vol. 4 n.4 - journal of aerospace technology and management

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Vol.4 N.4Oct./Dec.2012

ISSN 1984-9648ISSN 2175-9146 (online)

Journal of Aerospace Technology and Management

Vol. 4, n.4 (Oct./Dec. 2012) – São José dos Campos: Zeppelini Editorial, 2012Quartely issued

1. Aerospace sciences2. Technologies3. Aerospace engineering

CDU: 629.73

General Information

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EDITORIAL COMMITTEE

EDITOR IN CHIEF

Ana Cristina Avelar – IAE – São José dos Campos/SP – [email protected]

EXECUTIVE EDITOR

Ana Marlene F. Morais – IAE – São José dos Campos/SP – [email protected]

SCIENTIFIC COUNCIL

Angelo Passaro – IEAv – São José dos Campos/SP– BrazilAntonio Pascoal Del'Arco Jr. – IAE – São José dos Campos/SP– BrazilCarlos Antônio M. Kasemodel – IAE – São José dos Campos/SP– BrazilCarlos de Moura Neto – ITA – São José dos Campos/SP– BrazilEduardo Morgado Belo – EESC/USP – São Carlos/SP – BrazilFrancisco Carlos M. Pantoja – DIRENG - Rio de Janeiro/RJ– BrazilFrancisco Cristovão L. Melo – IAE – São José dos Campos/SP– BrazilJoão Marcos T. Romano – UNICAMP – Campinas/SP – BrazilMarco A. Sala Minucci – VALE Soluções em Energia – São José dos Campos/SP – BrazilMischel Carmen N. Belderrain – ITA – São José dos Campos – BrazilPaulo Tadeu de Melo Lourenção – EMBRAER– São José dos Campos/SP– Brazil Rita de Cássia L. Dutra – IAE – São José dos Campos/SP– Brazil

ASSOCIATE EDITORS

Acir Mércio Loredo Souza – UFRGS – Porto Alegre/RS – BrazilAdam S. Cumming – DSTL – Salisbury/Wiltshire–EnglandAdiel Teixeira de Almeida – UFPE – Recife/PE – BrazilAlain Azoulay – SUPELEC– Gif–Sur–Yvette – FranceAlexandre Queiroz Bracarense – UFMG – Belo Horizonte/MG – BrazilAltamiro Susin – UFRGS – Porto Alegre/RS – BrazilÁlvaro Damião – IEAv– São José dos Campos/SP– BrazilAndré Fenili – UFABC – Santo André/SP – BrazilAntonio F. Bertachini – INPE – São José dos Campos/SP–Brazil Antonio Henriques de Araújo Jr – UniFOA – Volta Redonda/RJ – BrazilAntonio Sergio Bezerra Sombra – UFC – Fortaleza/CE – BrazilBert Pluymers – KU – Leuven – BelgiumCarlos Henrique Marchi – UFPR – Curitiba/PR – BrazilCarlos Henrique Netto Lahoz – IAE – São José dos Campos/SP – BrazilCosme Roberto Moreira da Silva – UnB – Brasília/DF – BrazilCynthia Junqueira – IAE – São José dos Campos/SP– Brazil Daniel Alazard – ISAE – Toulouse – FranceDavid Murray–Smith – University of Glasgow – Glasgow – ScotlandEdson Aparecido de A. Querido Oliveira – UNITAU – Taubaté/SP – Brazil

Journal of Aerospace Technology and ManagementJ. Aerosp. Technol. Manag.Vol.4, No 4, Oct.-Dec., 2012

401

Edson Cocchieri Botelho – FEG/UNESP – Guaratinguetá/SP – BrazilElizabeth da Costa Mattos – IAE – São José dos Campos/SP– BrazilFabiano Fruett – UNICAMP – Campinas/SP – BrazilFabrice Burel – INSA – Lion – FranceFlamínio Levy Neto – UnB – Brasília/DF – BrazilFrancisco José de Souza – UFU – Uberlândia/MG – BrazilGilberto Fisch – IAE – São José dos Campos/SP– BrazilGilson da Silva – INPI – Rio de Janeiro/RJ – BrazilHugo H. Figueroa – UNICAMP – Campinas/SP – BrazilJoão Luiz F. Azevedo – IAE – São José dos Campos/SP – BrazilJosé Alberto Cuminato – ICMC/USP – São Carlos/SP– BrazilJosé Atílio Fritz F. Rocco – ITA – São José dos Campos/SP – BrazilJosé Leandro Andrade Campos – UC – Coimbra – Portugal José Rubens G. Carneiro – PUC Minas – Belo Horizonte – BrazilJosé Márcio Machado – IBILCE/UNESP – São José do Rio Preto/SP – BrazilJosé Maria Fonte Ferreira – UA – Aveiro – PortugalJosé Pissolato Filho – UNICAMP – Campinas/SP – BrazilJosé Roberto de França Arruda – UNICAMP – Campinas/SP – BrazilLuís Carlos de Castro Santos – EMBRAER– São José dos Campos/SP– BrazilLuiz Claudio Pardini – IAE – São José dos Campos/SP– BrazilMarcello Faraco de Medeiros – EESC/USP – São Carlos/SP – BrazilMárcia B. H. Mantelli – UFSC– Florianópolis/SC – B BrazilMarc Lesturgie – ONERA– Palaiseau–FranceMarcos Pinotti Barbosa – UFMG– Belo Horizonte/MG – BrazilMichael Gaster – Queen Mary University of London – London – EnglandMichele Leali Costa – FEG/UNESP – Guaratinguetá/SP – BrazilMirabel Cerqueira Rezende – IAE – São José dos Campos/SP– BrazilOthon Cabo Winter – FEG/UNESP – Guaratinguetá/SP – BrazilPaulo Celso Greco – EESC/USP – São Carlos/SP – BrazilPaulo Sérgio Varoto – EESC/USP – São Carlos/SP – Brazil Raimundo Freire –UFCG– Campina Grande/PB–BrazilRenato Machado Cotta – UFRJ – Rio de Janeiro/RJ – Brasil Roberto Costa Lima – IPqM – Rio de Janeiro/RJ – BrazilRomis R. F. Attux – UNICAMP – Campinas/SP– BrasilSamuel Machado Leal da Silva – CTEx – Rio de Janeiro /RJ– BrazilSandro Haddad – UnB– Brasília/DF–BrazilSelma Shin Shimizu Melnikoff – EP/USP – São Paulo/SP – BrazilSérgio Frascino M. Almeida – ITA – São José dos Campos/SP – BrazilUlrich Teipel – Georg Simon OHM – Nürnberg – GermanyValder Steffen Junior – UFU – Uberlândia/MG – Brazil

Waldemar de Castro Leite Filho – IAE – São José dos Campos/SP – BrazilWillian Roberto Wolf – IAE– São José dos Campos/SP – Brazil Wim P. C. de Klerk – TNO – Rijswijk/SH – The Netherlands

EDITORIAL PRODUCTION

Glauco da Silva – IAE – São José dos Campos/SP– BrazilHelena Prado A.Silva – IAE – São José dos Campos/SP– BrazilJanaina Pardi Moreira – IAE – São José dos Campos/SP– BrazilLucia Helena de Oliveira – DCTA – São José dos Campos/SP– BrazilMônica E. Rocha de Oliveira – INPE – São José dos Campos/SP–Brazil

402

CONTENTS

EDITORIAL405 Access to Space in Brazil – Current and future scenarios Carlos Antônio Magalhães Kasemodel

REVIEW ARTICLE407 Hypergolic Systems - A Review in Patents Gilson da Silva, Koshun Iha

