of Interest Live here Blunt Body Rationale 5 • Why is a blunt body used for planetary entry? – Slender body: low drag, highly maneuverable – Blunt body: high drag, not very maneuverable • Blunt bodies generate strong shock waves – Efficient energy dissipation.
    [Show full text]
  • Aerodynamic Heating of Inflatable Aeroshell in Orbital Reentry
    Title Aerodynamic heating of inflatable aeroshell in orbital reentry Author(s) Takahashi, Yusuke; Yamada, Kazuhiko Acta astronautica, 152, 437-448 Citation https://doi.org/10.1016/j.actaastro.2018.08.003 Issue Date 2018-11 Doc URL http://hdl.handle.net/2115/79647 © 2018. This manuscript version is made available under the CC-BY-NC-ND 4.0 license Rights http://creativecommons.org/licenses/by-nc-nd/4.0/ Rights(URL) http://creativecommons.org/licenses/by-nc-nd/4.0/ Type article (author version) File Information paper_titans_heatflux.pdf Instructions for use Hokkaido University Collection of Scholarly and Academic Papers : HUSCAP Aerodynamic Heating of Inflatable Aeroshell in Orbital Reentry Yusuke Takahashi1, Hokkaido University, Kita 13 Nishi 8, Kita-ku, Sapporo, Hokkaido 060-8628, Japan and Kazuhiko Yamada2 Japan Aerospace Exploration Agency, 3-1-1 Yoshinodai Chuo-ku, Sagamihara, Kanagawa 252-5210, Japan Keywords: Inflatable reentry vehicle, Aerodynamic heating, Hypersonic, Deformation, Coupled analysis Abstract The aerodynamic heating of an inflatable reentry vehicle, which is one of the innova- tive reentry technologies, was numerically investigated using a tightly coupled approach involving computational fluid dynamics and structure analysis. The fundamentals of a high-enthalpy flow around the inflatable reentry vehicle were clarified. It was found that the flow fields in the shock layer formed in front of the vehicle were strongly in a chemical nonequilibrium state owing to its low-ballistic coefficient trajectory. The heat flux tendencies on the surface of the vehicle were comprehensively investigated for various effects of the vehicle shape, surface catalysis, and turbulence via a parametric study of these parameters.
    [Show full text]
  • Columbia Accident Investigation Board
    COLUMBIA ACCIDENT INVESTIGATION BOARD Report Volume I August 2003 COLUMBIA ACCIDENT INVESTIGATION BOARD On the Front Cover This was the crew patch for STS-107. The central element of the patch was the microgravity symbol, µg, flowing into the rays of the Astronaut symbol. The orbital inclination was portrayed by the 39-degree angle of the Earthʼs horizon to the Astronaut symbol. The sunrise was representative of the numerous science experiments that were the dawn of a new era for continued microgravity research on the International Space Station and beyond. The breadth of science conduct- ed on this mission had widespread benefits to life on Earth and the continued exploration of space, illustrated by the Earth and stars. The constellation Columba (the dove) was chosen to symbolize peace on Earth and the Space Shuttle Columbia. In addition, the seven stars represent the STS-107 crew members, as well as honoring the original Mercury 7 astronauts who paved the way to make research in space possible. The Israeli flag represented the first person from that country to fly on the Space Shuttle. On the Back Cover This emblem memorializes the three U.S. human space flight accidents – Apollo 1, Challenger, and Columbia. The words across the top translate to: “To The Stars, Despite Adversity – Always Explore“ Limited First Printing, August 2003, by the Columbia Accident Investigation Board Subsequent Printing and Distribution by the National Aeronautics and Space Administration and the Government Printing Office Washington, D.C. 2 Report Volume I August 2003 COLUMBIA ACCIDENT INVESTIGATION BOARD IN MEMORIAM Rick D. Husband Commander William C.
