DOCSLIB.ORG

Launch Systems

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Launch Systems T09‐1 T09-Ot2013 LAUNCH SYSTEMS [email protected] T09‐2 Outline •Rocket propulsion fundamentals • Launch vehicle dynamics –Gravity and drag losses – Contribution of Earth’s rotation • Launcher design –Staging – Mass fraction • Launch vehicles by agency: –Performance table • Launch sites T09‐3 Rocket propulsion fundamentals t t+dt Conservation of momentum: mv (m dm)(v dv) dmv' But v,v' abs. velocities (inertial frame) ue ejection speed (relative) v'(v dv) Hence: mv (m dm)(v dv) dm(v dv ue ) 0 mdv dmue dm m dt dv m F m u dt e T09‐4 The rocket equation The total impulse I = variation of the momentum of the ejected gas (and hence of the rocket in the opposite direction) is dv I m dt Fdt m u dt dt e If ue = constant (which is not always true), then I u mdt u dm u m e e e p with m = ejected mass. p I F I u Specific impulse Isp =impulse per unit mass: sp e m p m F ‐2 In many applications: I sp (in seconds) with g0 = 9.81 ms m g 0 T09‐5 The rocket equation The effect of the rocket engine on the vehicle can be determined from Newton’s second law dv m F I m g t dt sp 0 with mt the total mass of the vehicle. Since dm t m dt with the – sign meaning that the vehicle mass decreases, then dm I m g I g t sp 0 sp 0 dt m v I g dm m dv tf exp sp 0 t t mti I sp g 0 T09‐6 Propulsion system’s classifications 1) Function‐based classification – Primary propulsion • Acceleration, overcoming drag and gravity (launch) • Orbit insertion, orbit change maneuvers – Secondary propulsion •For pitch, yaw, and roll control, spin control, momentum wheel and gyro unloading (= attitude control), stage separation 2) Mission‐based classification – Launch vehicles and ballistic missiles – Upper stages and orbital transfer vehicles – Secondary and auxiliary propulsion – Interplanetary and space propulsion T09‐7 Launch system: definition Launch system = launch vehicle incorporating one or more stages + infrastructure for ground support. It provides velocity to place the spacecraft in the desired orbit, creates a severe ascent environment and protects the spacecraft from its surroundings. Kennedy Space Flight Center T09‐8 Launch vehicle dynamics •Space launch systems are unique forms of transportation since they are the only systems that accelerate continuously throughout their performance envelope. •Velocity is the fundamental measure of performance for launch systems. •A launch system’s ability to achieve orbital velocity comes primarily from its propulsion efficiency, with vehicle weight and drag acting against it. T09‐9 Launch vehicle dynamics Consider the motion of a launcher in a vertical plane. The following forces are acting on it: •The weight W, acting at the center of gravity cg •The thrust T, ideally acting through the centerline x (x = centerline) . •The aerodynamics forces – lift L and drag D – cg acting at the centre‐of‐pressure, cp. Trajectory γ cg Local horizontal Note that cp in general does not coincide with the aerodynamic center (ca). It does, e.g., in the case of axisymmetric bodies (not STS) T09‐10 Launch vehicle dynamics • Ideally, thrust, T, acts through the centerline of the vehicle but we can develop a control torque by gimballing the engine nozzle. •So in general T acts along a direction which forms an angle δ with the centerline •In a configuration with cg ahead of cp, L and D produce restoring torques. Trajectory γ cg Local horizontal δ T09‐11 Launch vehicle dynamics •Let’s decompose the forces into two directions: α = angle of attack • Parallel to the flight velocity (V): γ = flight‐path angle δ = gimbal angle dV [1] M T cos Mg sin D θ = pitch angle dt • Normal to the flight direction (perp. to V): d MV 2 [2] MV T sin Mg cos L cos dt r See next page T09‐12 t t // V n // L V u // Mg ur u r u // local horizon n r tu cos θ sin unur ; sinθ cos ur uuuuθ rr; θ V cos V Vt V (cosu sin u ) ru ru r r r V Vt Vt Vt V (sin u cosur ) V (cosu sinu r ) Vt Vn V (cosur sinu ) Vt Vn Vn V 2 cos V Vt Vn n r T09‐13 Pitching moment and aerodynamic stability • Separation of centre of pressure away from the centre of mass gives a restoring turning moment in pitch axis (stability) IF cp is lower than cg (this obtained by adding lift in the lower side ‐‐‐ fins = aletas) Modern full scale rockets do not usually rely on aerodynamics for stability. Full scale rockets pivot their exhaust nozzles to provide stability and control. That's why you don't see fins on a Delta, Titan, or Atlas booster. [http://www.grc.nasa.gov/WWW/k‐12/rocket/rktstab.html] T09‐14 Pitching moment •If cp is ahead of cg, L and D produce destabilizing torques in pitch. • Stabilization is obtained by adding fins or spinning the rocket (military missiles) or controlling the direction of thrust (majority of launchers use this). •IfIP = moment of inertia in pitch axis, then dh d d2 p I I dt p dt P dt 2 p d2 I L cos Dsin l T sin l P dt 2 0 T [see http://microgravity.grc.nasa.gov/education/rocket/gimbaled.html] T09‐15 Losses in the rocket equation dV T D •For small , [1] becomes g sin [3] dt M M • Thrust is given by the rate of mass flow: dM T I sp g0 dt • Substitute into [3] and obtain: dV d D I g ln M g sin dt sp 0 dt M •Then integrate to get velocity at the end of the thrust (0,t): M t t D V V I g ln 0 g sin dt dt 0 sp 0 M 0 0 M ‘Ideal velocity ‘Aerodynamic loss’ change’ ‘Gravity loss’ T09‐16 Gravity loss & gravity turn t V g sin dt A way to reduce it is minimizing the burn time g 0 •It reflects the work made by the vehicle against the Earth’s gravity •This integral cannot be computed a priori as we do not know the profile of γ and g as functions of time ( numerical integration). •In general, in order to reduce this term, we need to minimize the burn time • Satellite launch vehicles take off vertically and, at injection into orbit, they must be flying parallel to the Earth’s surface •During the initial phase, the rocket builds up speed in an almost vertical trajectory taking it above the densest layers of the atmosphere •When it transitions to the thinner upper atmosphere, the trajectory bends over, trading vertical speed for horizontal speed so that the rocket can achieve orbital perigee velocity at burnout. •This gradual transition is caused by the force of gravity and is called gravity turn T09‐17 ...Gravity turn •At lift off, the rocket is vertical and γ = 90 deg. •After clearing the tower and gaining speed, gimbaling the main engines produces a small pitchover, establishing an initial flight path angle γ0 •Thereafter γ will continue to decrease at a rate dictated by [2] d MV 2 MV T sin Mg cos L cos dt r •For simplicity, assume no lift (L=0) and no gimbaling (α+δ=0): d V 2 V g cos dt r • E.g., at h = 2 km, if γ = 85 deg and V = 110 m/s, dγ /dt = ‐0.44 deg/s •Over the ascent, the coefficient of cos γ is negative •As the speed increases, the coefficient of cos γ decreases in modulus, so the the rate of change of γ decreases (in abs. value). •Gravity turn reduces gravity loss because this term depends on sin γ T09‐18 Drag loss or aerodynamic loss t D 1 2 V dt D V ACD D 2 0 M •It reflects the work made against the aerodynamic resistance •Since M is proportional to initial mass M0, then A A 1 VD M 0 Alrocket lrocket It means that drag loss is inversely proportional to rocket length and rocket density. Shorter rockets will suffer from worse drag losses. Increasing density and decreasing drag coefficient might help T09‐19 Ideal ΔV M t t D V V I g ln 0 g sin dt dt 0 sp 0 M 0 0 M Overall V V V0 VIdeal Vg VD •We can easily estimate the velocity that a launch vehicle should provide by VIdeal V Vg VD where V is the velocity required for the desired orbit. V •We add the velocity losses, Vg and VD , to the velocity to obtain the required design velocity. T09‐20 Gravity and drag loss •Gravity loss ≈ 18% of VIdeal •Drag loss ≈ 2% of VIdeal Launcher Gravity loss (m/s) Drag loss (m/s) Ariane 4 1576 135 Atlas I 1395 110 Delta 7925 1150 136 Saturn V is the tallest Shuttle 1222 107 and suffers the smallest Saturn V 1534 40 drag loss Titan IV /Centaur 1422 156 T09‐21 Velocity increase due to Earth’s rotation N Considering the rotation of the Earth: E VIdeal V Vg VD VROT VROT E RE cos()cosa 2 φ = launch latitude a = launch azimuth ωE ∙RE = VE = velocity of a point on the Earth equator = 0.465 km/s ωE ∙RE∙cos (φ) = velocity of launch site N a On the local horizon of E launch site (a = azimuth of launch) T09‐22 Effect on allowed azimuth range At each stage burnout, the corresponding structure is jettisoned.