ORIGINAL PAPERS413 Generation of an Atomic Beam by Using Laser Ablation for Isotope Separation Purposes Juliana Barranco de Matos, Márcio de Lima Oliveira, Emmanuela Melo de Andrade Sternberg, Marcelo Geraldo

Destro, Rudimar Riva, Nicolau André Silveira Rodrigues

4 1 Occurrence of Defects in Laser Beam Welded Al-Cu-Li S eets wit T- oint Con guration André Luiz de Carvalho Higashi, Milton Sérgio Fernandes de Lima

431 Multidisciplinary Design Optimization of Sounding Rocket Fins Shape Using a Tool Called MDO-SONDA Alexandre Nogueira Barbosa, Lamartine Nogueira Frutuoso Guimarães

443 Studies on In uence of Testing Parameters on Dynamic and Transient Properties of Composite Solid Rocket Propellants Using Dynamic Mechanical Analyzer

Vilas Wani, Mehilal, Sunil Jain, Praveen Prakash Singh, Bikash Bhattacharya

453 Kinematic Analysis of the Deployable Truss Structures for Space Applications Xu Yan, Guan Fu-ling, Zheng Yao, Zhao Mengliang

ISSN 1948 - 9648ISSN 2175 - 9146 (online)

Vol.04, N°4, Oct. – Dec. 2012

403

463 Wind Tunnel Simulation of the Atmospheric Boundary Layer for Studying of the Wind Pattern at Centro de Lançamento de Alcântara

Ana Cristina Avelar, Fabrício Lamosa Carneiro Brasileiro, Adolfo Gomes Marto, Edson R. Marciotto, Gilberto Fisch, Amanda Fellipe Faria

475 Propeller Induced Effects on the Aerodynamics of a Small Unmanned-Aerial-Vehicle Adnan Maqsood, Foong Herng Huei, Tiauw Hiong Go

481 Effect of Dielectric Barrier Discharge on the Air ow Around a Cylinder Ashraf El Droubi, Dawson Tadeu Izola

489 Electronic Simulator of the PLATO Satellite Imaging System Rafael Corsi Ferrão, Sergio Ribeiro Augusto, Tiago Sanches da Silva, Vanderlei Cunha Parro

THESIS ABSTRACTS495 A Comprehensive Investigation of Retrodirective Cross-Eye Jamming Warren Paul du Plessis

495 Ant Colony Optimization Applied to Laminated Composite Materials Rubem Matimoto Koide

496 Classifying Low Probability of Intercept Radar Using Fuzzy ARTMAP Pieter Frederick Potgieter

497 AD HOC REFEREES

499 INSTRUCTIONS TO AUTHORS

404

EditorialAccess to Space in Brazil – Current and future scenariosBrig. Eng. Carlos Antônio M. Kasemodel*Director of Instituto de Aeronáutica e Espaç[email protected]

The search for autonomy to access to space has always been the objective of the Brazilian Space Program. At the end of the 1970s, the Complete Brazilian Space Mission – MECB established the goal to develop a national satellite to be launched, by a launch vehicle designed and manufactured in the country, from a launch site located in Brazil. In order to master the critical technologies to build this launch vehicle, sounding rockets of a family named SONDA were developed. With these, technologies were acquired to produce solid propellants, thermal protections, stage separation systems, motor structures made of highly resistant steel, structures of composite materials, attitude control systems, pyrotechnic devices, on-board electronics systems, as well as the associated ground support equipment. Even though MECB goal has not been completely achieved as initially planned, due to misalignments in the stage of development on its three segments (while the Launch Center was created in 1987, the satellite was concluded in 1992, the

consolidation of the development of space technology in the country.

reviewed, resulting in the re-design of electrical and pyrotechnic networks, besides many minor changes in other systems, as well as the conducting of new ground tests. Nowadays, the construction of three other prototypes is predicted for the conclusion

national inertial navigation system; the second one concerns the complete test of the vehicle with a technological payload; and the third one aims to launch a national satellite into orbit. In 2005, in order to establish long term goals to develop launchers in the country, the Cruzeiro do Sul Program was proposed,

into geostationary transfer orbit.

low national demand for geostationary orbit satellites and the existence of a binational company, Alcântara Cyclone Space, with

Cruzeiro do Sul program, the

Applied Physics by the Naval Postgraduate School, in the United States, in 1999. Master of Business Administration in Advanced Development of Executives – department of Institutional Strategic Management, by Universidade Federal Fluminense, in 2007, and postgraduation in Air

J. Aerosp. Technol. Manag., São José dos Campos, Vol.4, No 4, pp. 405-406, Oct.-Dec., 2012 405

vehicle, and all of them will use solid propellant and motor cases in composite material.

Studies show that with the initiatives proposed in this review, 75% of the national needs concerning satellite launches will be met, and also the knowledge of critical technologies to access space will be ensured, including the development of larger vehicles, in case of future needs. As established by the National Program of Space Activities – PNAE and by the National Defense Strategy – END, considering that Brazil is a country of large dimensions, with extensive land and sea borders, it cannot give up the knowledge of space technology and the autonomous capability to access space. Also, according to the last document, “Whoever does not master critical technologies is neither independent for defense nor for development”.

J. Aerosp. Technol. Manag., São José dos Campos, Vol.4, No 4, pp. 405-406, Oct.-Dec., 2012406

LIST OF SYMBOLS

DETA DiethylenetriamineEDA EthylenediamineTNT 2, 4, 6-trinitrotolueneHMX OctogenRDX HexogenDNAZ 2-(N,N-dimethylamino)ethylazideTMEDA N,N,N’,N’-tetramethyl-ethylene-diamineTMPDA N,N,N’,N’-tetramethyl-1,3-diaminopropaneTMBDA N,N,N’,N’-tetramethyl-1,4-diaminobutaneAU Astronomical unit (149,597,870,700 km)LH Liquid hydrogenLOX Liquid oxygenMON Mixed oxides of nitrogenGLP Gelled liquid propaneMMH MonomethylhydrazineHGF Hypergolic green fuelHEH HydroxyethylhydrazineHEHN Hydroxyethylhydrazinium nitrateFTIR Fourier transform infrared spectroscopyNTO Nitrogen tetroxide

INTRODUCTION

The term ‘hypergolic’ includes igniting spontaneously upon contact with the complementary explosive or energetic substance. Then, hypergolicity is the propriety of self-ignition

within milliseconds after fuel and oxidizer contact (Hawkins et al., 2011). This propriety is very important in propellant systems, because it can substitute the multistage rocket with the separate ignition system, resulting in high combustion

hypergolic reactions can be used to improve ignition of nitroarene explosives, applied in unexploded ordnance, like bombs and mines, to neutralize these explosives, in agreement with the research of Koppes et al. (2010). Koppes et al. (2010) taught a method to chemically neutralize a nitroarene explosive composition comprising in

and an accelerant by applying the nitroarene hypergolic to the

of the nitroarene hypergolic may include linear polyamines with or without nitrogen or other heteroatom within the structure of the compound, such as diethylenetriamine (DETA), ethylenediamine (EDA), propanediamine, and so on. The accelerant of the nitroarene hypergol may include appropriate hydridoborate salts (M+BH4), hydrazine, alkylated derivatives of hydrazine, or combinations. The nitroarene hypergol provides a decrease in the delay to ignition of 90% or more, in agreement with Koppes et al. (2010), and an increase in heat generation. The nitroarene compounds that can be neutralized with the hypergolic system include nitrotoluenes, nitrobenzenes, nitronaphthalenes, nitrophenoxyalkyl nitrates, and their derivatives. The hypergols are added as pure liquids or as mixtures with other liquid or solid hypergols. In an unusual coupling with TNT, the amines with terminal amine groups (primary amines), i.e.,

to provide a TNT-amine-TNT bridged product, with amines attaching at the ring carbons of the TNT bearing the methyl

doi: 10.5028/jatm.2012.04043812

Hypergolic Systems: A Review in PatentsGilson da Silva1*, Koshun Iha2

1Instituto Nacional da Propriedade Industrial – Rio de Janeiro/RJ – Brazil2Instituto Tecnológico de Aeronáutica – São José dos Campos/SP – Brazil

Abstract: Hypergolic reactions may be useful in civil and military applications. In the area of rocket propulsion, they constitute a potential eld due to the reduced eight and comple ity of fuel in ection systems, allo ing yet controlled use of the propulsors. This manuscript aimed at presenting different hypergolic systems and their particularities, comparing them ith chemical propulsion systems, hich are most commonly employed in rocket motors, for e ample.