    [Show full text]
  • Space Planes and Space Tourism: the Industry and the Regulation of Its Safety
    Space Planes and Space Tourism: The Industry and the Regulation of its Safety A Research Study Prepared by Dr. Joseph N. Pelton Director, Space & Advanced Communications Research Institute George Washington University George Washington University SACRI Research Study 1 Table of Contents Executive Summary…………………………………………………… p 4-14 1.0 Introduction…………………………………………………………………….. p 16-26 2.0 Methodology…………………………………………………………………….. p 26-28 3.0 Background and History……………………………………………………….. p 28-34 4.0 US Regulations and Government Programs………………………………….. p 34-35 4.1 NASA’s Legislative Mandate and the New Space Vision………….……. p 35-36 4.2 NASA Safety Practices in Comparison to the FAA……….…………….. p 36-37 4.3 New US Legislation to Regulate and Control Private Space Ventures… p 37 4.3.1 Status of Legislation and Pending FAA Draft Regulations……….. p 37-38 4.3.2 The New Role of Prizes in Space Development…………………….. p 38-40 4.3.3 Implications of Private Space Ventures…………………………….. p 41-42 4.4 International Efforts to Regulate Private Space Systems………………… p 42 4.4.1 International Association for the Advancement of Space Safety… p 42-43 4.4.2 The International Telecommunications Union (ITU)…………….. p 43-44 4.4.3 The Committee on the Peaceful Uses of Outer Space (COPUOS).. p 44 4.4.4 The European Aviation Safety Agency…………………………….. p 44-45 4.4.5 Review of International Treaties Involving Space………………… p 45 4.4.6 The ICAO -The Best Way Forward for International Regulation.. p 45-47 5.0 Key Efforts to Estimate the Size of a Private Space Tourism Business……… p 47 5.1.
    [Show full text]
  • Post-Flight Assessment of Avcoat Thermal Protection System for The
    Post-Flight Assessment of Avcoat Thermal Protection System for the Exploration Flight Test-1 Deepak Bose, Jose Santos, Erika Rodriguez, Milad Mahzari, Brian Remark, and Suman Muppidi NASA Ames Research Center, Moffett Field, CA NASA Johnson Space Center, Houston, TX On December 5, 2014 NASA conducted the first flight test of its next generation human-class Orion spacecraft. The flight was called the Exploration Flight Test -1 (EFT-1) which lasted for 4 hours and culminated into a re-entry trajectory at 9 km/s. This flight test of the 5-meter Orion Crew Module demonstrated various sub-systems including the Avcoat ablative thermal protection system (TPS) on the heat shield. The Avcoat TPS had been developed from the Apollo-era recipe with a few key modifications. The engineering for thermal sizing was supported by modeling, analysis, and ground tests in arc jet facilities. This paper will describe a postlfight analysis plan and present results from post-recovery inspections, data analysis from embedded sensors, TPS sample extraction and characterization in the laboratory. After the recovery of the vehicle, a full photographic survey and surface scans of the TPS were performed. The recovered vehicle showed physical evidence of flow disturbances, varying degrees of surface roughness, and excessive recession downstream of compression pads. The TPS recession was measured at more than 200 locations of interest on the Avcoat surface. The heat shield was then processed for sample extraction prior to TPS removal using the 7-Axis Milling machine at Marshall Space Flight Center. Around 182 rectangular TPS samples were extracted for subsequent analysis and investigation.
    [Show full text]
  • MIT Japan Program Working Paper 01.10 the GLOBAL COMMERCIAL
    MIT Japan Program Working Paper 01.10 THE GLOBAL COMMERCIAL SPACE LAUNCH INDUSTRY: JAPAN IN COMPARATIVE PERSPECTIVE Saadia M. Pekkanen Assistant Professor Department of Political Science Middlebury College Middlebury, VT 05753 [email protected] I am grateful to Marco Caceres, Senior Analyst and Director of Space Studies, Teal Group Corporation; Mark Coleman, Chemical Propulsion Information Agency (CPIA), Johns Hopkins University; and Takashi Ishii, General Manager, Space Division, The Society of Japanese Aerospace Companies (SJAC), Tokyo, for providing basic information concerning launch vehicles. I also thank Richard Samuels and Robert Pekkanen for their encouragement and comments. Finally, I thank Kartik Raj for his excellent research assistance. Financial suppport for the Japan portion of this project was provided graciously through a Postdoctoral Fellowship at the Harvard Academy of International and Area Studies. MIT Japan Program Working Paper Series 01.10 Center for International Studies Massachusetts Institute of Technology Room E38-7th Floor Cambridge, MA 02139 Phone: 617-252-1483 Fax: 617-258-7432 Date of Publication: July 16, 2001 © MIT Japan Program Introduction Japan has been seriously attempting to break into the commercial space launch vehicles industry since at least the mid 1970s. Yet very little is known about this story, and about the politics and perceptions that are continuing to drive Japanese efforts despite many outright failures in the indigenization of the industry. This story, therefore, is important not just because of the widespread economic and technological merits of the space launch vehicles sector which are considerable. It is also important because it speaks directly to the ongoing debates about the Japanese developmental state and, contrary to the new wisdom in light of Japan's recession, the continuation of its high technology policy as a whole.
    [Show full text]