Load more

Recommended publications
  • Coplanar Air Launch with Gravity-Turn Launch Trajectories
    View metadata, citation and similar papers at core.ac.uk brought to you by CORE provided by AFTI Scholar (Air Force Institute of Technology) Air Force Institute of Technology AFIT Scholar Theses and Dissertations Student Graduate Works 3-23-2004 Coplanar Air Launch with Gravity-Turn Launch Trajectories David W. Callaway Follow this and additional works at: https://scholar.afit.edu/etd Part of the Aerospace Engineering Commons Recommended Citation Callaway, David W., "Coplanar Air Launch with Gravity-Turn Launch Trajectories" (2004). Theses and Dissertations. 3922. https://scholar.afit.edu/etd/3922 This Thesis is brought to you for free and open access by the Student Graduate Works at AFIT Scholar. It has been accepted for inclusion in Theses and Dissertations by an authorized administrator of AFIT Scholar. For more information, please contact [email protected]. COPLANAR AIR LAUNCH WITH GRAVITY-TURN LAUNCH TRAJECTORIES THESIS David W. Callaway, 1st Lieutenant, USAF AFIT/GAE/ENY/04-M04 DEPARTMENT OF THE AIR FORCE AIR UNIVERSITY AIR FORCE INSTITUTE OF TECHNOLOGY Wright-Patterson Air Force Base, Ohio APPROVED FOR PUBLIC RELEASE; DISTRIBUTION UNLIMITED. The views expressed in this thesis are those of the author and do not reflect the official policy or position of the United States Air Force, Department of Defense, or the United States Government. AFIT/GAE/ENY/04-M04 COPLANAR AIR LAUNCH WITH GRAVITY-TURN LAUNCH TRAJECTORIES THESIS Presented to the Faculty Department of Aeronautics and Astronautics Graduate School of Engineering and Management Air Force Institute of Technology Air University Air Education and Training Command In Partial Fulfillment of the Requirements for the Degree of Master of Science in Aeronautical Engineeering David W.
    [Show full text]
  • Trade Studies Towards an Australian Indigenous Space Launch System
    TRADE STUDIES TOWARDS AN AUSTRALIAN INDIGENOUS SPACE LAUNCH SYSTEM A thesis submitted for the degree of Master of Engineering by Gordon P. Briggs B.Sc. (Hons), M.Sc. (Astron) School of Engineering and Information Technology, University College, University of New South Wales, Australian Defence Force Academy January 2010 Abstract During the project Apollo moon landings of the mid 1970s the United States of America was the pre-eminent space faring nation followed closely by only the USSR. Since that time many other nations have realised the potential of spaceflight not only for immediate financial gain in areas such as communications and earth observation but also in the strategic areas of scientific discovery, industrial development and national prestige. Australia on the other hand has resolutely refused to participate by instituting its own space program. Successive Australian governments have preferred to obtain any required space hardware or services by purchasing off-the-shelf from foreign suppliers. This policy or attitude is a matter of frustration to those sections of the Australian technical community who believe that the nation should be participating in space technology. In particular the provision of an indigenous launch vehicle that would guarantee the nation independent access to the space frontier. It would therefore appear that any launch vehicle development in Australia will be left to non- government organisations to at least define the requirements for such a vehicle and to initiate development of long-lead items for such a project. It is therefore the aim of this thesis to attempt to define some of the requirements for a nascent Australian indigenous launch vehicle system.