Keywords: Hypergolic, ropellant, onomethylhydra ine, Hydro yethylhydra ine, i uid hydrogen, i uid o ygen.

Received: 05/07/12 Accepted: 30/07/12*author for correspondence: [email protected]ça Mauá, 7 – CentroCEP 20.081-240 Rio de Janeiro/RJ – Brazil


group. DETA does this by leaving the central amine function unreacted. To the extent that this bridging could be maximized on the surface of TNT, there would be an energy release compressed into the smallest time scale, which is a condition favoring the evolution and conservation of heat, and thus a decrease in delayed ignition.

CIVIL APPLICATION OF A HYPERGOLIC SYSTEM

reaction is taught by Blackburn (2006). In agreement with

many occupant restraint ones tends to be the heaviest and most complex component of the restraint system. Then, Blackburn (2006) systems can simplify the design and manufacturing of

The device comprises a cartridge formed from a container, two materials stored in the container; however, the second one

materials forms a hypergolic mixture upon contact with each other. Under exposure of the gas generating to an elevated

and second materials is breached, enabling the materials to combine to form the hypergolic mixture. A propellant charge may be used to generate gas, being the decomposition of the propellant initiated by the hypergolic reaction. The propellant can be a composition such as the ammonium nitrate. According to Blackburn (2006), the hypergolic reaction

form, like glycerol or any suitable alcohol such as polyvinyl alcohol, and a second material comprising potassium

increase the surface interaction between the materials.

MILITARY APPLICATION OF A HYPERGOLIC SYSTEM

Hypergolic reactions can be used to defense systems according to Thuman et al. (2011). A projectile comprising a reactive charge is used to promote the destruction of an explosive-charged weapon, such as bombs, homemade explosive devices, and air, water or ground craft comprising explosives. An explosive can be made to detonate by the shock effect, which is generated by a splinter when it hits the explosive at high speed or by a pressure wave from an explosive charge (blasting). Then, the systems proposed by Thuman et al. (2011)

shell upon impact so that a passage is opened into the explosive of the shell, through which passage the reactive charge is

the kinetic energy of the projectile. The projectile has a reactive charge disposed in at least one gas-and liquid-tight cavity to react and start a hypergolic reaction with the explosive. The projectile can have the gas- and/or liquid-tight container charged with zinc, zinc stearate, zirconium, magnesium perchlorate, bismuth trioxide, or a liquid, such as pyrrolidine. The gas- and liquid-tight container is constituted by the all-covering metal foil for preventing undesirable reactions with the surrounding atmosphere. When the reactive charge of the projectile is mixed with the weapon explosive, under effect from the kinetic energy of the projectile penetration, a reaction with the explosive occurs. Gas that is formed in the course of the burning generates an overpressure inside the weapon unit, which leads to splitting and destruction of the weapon unit. A suitable composition, being 99% by weight zinc and 1% by weight zinc stearate, is used, like a termed hypergolic composition, which, upon contact with the explosive weapon, spontaneously reacts.

PROPELLANTS

Useful propellant compositions were taught by Fawls et al. (2005), who described the effect of the oxygen in the metal passivation in the compositions of explosives and propellants. During the combustion process, the metal ingredients have an oxide shell formed in the surface that inhibits the oxidation of the metal, thereby reducing the overall available energy and forming a totally oxidized metal.

et al. (2005) taught how to increase the metal surface area, by means of the nanosized metallic particles, in combination with a halogenic oxidizer, to enhance the combustion of the metal by means of preventing the chemically-inhibiting

species is pyrolytically or chemically degraded in the combustion or explosive zones, releasing halogens in the

increasing the overall energy released. In general, the conventional metal nanoparticles can

Silva, G., Iha, K.


®, Viton®, or

energetic ingredients may be added to the fuel or hybrid grain to improve the energy output of these propellants, such as HMX, RDX, or other energetic ingredients. Liquid rocket propellant systems produce thrust by means of expulsion of high velocity exhaust gases made by the reaction between a fuel and an oxidizer. Non hypergolic systems are useful, but they need complex ignition systems with igniters and/or catalyst beds, which are expensive, introduce extra weight to the thrust, and increase the risk of failure (Natan et al., 2011).

HYPERGOLIC PROPELLANTS

The conventional hypergolic system is composed of hydrazine as the fuel component, being very toxic. Hawkins et al. (2011) taught a bipropellant fuel based upon salts containing the dicyanamide anion, employing nitrogen-containing, heterocyclic-based cations such as the imidazolium cation. While salt molecules contain highly energetic (formation of enthalpy), high nitrogen anions, the dicyanamide-based molecule solely displays fast ignition. The fuel proposed by Hawkins et al. (2011) is stored in a propulsion system fuel tank and the oxidizer in a separate one. The ignition happens just after the contact of the fuel and oxidizer sprayed into a chamber in a rocket. The ionic liquid fuel can provide greater than 40% improvement in density over hydrazine fuels. Watkins (2004) suggested a hypergolic fuel system comprising hydrogen peroxide, silane, and liquid fuel. In agreement with Watkins, in order that hydrogen peroxide is used as a propellant in rocket engines, a decomposition catalyst is required, which accelerates the decomposition of the hydrogen peroxide, however this technology is expensive. Therefore, a system where the hydrogen peroxide (H2O2) and silane (SiH4) are contacted to improve the decomposition of the peroxide, forming a gas that is contacted with a liquid fuel, igniting the liquid fuel (such as kerosene), was proposed. To thrust a rocket engine, the hydrogen peroxide is contacted with the silane in the combustion zone, at room temperature, to provide effective ignition. Upon ignition, a liquid fuel is fed to the combustion zone and combusted therein to provide thrust, since exhaust gases exit combustion zone of the rocket chamber through the exhaust outlet port. A hypergolic fuel propulsion system containing a fuel composition (azide compound) and an oxidizer composition (hydrogen peroxide) is showed by Hallit and Bauerle (2004).