    [Show full text]
  • Corporate Profile
    2013 : Epsilon Launch Vehicle 2009 : International Space Station 1997 : M-V Launch Vehicle 1955 : The First Launched Pencil Rocket Corporate Profile Looking Ahead to Future Progress IHI Aerospace (IA) is carrying out the development, manufacture, and sales of rocket projectiles, and has been contributing in a big way to the indigenous space development in Japan. We started research on rocket projectiles in 1953. Now we have become a leading comprehensive manufacturer carrying out development and manufacture of rocket projectiles in Japan, and are active in a large number of fields such as rockets for scientific observation, rockets for launching practical satellites, and defense-related systems, etc. In the space science field, we cooperate with the Japan Aerospace Exploration Agency (JAXA) to develop and manufacture various types of observational rockets named K (Kappa), L (Lambda), and S (Sounding), and the M (Mu) rockets. With the M rockets, we have contributed to the launch of many scientific satellites. In 2013, efforts resulted in the successful launch of an Epsilon Rocket prototype, a next-generation solid rocket which inherited the 2 technologies of all the aforementioned rockets. In the practical satellite booster rocket field, We cooperates with the JAXA and has responsibilities in the solid propellant field including rocket boosters, upper-stage motors in development of the N, H-I, H-II, and H-IIA H-IIB rockets. We have also achieved excellent results in development of rockets for material experiments and recovery systems, as well as the development of equipment for use in a space environment or experimentation. In the defense field, we have developed and manufactured a variety of rocket systems and rocket motors for guided missiles, playing an important role in Japanese defense.
    [Show full text]
  • Modeling Rocket Flight in the Low-Friction Approximation
    Undergraduate Journal of Mathematical Modeling: One + Two Volume 6 | 2014 Fall Article 5 2014 Modeling Rocket Flight in the Low-Friction Approximation Logan White University of South Florida Advisors: Manoug Manougian, Mathematics and Statistics Razvan Teodorescu, Physics Problem Suggested By: Razvan Teodorescu Follow this and additional works at: https://scholarcommons.usf.edu/ujmm Part of the Mathematics Commons UJMM is an open access journal, free to authors and readers, and relies on your support: Donate Now Recommended Citation White, Logan (2014) "Modeling Rocket Flight in the Low-Friction Approximation," Undergraduate Journal of Mathematical Modeling: One + Two: Vol. 6: Iss. 1, Article 5. DOI: http://dx.doi.org/10.5038/2326-3652.6.1.4861 Available at: https://scholarcommons.usf.edu/ujmm/vol6/iss1/5 Modeling Rocket Flight in the Low-Friction Approximation Abstract In a realistic model for rocket dynamics, in the presence of atmospheric drag and altitude-dependent gravity, the exact kinematic equation cannot be integrated in closed form; even when neglecting friction, the exact solution is a combination of elliptic functions of Jacobi type, which are not easy to use in a computational sense. This project provides a precise analysis of the various terms in the full equation (such as gravity, drag, and exhaust momentum), and the numerical ranges for which various approximations are accurate to within 1%. The analysis leads to optimal approximations expressed through elementary functions, which can be implemented for efficient flight
    [Show full text]
  • A Value Proposition for Lunar Architectures Utilizing On-Orbit Propellant Refueling
    A VALUE PROPOSITION FOR LUNAR ARCHITECTURES UTILIZING ON-ORBIT PROPELLANT REFUELING By James Jay Young In Partial Fulfillment of the Requirements for the Degree of Doctor of Philosophy in the School of Aerospace Engineering Georgia Institute of Technology May 2009 Copyright © 2009 by James J. Young A VALUE PROPOSITION FOR LUNAR ARCHITECTURES UTILIZING ON-ORBIT PROPELLANT REFUELING Approved by: Dr. Alan W. Wilhite, Chairman Dr. Douglas Stanley School of Aerospace Engineering School of Aerospace Engineering Georgia Institute of Technology Georgia Institute of Technology Dr. Trina M. Chytka Dr. Daniel P. Schrage Vehicle Analysis Branch School of Aerospace Engineering NASA Langley Research Center Georgia Institute of Technology Dr. Carlee A. Bishop Electronics Systems Laboratory Georgia Tech Research Institute Date Approved: October 29, 2008 ACKNOWLEDGEMENTS As I sit down to acknowledge all the people who have helped me throughout my career as a student I realized that I could spend pages thanking everyone. I may never have reached all of my goals without your endless support. I would like to thank all of you for helping me achieve me goals. I would like to specifically thank my thesis advisor, Dr. Alan Wilhite, for his guidance throughout this process. I would also like to thank my committee members, Dr. Carlee Bishop, Dr. Trina Chytka, Dr. Daniel Scharge, and Dr. Douglas Stanley for the time they dedicated to helping me complete my dissertation. I would also like to thank Dr. John Olds for his guidance during my first two years at Georgia Tech and introducing me to the conceptual design field. I must also thank all of the current and former students of the Space Systems Design Laboratory for helping me overcome any technical challenges that I encountered during my research.
    [Show full text]
  • 5.0 Launch Vehicle Performance
    Final Design Report on Converting the Minuteman Missile into a Small Satellite Launch System Submitted to Dr. George W. Botbyl USRA Design Professor Department of Aerospace Engineering The University of Texas at Austin by The Minotaur Design Team Rodrick McHaty - Team Leader Rodolfo Gonzalez - Chief Engineer Vu Pham - Chief Administrative Officer Engineers: Gordon MacKay Greg Humble Bill Alexander November 24, 1993 EXECUTIVE SUMMARY m Introduction Due to the Strategic Arms Reduction Talks (START) treaty between the United States and Ex-Soviet Union, 450 Stage III Minuteman II (MMII) missiles were -q recently taken out of service. Minotaur Designs Incorporated (MDI) intends to StagelI convert the MMII ballistic missile from a "" nuclear" warhead carrier into a small- "] L__________ satellite launcher. MDI will perform this conversion by acquiring the Minuteman Stage I stages, purchasing currently available control wafers, and designing a new ] shroud and interfaces for the satellite. Figure 1. MDI missile MDI is also responsible for properly integrating all systems. The new MDI system still System Description incorporates the original range-safety raceway and attitude-control actuators. MDI plans to purchase the 52 inch Figure 1 shows a representation of diameter avionics, range-safety, and the MDI missile. The stages that MDI attitude-control wafers from Martin will acquire from the Air Force are the Marietta's Multi-Service Launch System MMII stage I and II, and the MMIII stage (MSLS), "D" configuration missile, III. These stages define the propulsion which is currently under development. system of the MDI missile, and an MDI has designed a mechanical analysis of attainable orbits is performed.