The azide compound in the fuel has at least one tertiary nitrogen and one azide functional group in combination with a catalyst, which has at least one transition metal compound (preferably, cobalt and manganese). An azide functional group is represented by -N3. Upon oxidation after contact with the oxidizer composition, the azide compound loses nitrogen and reacts to produce the energy needed to provide thrust. The catalyst is added to the azide compound to produce a transition metal level in the fuel composition of about 0.2% or greater. The fuel composition and the oxidizer composition are brought into contact in stoichiometric ratios, which lead to the desired ignition. In general, with hypergolic fuels in rocket motor or missile engines, the oxidizer to fuel ratio may vary over a relatively wide range, depending on the performance desired, propellant tank pressures, and other operating parameters. In such bipropellant mixtures, the fuel and oxidizer are unstable when mixed together, and they are generally stored separately. Bipropellant rocket motor propulsion systems consist of oxidizer and fuel propellant tanks, pressurizing system, plumbing, valves, and engine. Currently known hypergolic, bipropellant rocket propulsion systems have a number of drawbacks. For example, one system consists of monomethylhydrazine (MMH) and red fuming nitric acid. Stevenson et al. (2011) proposed a fuel mixture to use as hypergolic liquid or gel fuel in bipropellant propulsion systems, with the chemical compounds preferably having similar ignition characteristics as monomethyl hydrazine, and not

combinations consist of one or more of a family of hypergolic

compound), mixed with one or more hypergolic tertiary diamine compounds (second compound). The hypergolic amine azides have the general structure (R1)(R2)(R3)N, in which R1, R2, an R3 can be an hydrogen and an aliphatic alkene, alkyne, or cycloalkyl group, without hetero-atoms or heterocyclic atoms, but where at least one of the R groups have an azide. Examples of hypergolic amine azides are the 2-(N, N-dimethylamino)ethylazide (DNAZ), 2-(N-cyclo-propylamino)ethylazide, bis(2-azidoethyl)methylamine, and so on. The tertiary diamines have the general formula R4R5N-R6-NR7R8, where R4, R5, R7 e R8 are aliphatic groups and R6 may be aliphatic, alkene, or alkyne groups. Examples of hypergolic diamines include the N,N,N’,N’-tetramethyl-ethylene-diamine (TMEDA), N,N,N’,N’-tetramethyl-1,3-diaminopropane (TMPDA), N,N,N’,N’-tetramethyl-1,4-diaminobutane (TMBDA), etc.

Hypergolic Systems: A Review in Patents


values, it is generally desirable to incorporate into the fuel the maximum percentage of amine azide compound, which will still allow an acceptably low ignition delay of about 3 milliseconds to about 15 milliseconds. For example, a fuel containing about 33.3% DMAZ and about 66.7% TMEDA has an ignition delay of about 9.0 milliseconds. DiSalvo (2012) described problems to prepare a low-storage temperature bipropellant for missions far from the sun greater than 3 AU, because the portion of the power budged consumed by heaters to prevent propellant freezing increases

points such as liquid hydrogen (LH) and liquid oxygen (LOX) are not suitable for using on planetary probes, because they require cryogenic storage vessels capable of containing them within several AU of the sun. In agreement with such author, propane is a potential low-temperature propellant, because

-189.9 ºC, and boils at -42.2 ºC. However, mixed oxides of nitrogen (MON) have freezing points not low enough to be ideal on deep space missions. Then, DiSalvo (2012) taught a method for preparing a bipropellant system comprising gelled liquid propane (GLP) fuel, which is well-suited for outer planet missions with additives, such as powders of boron, carbon, lithium and/or aluminum added to the fuel to improve its performance and enhance hypergolicity. The gelling agent can be silicon dioxide, clay, carbon, organic or inorganic polymers. The oxidizer for the low-temperature propellant combination is MON-30 (70%N2O4+30%NO), produced by an exothermic reaction (6000 kcal/kg) between nitric oxide and dinitrogen tetroxide/nitrogen dioxide. The reaction should be done in vacuum system with the tank into the ice water bath to maintain the temperature of the reactants at 0 ºC. The MON-30 can be gelled at around -25 ºC with 3% of fumed silica by weight, using a plate churn mixer and its freezing point is of -81 ºC. Propane can be gelled using a plate churn mixer placed

ethylene glycol mixture cooled at -55 ºC. A total of 20g of fumed silica is introduced into the mixing vessel, which is attached to a vacuum pump and cooled in dry ice. 500 grams of liquid propane is introduced into the mixing vessel. For the churning phase, the system is submerged in a 70/30-ethylene glycol/water bath and cooled to -55 ºC. The gelled propane has a freezing point of -189.9 ºC.

GREEN PROPELLANTS

Natan et al. (2011) showed a composition comprising a gelled fuel in which catalyst or reactive particles are suspended. The particles can ignite hypergolically with an oxidizer. The fuel and the oxidizer can be chosen from a wide spectrum of materials that are environmentally friendly (green propellants), without the need of carrying a complex ignition system. The catalyst or reactive particles can react spontaneously with an oxidizer or can be as catalysts to promote the ignition reaction. The rheological properties of the gelled fuel, e.g., the yield stress and the high viscosity while at rest, assure that no particle sedimentation takes place even at high acceleration levels of the vehicle. The hypergolic composition taught by them comprises at least one fuel in the form of a gel, at least one particulate ignition agent suspended in the fuel, and one oxidizer. The ignition agent is selected from the group consisting of hydrazine, its derivatives and a metal hydride (selected from the group consisting of sodium borohydride, lithium borohydride and potassium borohydride), it can comprise a hypergolic catalyst too, like an alkyl-substituted amine and metal salt (selected from the group consisting of an alkyl-substituted diamine and triamine and metal salt of an aliphatic carboxylic acid – such as acetate, propionate, and butyrate). The fuel is chosen from the group consisting of hydrocarbons, alcohols, amines, amides, metal-organic liquid compounds, alkaloids, and liquid hydrogen, with a gelling agent (nano-silica fumed powder, aluminum stearate and gelling polymers), and an oxidizer (hydrogen peroxide, liquid oxygen, nitrous oxide, nitrous acid, nitric acid, perchloric acid,

The method for preparing a hypergolic composition for rocket propellant comprises adding a gelling agent to the liquid fuel and suspending a particulate ignition agent in this fuel, upon contact with an oxidizer, the ignition agent initiates the reaction between the fuel and the oxidizer. MMH is a widely employed fuel in hypergolic and bipropellant systems. It has desirable propellant properties, but it is highly toxic, carcinogenic, and corrosive. A rocket fuel composition comprising one or more tertiary amine azides is taught by Sengupta (2008). The fuel is hypergolic when combined with a strong oxidizer, such as red fuming nitric acid, hydrogen peroxide, nitrogen tetroxide, or hydroxyl ammonium nitrate.

Silva, G., Iha, K.


In agreement with Smith et al. (2010), hypergolic green fuel (HGF) can be produced from 2-hydroxyethylhydrazine (HEH) by means of its nitration, resulting in hydroxyethylhydrazinium nitrate-acetone. At the beginning, the precursor HEH is pumped into a reactor under nitrogen atmosphere, set at 0.5 ºC, and deionized water is slowly added to the reactor under agitation. At 1 ºC and under agitation, nitric acid (HNO3) is slowly added, in order to prevent the temperature increasing above 10 ºC. The pH of the mixture is about 8 to 9 at the beginning and the nitric acid should be slowly added until pH is in the range of 4.8 to 5.0. The produced hydroxyethylhydrazinium nitrate (HEHN) is light yellow and should be transferred into a rotary evaporator, where the water is removed and neutralized until reaches 10% of water. The remaining water is removed by means of sparging the HEHN with nitrogen, then from the rotary evaporator and put into a storage vessel under nitrogen. To produce a hypergolic green fuel, HEHN produced must be mixed with the solvent acetone. Smith et al. (2010) showed a one-step synthesis process to prepare a HGF propellant from HEH. In agreement with them, the process included: providing a solution of acetone in 2-hydroxyethylhydrazine, wherein the solution is about 15 to 50% by volume acetone in HEH; and adding nitric acid containing less than 5% water to the acetone-HEH solution to form the HGF propellant, wherein the molar ratio of nitric acid to HEH is less than about 0.05:1 to 1.4:1. The process

et al. (2010) can be evaluated by means of a comparison between the sample and the reference Fourier transform infrared (FT-IR) spectrum.