    [Show full text]
  • A Tool for Preliminary Design of Rockets Aerospace Engineering
    A Tool for Preliminary Design of Rockets Diogo Marques Gaspar Thesis to obtain the Master of Science Degree in Aerospace Engineering Supervisor : Professor Paulo Jorge Soares Gil Examination Committee Chairperson: Professor Fernando José Parracho Lau Supervisor: Professor Paulo Jorge Soares Gil Members of the Committee: Professor João Manuel Gonçalves de Sousa Oliveira July 2014 ii Dedicated to my Mother iii iv Acknowledgments To my supervisor Professor Paulo Gil for the opportunity to work on this interesting subject and for all his support and patience. To my family, in particular to my parents and brothers for all the support and affection since ever. To my friends: from IST for all the companionship in all this years and from Coimbra for the fellowship since I remember. To my teammates for all the victories and good moments. v vi Resumo A unica´ forma que a humanidade ate´ agora conseguiu encontrar para explorar o espac¸o e´ atraves´ do uso de rockets, vulgarmente conhecidos como foguetoes,˜ responsaveis´ por transportar cargas da Terra para o Espac¸o. O principal objectivo no design de rockets e´ diminuir o peso na descolagem e maximizar o payload ratio i.e. aumentar a capacidade de carga util´ ao seu alcance. A latitude e o local de lanc¸amento, a orbita´ desejada, as caracter´ısticas de propulsao˜ e estruturais sao˜ constrangimentos ao projecto do foguetao.˜ As trajectorias´ dos foguetoes˜ estao˜ permanentemente a ser optimizadas, devido a necessidade de aumento da carga util´ transportada e reduc¸ao˜ do combust´ıvel consumido. E´ um processo utilizado nas fases iniciais do design de uma missao,˜ que afecta partes cruciais do planeamento, desde a concepc¸ao˜ do ve´ıculo ate´ aos seus objectivos globais.
    [Show full text]
  • Russian and Chinese Responses to U.S. Military Plans in Space
    Russian and Chinese Responses to U.S. Military Plans in Space Pavel Podvig and Hui Zhang © 2008 by the American Academy of Arts and Sciences All rights reserved. ISBN: 0-87724-068-X The views expressed in this volume are those held by each contributor and are not necessarily those of the Officers and Fellows of the American Academy of Arts and Sciences. Please direct inquiries to: American Academy of Arts and Sciences 136 Irving Street Cambridge, MA 02138-1996 Telephone: (617) 576-5000 Fax: (617) 576-5050 Email: [email protected] Visit our website at www.amacad.org Contents v PREFACE vii ACRONYMS 1 CHAPTER 1 Russia and Military Uses of Space Pavel Podvig 31 CHAPTER 2 Chinese Perspectives on Space Weapons Hui Zhang 79 CONTRIBUTORS Preface In recent years, Russia and China have urged the negotiation of an interna - tional treaty to prevent an arms race in outer space. The United States has responded by insisting that existing treaties and rules governing the use of space are sufficient. The standoff has produced a six-year deadlock in Geneva at the United Nations Conference on Disarmament, but the parties have not been inactive. Russia and China have much to lose if the United States were to pursue the programs laid out in its planning documents. This makes prob - able the eventual formulation of responses that are adverse to a broad range of U.S. interests in space. The Chinese anti-satellite test in January 2007 was prelude to an unfolding drama in which the main act is still subject to revi - sion.
    [Show full text]
  • Small Satellite Access to ESPA Standard Service
    SSC10-IX-1 Small Satellite Access to ESPA Standard Service Mr. Ted Marrujo DoD Space Test Program (STP), Mission Design 3548 Aberdeen Ave. SE, Kirtland AFB NM 87117; (505) 853-3338 Ted. [email protected] Lt Jake Mathis Space and Missile Systems Center, Launch and Range Systems Wing, Engineering 483 N. Aviation Blvd, Los Angeles AFB, CA; (310) 653-3392 [email protected] Mr. Caleb C. Weiss United Launch Alliance, Mission Integration 12257 S. Wadsworth Blvd, TSB-B7140, Littleton, CO 80125; (303) 977-0843 [email protected] ABSTRACT The DoD Space Test Program (STP), the Air Force Launch and Range Systems Wing (LRSW), and United Launch Alliance (ULA) are teaming up to provide a rideshare service to small satellites (<400lb) using an Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA). This rideshare service is an opportunity on EELV missions with margin to carry auxiliary payloads (APLs). This paper will define the ESPA, the standard rideshare service provided to APLs, and how APLs can access this service. We will discuss the roles and responsibilities the different government organizations, ULA, and the small satellite provider have in accessing and implementing ESPA Standard Service. In brief, ULA builds the EELV and performs the launch service, LRSW is responsible for developing and acquiring EELVs from ULA, and STP is responsible for identifying and manifesting APLs that meet ESPA Standard Service requirements. We will further define the processes and procedures required to implement ESPA Standard Service to include: how a particular EELV mission is selected to host ESPA Standard Service, the selection process for auxiliary satellites to utilize the capability, the requirements and timelines small satellites must meet to qualify, and the scope of services provided by ULA as part of Standard Service.