INJECTORS

Bipropellant injection elements are useful in a typical liquid propellant rocket engine to facilitate the injection, distribution, mixing, and combustion of the elements in a combustion chamber. The injector may be composed by a system assembly, spark, to ignite the propellants by creating

from the propellant one. The size and mass are undesirable characteristics to this kind of system assembly, when used in small rocket engines (Fisher, 2009). There are also known spark ignition systems for providing ignition sparks within a reaction zone in the combustion

fabricating system components and pose problems with component degradation during use. For instance, special

oxidizer and to create an easily ignited mixture of propellants at the exposed electrodes. Direct spark ignition systems through an injector faceplate can also add weight, increase design complexity, and typically operate at off-optimum mixture ratios (usually at fuel-rich rations) to preclude thermal damage to the electrodes, but which lower overall combustion performance. Brown et al. (2010) taught a fuel manifold for the injector of a hypergolic rocket engine. According to them, hypergolic rocket engines that use the MON-25/MMH ((25% mixed oxides of nitrogen and 75% nitrogen tetroxide)/(Monomethylhydrazine)) propellant combination may be relatively sensitive to pulsing frequencies imparted form the propellant system. Thus, compact vehicles that provide relatively small packaging envelopes may only further complicate this sensitivity. The rocket engine proposed by Brown et al. (2010)

comprises a main fuel chamber that is generally frustroconical

formed within the injector body, generally along the axis such

of oxidizer manifold. The rocket engine includes yet a combustion chamber having an acoustic resonance frequency and a fuel manifold having a resonance frequency, which is at least an order of magnitude lower than the acoustic resonance frequency. An engine generally includes a thrust chamber assembly powered by a propellant system having a fuel and an oxidizer system. The fuel and oxidizer systems provide a fuel and an oxidizer into the thrust chamber assembly. The propellant combination self-ignites within the thrust chamber assembly to provide reliable performance and thrust. MON-25 is highly reactive with MMH and has a tendency to drive unstable combustion processes. It should be understood that other oxidizers, such as nitrogen tetroxide (NTO) and other fuels, may alternatively or additionally be utilized. The combustion chamber is retained adjacent to an injector body through a chamber retention ring. A valve system selectively communicates the propellant combination into the injector body. The oxidizer manifold may be at least partially

The fuel manifold may be utilized for any bipropellant rocket engine that operates at several thrust levels from, for example, relatively small thrust attitude control thrusters, medium thrust divert engines, or large axial engine rocket engines.

Hypergolic Systems: A Review in Patents


FINAL CONSIDERATIONS

The useful common propulsion systems need a complex ignition system, which are expensive and introduce extra weight to the thrust. Solid rocket systems do not allow a controlled actuation of the propulsion and show increased risk of the failure in the ignition step. The liquid propulsion brings security (LH/LOX) and/or healthy (MMH) problems. Then, a hypergolic propellant system seems to be the most secure and controlled system to be developed. On the other hand, considering the interaction between liquid compounds for hypergolic reaction, it is important to give special attention

contact area between the hypergolic system components.

REFERENCES

Blackburn, J., 2006, “Gas generating system with autoignition device”, World Intellectual Property Organization, WO2006/105412 A2.

Brown, W.S. et al., 2010, “Low velocity injector manifold for hypergolic rocket engine”, U.S. Patents 2010/0037590 A1.

DiSalvo, R., 2012, “High energy, low temperature gelled bi-propellant formulation preparation method”, U.S. Patents 2012/0073713 A1.

Fawls, C.J. et al., 2005, “Propellants and explosives with

U.S. Patents 6,843,868 B1.

Fisher, S.C., 2009, “Coaxial ignition assembly”, U.S. Patents 2009/0320447 A1.

Hallit, R.E.A. and Bauerle, G., 2004, “Hypergolic azide fuels with hydrogen peroxide”, U.S. Patents 2004/0221933 A1.

Hawkins, T.W. et al., 2011, “Hypergolic fuels”, U.S. Patents 8,034,202 B1.

Koppes, W.M. et al., 2010, “Reagents for hypergolic ignition of nitroarenes”, U.S. Patents 7,648,602 B1.

Natan, B. et al., 2011, “Hypergolic ignition system for gelled rocket propellant”, World Intellectual Property Organization, WO2011/001435 A1.

Sengupta, D., 2008, “High performance, low toxicity hypergolic fuel”, U.S. Patents 2008/0202655 A1.

Smith, J.R. et al., 2010, “Hydroxyethylhydrazinium nitrate-acetone formulations and methods of making hydroxyethylhydrazinium nitrate-acetone formulations”, U.S. Patents 2010/0287824 A1.

Stevenson, III H.W. et al., 2011, “Hypergolic liquid or gel fuel mixtures”, U.S. Patents 2011/0272071 A1.

Thuman, C. et al., 2011, “Method for combating explosive-charged weapon units, and projectile designed for the same”, World Intellectual Property Organization, WO2011/053211 A1.

Watkins, W.B., 2004, “Hypergolic fuel system”, U.S. Patents US2004/0177604 A1.

Silva, G., Iha, K.


INTRODUCTION

Both stable and radioactive isotopes have many important applications in the aerospace area. Nuclear propulsion, based in very compact and highly enriched 235U demanding nuclear reactors, has been pointed as having great potential for deep space navigation (Bennet, 2006). Electricity generation in spacecrafts that travel far from the sun are in general based on radioisotope thermoelectric generator (Bennet, 2006; Flicker et al., 1964). Inertial sensors (accelerometers and gyrometers)

be produced by combining different Si isotopes (Kato and Lamont, 1977). Lithium niobate is an optical material broadly used in electro-optics devices and circuits used in inertial optical platforms: lithium niobate with low content of 6Li is shown to be less sensitive to cosmic radiation and to have a longer operational life than the ordinary LiNb (Riley, 1999). Magnetic and magneto-optical sensors responsivity can be improved if isotope contents are considered (Itoh et al., 1999;Kamada et al., 2009).

Isotopes are separated in many different ways, depending on

and application. The most used isotope separation methods are those based on centrifuges (Beams and Haynes, 1936; Kholpanov et al., 1997), gas diffusion (Naylor and Backer, 1955), electromagnetic methods (Martynenko, 2009), thermal diffusion (Furry et al., 1939; Rutherford, 1986), aerodynamic method (Becker et al., 1967), ion exchange and chemical separation (Calusaru and Murgulescu, 1976; Kim et al., 2001), plasma centrifuge (Prasad and Krishnan, 1987; Del Bosco et al., 1987), ion cyclotron resonance – ICR (Dolgolenko and Muromkin, 2009; Louvet, 1995), atomic vapor laser isotope separation – AVLIS (Schwab et al., 1998; Paisner, 1988), and molecular LIS – MLIS (Schwab et al., 1998; Jensen et al., 1982). The main differences between two distinct isotopes are mass, nuclear volume and nuclear spin. Most of the isotope separation processes are based on mass difference, however, the methods based on lasers, generally called LIS methods rely on the subtle difference on electromagnetic radiation absorption spectra (Mack and Arroe, 1956). The Institute for Advanced Studies (IEAv) has studied isotope separation, both in MLIS and AVLIS, mainly in uranium enrichment for nuclear fuel production (Schwab et al., 1998).

doi: 10.5028/jatm.2012.04041712

Generation of an Atomic Beam by Using Laser Ablation for Isotope Separation PurposesJuliana Barranco de Matos1, Márcio de Lima Oliveira1, Emmanuela Melo de Andrade Sternberg1, Marcelo Geraldo Destro2, Rudimar Riva2, Nicolau André Silveira Rodrigues2*1Instituto Tecnológico de Aeronáutica – São José dos Campos/SP – Brazil2Instituto de Estudos Avançados – São José dos Campos/SP – Brazil

Abstract: Atomic vapor laser isotope separation has been studied at the Institute for Advanced Studies for nuclear purposes since 1982, and recently it has been questioned about its potentialities for the aerospace area. Many applica-tions from nuclear propulsion to electricity generation and space navigation have been found, which justify the study of isotope separation for aerospace applications. ne of the ey process, and the rst step for atomic vapor laser isotope separation, is the production of a neutral vapor jet. This paper discussed the potentiality of using laser ablation as a tool to generate neutral metal vapor jet for isotope separation purposes. The basis for the discussion is a set of experimental results obtained at the Institute for Advanced Studies. The experiments were described, the results were analyzed using basic theoretical treatment found in the literature, and it was concluded that laser ablation is a potential tool for the generation of a neutral vapor jet for atomic vapor laser isotope separation purposes.

Keywords: Laser ablation, Laser isotope separation, Neutral jet generation.

Received: 02/05/12 Accepted: 22/08/12*author for correspondence: [email protected] Coronel Aviador José Alberto Albano do Amarante, 1 – PutimCEP 12.228-001 São José dos Campos/SP – Brazil


The isotope separation based on AVLIS process follows three steps: production of a neutral atomic beam; selective photoionization of the desired isotope, and collection of the photoionized atoms.

simple resistive heating can cause evaporation in low-melting temperature materials, whereas, for high melting temperature materials, electron beam heating is generally used (Schiller et al., 1983). However, there are refractory materials for which even electron beams cannot produce the desired vapor. Short-laser pulses can remove a fraction of a target surface, generating a plume made of neutral atoms, ions, clusters, and

et al., 1999; Capitelli et al., 2004; Noll, 2012). Such process is called laser ablation. Near the target surface, both atoms and ions are in excited states and tend to decay to the ground or metastable state as the plume expands (Kools et al., 1992). As far as the neutral fraction can be separated from the rest of the plume, laser ablation could be used as a neutral jet source for AVLIS. All the references mentioned in the last paragraph deal with the behavior of the plasma in the target vicinity. However, the interest for isotope separation is far from the target surface, where the plume is not collisional anymore and most of the ions and neutral atoms have decayed to the ground or metastable states. This paper presents an experimental study in order to investigate the assumption that laser ablation can be used to prepare a neutral atomic jet for AVLIS purposes. Firstly, laser ablation is discussed in the thermal regime, and the relevant parameters are presented. The experiments are described and the results are analyzed using basic theoretical models found in the literature. It is concluded that, under our experimental conditions, laser ablation is indeed a potential method for neutral jet production for further isotope separation, at least when tiny amounts of material are desired.

LASER ABLATION

Laser ablation is the general designation for the material removal, from a solid or liquid surface, by short laser pulses. In this paper, only the so-called thermal laser ablation will be considered since it corresponds to the present experimental conditions (Amoruso et al., 1999). If a laser pulse illuminates an area A of a solid target, a fraction aA of the pulse will be absorbed. If the target is metallic or a strong absorber, the laser energy will be absorbed in a very thin layer with the thickness given by the optical penetration length. For instance,

for metals illuminated by light in the visible, the optical penetration length is typically much smaller than the radiation wavelength (Born and Wolf, 2002). It is reasonable to suppose that all the laser pulse energy is absorbed in the target surface and transmitted to the target volume through heat conduction. During the pulse duration , the heat penetrates the sample a depth given by the thermal diffusion length (Eq. 1):

,L T4D κ= (1)

where,

Usually, in thermal ablation experimental conditions, the diameter of the illuminated area is much larger then LD,

solid uniformly illuminated’ problem and the target surface temperature, at the end of the laser pulse, will be as in Eq. 2 (Duley, 1976):

( ) ,T TKa I

02 /

A0

01 2

rlx= + ` j (2)

where,T0: is the target temperature before the laser pulse,I0: is the laser intensity (power/illuminated area), andK: is the thermal conductivity.

The mass mE removed from one single laser pulse can be estimated by using calorimetry and neglecting the solid-liquid phase transformation and the temperature dependence on the

cE as in Eq. 3:

,mc T L

aE

E V

A pf

D=

+ (3)

where,

P: is the pulse energy,T: is the temperature variation from the room until the

boiling temperature, andLV: is the vaporization enthalpy.

EXPERIMENTS

Figure 1 presents the basic experimental setup: the laser beam is focused on the sample surface, forming a 45º angle by

to a computer controlled xy table, and the sensor is placed in the plume path, at different distances from the sample.

Matos, J.B. et al.


These experiments were performed in vacuum (~10-5 mbar), and three different sensors were used: a mass spectrometer, a

electrostatic probe. Three different lasers were used with the characteristics given in the Table 1. A set of experiments of emission spectroscopy was accom-plished in air. In such cases, a fraction of the light emitted by the plume was collected by a quartz lens, coupled to 500 um

model HR4000 and later analyzed.

Mass spectrometry

In this set of experiments, the measuring device was a Pfeiffer quadrupole mass spectrometer model QMS-300, placed either at 10 or 25 cm from the target, with the ion collector placed orthogonally to the plume pathway. The lasers used in these experiments were the CuHBr and the CVL. Stainless steel, nickel, copper, tungsten, and tantalum targets

mass spectrometer ionization sector turned off and the second with the ionization sector turned on. With the ionization off, the ions captured by the mass spectrometer were only those produced in the laser ablation. With the ionization sector on, ions produced by impact of neutral atoms with electrons were added to those produced by laser ablation. For the stainless steel sample and the ionization sector turned off, peaks for Fe and Cr were observed, the main components of steel, and only singly ionized single atoms were present in the spectrograms, or rather, there were no peaks corresponding to ions doubly or highly ionized, not even for particles with the double of the unitary mass. For the ionization sector turned on, the spectra remained relatively the same, only the amplitude increased by a factor of about two. Although it is not possible to obtain quantitative information from this factor, because it is not clear which fraction of the neutral atoms are ionized in the ionization sector, it indicates that the ion and neutral populations have roughly the same order of magnitude. The same behavior was observed with all the remaining samples: only peaks due to single atoms singly ionized were observed. It is known that the plume generated by laser ablation is always followed by droplets; however, they were not seen in the mass spectrograms because the droplet mass was much above the mass spectrometer upper limit (300 amu). Thus, in our experimental conditions, except for the droplets, the monitored plume was mainly made of single atoms (neutral or singly ionized).

PVDF sensor

The PVDF is a polymer that exhibits pyro and piezoelectric

very convenient to measure the plume drift velocity (center of mass velocity) and translational temperature both for neutral or ionized atoms. It is well-accepted that the plume generated by laser ablation of single element targets, far from the target surface, is mainly made of a bunch of atoms, which expand according to a maxwellian velocity distribution with a drift

et al., 2004):

.expS tt k T

mt

v1

2l

B Z5 0

2

- -^ `h j; E (4)

possible to obtain the drift velocity v0 and the translational temperature Tz. These experiments were performed with a

Table 1. Laser parametersParameter CVL CuHBrWavelength (nm) 511/578 355Pulse width (ns) 40 35 25Repetition rate (kHz) 5 16 2Pulse energy (mJ)* 2.5 1 0.23Peak power (kW)* 60 29 9.2Average power (W)* 12.5 16 0.46Illuminated area (cm2) 1.1 × 10-4 1.5 × 10-5 6.3 × 10-6

Beam quality – M2 16.7 6 1.7Peak power density (W/cm2)** 5.3 × 108 2.0 × 109 1.5 × 109

Fluency (J/cm2) 22 66 37

* maximum values; ** maximum value at target surface.