    [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]
  • Successful Launch of the First Epsilon Launch Vehicle
    Mitsubishi Heavy Industries Technical Review Vol. 51 No. 1 (March 2014) 59 Contribution to Japan's Flagship Launch Vehicle – part 2 -Successful Launch of the first Epsilon Launch Vehicle- TATSURU TOKUNAGA*1 NOBUHIKO KOHARA*2 KATSUYA HAKOH*3 TSUTOMU TAKAI*4 KYOICHI UI*5 TETSUYA ONO*6 On September 14, 2013, the first Epsilon Launch Vehicle was launched from (Independent Administrative Institution) Japan Aerospace Exploration Agency (JAXA) Uchinoura Space Center, and succeeded in properly injecting a satellite into orbit. In Epsilon Launch Vehicle development, we participate in the development/manufacture of the second-stage reaction control system(RCS) and modification maintenance of the launcher for the Epsilon launch system. The development/maintenance details and launch results are introduced in this report. |1. Introduction JAXA started development of the Epsilon Launch Vehicle in 2010, going through the stages of vehicle development, manufacturing, and maintenance of launch-related facilities, until the first Epsilon Launch Vehicle was launched from Uchinoura Space Center in 2013. We contributed to the successful launch of the first Epsilon Launch Vehicle through development of the second-stage reaction control system, which was equipped onto the launch vehicle, and modification maintenance of the launcher. Details of the development/maintenance and the launch results are introduced here. |2. Approach to Epsilon Launch Vehicle Development 2.1 General The Epsilon Launch Vehicle is the three-staged solid rocket developed by JAXA since 2010, and is the successor to the M-V launch vehicle technology, which completed operations in 2006, and develops to organize technical application/commonality of the H-IIA launch vehicle.
    [Show full text]
  • Hybrid Rocket Engine Design Optimization at Politecnico Di Torino: a Review
    aerospace Review Hybrid Rocket Engine Design Optimization at Politecnico di Torino: A Review Lorenzo Casalino † , Filippo Masseni † and Dario Pastrone *,† Politecnico di Torino, Dipartimento di Ingegneria Meccanica e Aerospaziale, Corso Duca degli Abruzzi 24, 10129 Torino, Italy; [email protected] (L.C.); fi[email protected] (F.M.) * Correspondence: [email protected] † These authors contributed equally to this work. Abstract: Optimization of Hybrid Rocket Engines at Politecnico di Torino began in the 1990s. A comprehensive review of the related research activities carried out in the last three decades is here presented. After a brief introduction that retraces driving motivations and the most significant steps of the research path, the more relevant aspects of analysis, modeling and achieved results are illustrated. First, criteria for the propulsion system preliminary design choices (namely the propellant combination, the feed system and the grain design) are summarized and the engine modeling is presented. Then, the authors describe the in-house tools that have been developed and used for coupled trajectory and propulsion system design optimization. Both deterministic and robust-based approaches are presented. The applications that the authors analyzed over the years, starting from simpler hybrid powered sounding rocket to more complex multi-stage launchers, are then presented. Finally, authors’ conclusive remarks on the work done and their future perspective in the context of the optimization of hybrid rocket propulsion systems are reported. Citation: Casalino, L.; Masseni, F.; Pastrone, D. Hybrid Rocket Engine Keywords: Hybrid Rocket Engines; multidisciplinary optimization; robust optimization Design Optimization at Politecnico di Torino: A Review. Aerospace 2021, 8, 226.
    [Show full text]