HyBrID-copper laser and Fig. 2 provides a typical PVDF

sample; the solid lines indicate the PVDF signal and the

Electrostatic probe

The electrostatic probe is convenient to measure drift velocity, translational temperature of ions, ion density, electron

placed transversely to the plume, at distances ranging from 4 to 20 cm from the target, was used to study the plasma generated

ions was studied with the probe polarized at a -10 V voltage. Later, the electrostatic probe was used to evaluate the charge densities and electron temperature. Figure 3 shows the probe current signal time behavior for different probe electric potentials in experiments with the CuHBr laser and W targets. The noise around 5 × 10-6

s is from laser electric discharge pulses, and it was taken as reference for triggering the oscilloscope. The peak values of the probe current curves were plotted against the polarization voltage, giving rise to the Langmuir curve. From this curve, the charge densities

and electron temperature were evaluated, considering the hydrodynamic expansion approach (Koopman, 1971). This experiment was repeated with aluminum and tungsten targets, ablated by CVL laser, and for copper samples, ablated by the

Figure 3. Temporal behavior of the probe current in different voltages, for ablation of tungsten samples with the CuHBr laser.

Table 3. Plasma parameters for ablation of different targets and lasers, measured with the electrostatic probe. All the listed

Target Copper Aluminum Tungsten Tungsten

Laser CVL CVL CuHBr CVL

Ion density (m-3) 3.4×1016 2.3×1015 2.4×1017 1.3×1016

Electron density (m-3) 1.2×1015 3.2×1014 6.3×1015 3.9×1014

Electron temperature (eV) 15 19 28 15

Drift velocity (km/s) 8 – 10 8 – 13 5.4 6 – 10

Emission spectra

Laser induced breakdown spectroscopy (LIBS) experiments in air were made with copper, graphite, molybdenum, alumina, and beach sand samples in order to investigate the composition of the expanding plume. The plume light emission in our experimental conditions vanishes for distances larger than 3 or 4 mm and thus the experiments were performed just above the target surface and not at the same distances, as in the case of the PVDF and electrostatic probe experiments.

samples, the line attribution was made by comparing

Matos, J.B. et al.

line) for ablation of tungsten with CuHBr laser.

Table 2. Plume parameters for ablation of tungsten targets measured with the PVDF sensor.

Target TungstenLaser CuHBrTranslational temperature (K) 8.8 × 104 – 9.1 × 104

Drift velocity (km/s) 4.65 – 4.74


considering only the sample constituent element. The synthetic spectrum was built by adding lorentzian curves with peak values given by the Einstein Athe National Institute of Standards and Technology transitions database (NIST, 2012), and with the linewidth equals to the instrument resolution (1.5 nm). Figure 4 compares the measured (black) and synthetic (gray) spectra for the copper sample, in 450 to 550 nm. In the case of alumina targets, besides the aluminum and the oxygen, nitrogen and sodium in the synthetic spectra were also considered. The experiments with beach sand were performed in order to examine the separation potentiality of our experimental

heat dried, pressed and sintered with the purpose of getting compact samples, which have a very complex composition and are inhomogeneous, i.e., the spectrum depends on the position the laser strikes the sample surface. It was not possible to build a synthetic spectrum because of the complexity and lack of information about the sample composition. We focused our attention to some peaks that are repetitive and very distinct from the background, by comparing their resonance wavelength with the NIST database. Several of the more intense observed lines were due to sodium and silicon, as shown in Fig. 5. With regards to Mo samples, a Jobyn-Yvon spectrometer model TRIAX 550, with a 0.025 nm resolution (at 546 nm), was also used and plasma temperature was measured by means of the Boltzmann plot method (Amoruso et al., 1999). The plasma temperature was about 0.9 eV, in agreement with results found in the literature for spectroscopic measurements (Noll, 2012;

Capitelli et al., 2004), but in contrast with the results obtained in this work with Langmuir probe for other metallic samples.

ANALYSIS

In order to establish the magnitudes this work refers to, let us consider the experiments performed with tungsten targets and the CuHBr laser. The tungsten properties are: aA = 0.493 for = 511 nm; K = 1.74 W/(cm ºC), = 0.7 cm2/s , atomic mass MW = 183.84 amu, cE = 0.133 J/(g ºC), LV = 4.48 kJ/g, and W = 19.3 g/cm3. The laser parameters are provided in Table 1. With these, Tmuch higher than the tungsten boiling point of 5,930 K (Lide, 1996), and before the surface had achieved this temperature level, a fraction of the sample had been evaporated and ejected, starting the formation of the ablation plume. This evaluation was done without taking changes of the thermal parameters with temperature into consideration, and without considering the interaction of the laser beam with the ejected plume. However, the value is in the same order of magnitude as the translational temperature measured, both with the PVDF sensor and the electrostatic probe. It means that the laser pulse energy is in some way delivered to the ejected plume. Using Eq. 3, the mass that is removed in one single pulse is estimated in mE = 1.0 × 10-7 g, which implies that, taking the laser repetition rate of 16 kHz, the ejected mass rate is about 5 g/h. This evaluation requires some care, and some fraction of the removed material expands as clusters and/or droplets. The following calculations assume that all the removed material is

must be faced as limit values, which are useful only to provide orders of magnitudes for the analyzed parameters.

Figure 5. A typical sand LIBS spectrum. The saturated peak around 355 nm is due to the scattering of the laser beam.

Figure 4. Comparison between measured (black line) and synthetic (gray line) spectra for copper samples.


If one takes an intermediate value for the expansion velocity (drift velocity) of vD = 5 × 103 m/s, at the end of the laser pulse the plume will expand a distance given by LE = vD

of volume) at the end of the laser pulse can estimate if it is assumed that all evaporated material had been expanded to a volume given by a hemisphere with radius LE as in Eq. 5:

/. .n

mm

Lcm

2 31

2 9 10w

E

E3

19 3#cr

= - (5)

The mean free path is given by Eq. 6:

,Ln

nm1

560pcc

v= (6)

where:

C: the collisional cross section was estimated taking the tungsten atomic radius (1,41 Å) and thus C

Time between two successive collisions is given by Eq. 7:

vL

ps166CT

pcx = (7),

where:vT: is the thermal velocity, calculated by taking the temperature of 100.000 K previously estimated.

duration, more than 200 collisions between particles happen in the plume. Kools et al. (1992), using Monte Carlo calculations, showed that only about four collisions are necessary for thermalization in the expanding plume. Thus, it is reasonable to consider that the plume expands similarly to a gas in equilibrium

which is suddenly released to expand into vacuum. The particle density decreases as the plume expands and, after some distance, the plume is not collisional anymore. To estimate this distance, it is considered that the particle density decays with the distance from the target surface according to Eq. 8,

( )n z n L zL

EE

3

= ^ `h j (8)

and that the plume stops being collisional when the mean free path is in the same order of magnitude as the plume size LP, thus substituting Eq. 8 into 6, one has Eq. 9:

. .L n L L mm3 1NC E C E3v= =^ h (9)

If one considers a 5 km/s expanding velocity, the time to expand until LNC is in the order of 600 ns. Therefore, for the experimental conditions presented in this work, after an expansion of about 3 mm, the plume is not collisional anymore and the interaction between particles from this point on is essentially electrostatic. Basically, two kinds of behavior are expected, for low densities the charged particles behave like free particles and for high densities they

allows identifying the particles’ behavior is the Debye Length LDb, given by Eq. 10:

.Lq Nk T /

Dbe

B e2

01 2f

= c m (10)

Taking the electron densities and temperatures from Table 3, the Debye length ranges from 0.5 to 2.6 mm in the present experimental conditions. Therefore, the plume typical dimensions are in the same order of magnitude as the Debye length and the particles’ behavior is in the transition between the individual particles and plasma behavior regimes. It suggests that the charged particles (ions mainly) can be separated by

from distances in the range of few millimeters and that the remaining plume fraction will be made of single atoms. The same calculation was repeated for copper and aluminum, and the results, together with the values for

and the same comments made for tungsten are also applicable for copper and aluminum.

Table 4. Estimated plume parameters for tungsten, copper, and aluminum, using the same calculation procedure described in “Analysis”. The material constants were taken from Lide (1996), laser parameters from the CVL laser in Table 1 and plasma parameters from Table 3.

Parameters Tungsten Copper AluminumTemperature at the surface (K)* 1.0×105 1.0×104 1.4×104

Removed mass per pulse(kg/pulse) 1.0×10-10 1.6×10-11 6.7×10-12

Atom density (m-3)* 2.9×1025 1.3×1025 1.3×1025

Mean free path (m) 5.6×10-7 1.5×10-6 1.2×10-6

LNC (m) 3.1×10-3 1.9×10-6 2.1×10-6

Debye length (m) 5.0×10-4 8.0×10-4 1.8×10-3

* at the end of the laser pulse.

Matos, J.B. et al.


CONCLUSIONS

In this work, several materials were evaluated in different experiments of laser ablation, using low energy (~ mJ per pulse), high repetition rate (~ tens of kHz) lasers in the visible and in the near ultraviolet, with pulse width in the range of tens of nanoseconds. The ablation experiments were in the thermal regime, with energy density in the range of tens of J/cm2 and intensities of about 109 W/cm2. The set of results for these experimental conditions leads to the following conclusions:

ablation is made of single atoms (neutral or ionized), even if complex targets are used;

magnitude;

longer collisional;

length is such that the charged fraction of the plume can be

In short, it is possible, using laser ablation, to generate an atomic beam adequate for AVLIS purpose. This is possible even for very complex targets, such as ores. The main limitation is the small amount of material that is removed, limiting the method for the separation of small amounts of material. This is a severe limitation for the separation of materials that are needed in large amounts, such as uranium, however it is adequate for the separation of materials used in photonics or in magneto-optic sensors, which require small amounts of isotopes.

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Noll, R., 2012, “Laser-induced breakdown spectroscopy”, Springer-Verlag, Heidelberg, 543p. doi:10.1007/978-3-642-20668-9

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pp. 253-260. doi:10.1007/BF00692883

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Matos, J.B. et al.


INTRODUCTION

In the aerospace industry, the two routes of manufacturing technologies for structures have been constantly improved. One of them is the employment of polymer matrix composite materials in aircraft structures, which has been growing over the years (Mangalgiri, 1999). The other one is the use of conventional metallic materials with enhanced mechanical and physical properties (King et al., 2009). This later could be considered as a safer route due to the very large experience of metallic alloys engineering use. The rising competition between composites and metals took the aluminum alloys producers to develop, jointly with aircraft manufacturers, lighter alloys, with high mechanical strength and high damage tolerance. The aluminum-copper-lithium alloy AA2198 is an example of these new generation

alloys. Typically, the AA2198 alloy composition is 3.2% Cu and 1.0% Li, falling in the Al solid solution above 500 °C. Aging at lower temperatures promotes the formation of intermetallics responsible by strengthening effect (Bordesoules, 2007). In addition to these new alloys, materials-joint techniques have been improved aiming at the reduction of weight, costs, and lead-time. Although the riveting process is highly automated, which is largely used by aircraft manufacturers, this process reached its development potential limit, and no

can be expected. Thus, many joint techniques that could

industry, adding low weight and mechanical properties suitable with structural demands during aircraft lifetime operation. Among the available welding processes, friction stir welding (FSW) and laser beam welding (LBW) have presented advances over the past years, becoming attractive for the aerospace industry worldwide.

doi: 10.5028/jatm.2012.04044212

Occurrence of Defects in Laser Beam Welded Al-Cu-Li Sheets

1 1,2*1Instituto Tecnológico de Aeronáutica – São José dos Campos/SP – Brazil2Instituto de Estudos Avançados – São José dos Campos/SP – Brazil

Abstract: In the aerospace industry, laser beam welding has been considered as one of the most promising routes among the new manufacturing processes. Substitution of riveting by laser beam welding of aircraft structures has contributed to weight and cost savings. Concurrently, new aluminum alloys have been developed with the addition of lithium with better mechanical properties and lower density. The Al-3.5%Cu-1.1%Li alloy (AA2198) is one of these new generation alloys. However, laser beam welding of Al-alloys expectations might be greatly reduced by the occurrence of two main defects:

On the other hand, hot cracking happens due to the conjunction of tensile stresses, which are transmitted to the mushy

the weld beads presented high porosity level, but with a decreasing tendency when welding from both sides. The use of the

Keywords: Laser, Laser beam welding, Aluminum alloys, Aerospace.

Received: 11/07/12 Accepted: 04/09/12

Trevo Coronel Aviador José Alberto Albano do Amarante, 1 – PutimCEP 12.228-001 São José dos Campos/SP – Brazil


LBW already found its place in industrial production, with a market in full expansion. Initially, the automotive industry developed LBW processes for sheets of different thicknesses prior to forming, which are called tailored-blanks welding, and then body-in-white laser welding. The portfolio of available

alloys, presenting productivity and welding quality gains (Pallett and Lark, 2001). Afterwards, the aerospace structure manufacturers began to substitute some riveted panels by laser-welded ones. The utilization of laser as a welding process allows not only weight reductions, but also the development of optimized panels, decreasing the manufacturing lead-time. If compared to the typical riveting speeds (200 to 400 mm/min), the laser welding clearly shows itself more productive, reaching speeds exceeding 6 m/min. Even with more rigorous inspections and

become less susceptible to corrosion, since holes in the skin are avoided and gaps in butt joints are eliminated (Rötzer, 2007). In spite of these advantages, LBW like other fusion methods is subject to common metallurgical problems, such as hot cracking and porosities. The control of welding defects is of utmost importance to control the mechanical properties under the extreme conditions aerospace materials are subjected to. Two of the major problems in fusion welding of aluminum alloys are related to porosity and hot cracks. These problems will be further analyzed. Porosities are intrinsically related to the weld, occurring due to a large number of factors: alloy composition, surface contaminants, improper gas shielding, keyhole collapse, and hydrogen release. In case of alloys with highly volatile elements, such as Li in the present case, boiling could also happen. Pores in Al-Li welds beads had been previously reported (ASM, 1993) as a result of hydrogen contamination leading to interdendritic microporosity.

metallic alloys. This is the result of inadequate melt feeding initiating micropores and severe deformation leading to the opening and propagation of such defects. This type of

solid fraction is high (Piwonka and Flemmings, 1966). A large

dendrite roots increasing the tendency to hot cracking.

and Davies (1981) is formulated as the ratio between the vulnerable time period (tv), and the time available for stress-

relief process (tR), i.e. the time spent in the interdendritic

fraction. Equation 1 presents the hot cracking susceptibility (HCS) followed by Clyne and Davies (1981).

HCStt

t tt t

R

V

90 40

99 90= =--

(1)

Since the solid fraction is a function of temperature and

is a suitable way to decrease the vulnerable time. This is usually accomplished when an eutectic forming compound is added to the weld. For example, it is well known that silicon reduces tv

reduced and the fraction of the eutectic phase is increased. Drezet et al. (2008) also proposed that hot cracking could be diminished when two laser sources are used together. The main

the middle of the weld bead, so the liquid permeability increases and the thermal gradient decreases. It has been proved that a process using two laser sources improves the high temperature toughness of the AA6013 aluminum joints (Lima et al., 2001). The use of two laser sources could be unpractical in some weld geometries, but using two weld runs could reduce the

of an aircraft panel, two runs could be envisaged: one at the joint between skin and stringer and another in the opposite face. For this, two challenges must be attained: the laser beam must be very accurately positioned at the interface between

inserted in some way it is not obstructing the beam. Therefore,

this ribbon directly at the joint intersection. This work intended at contributing to the study of

experimental results of the weldability are missing in the

produce a new insight on the matter. The weld geometry is similar to that of a stringer-skin T-joint both autogenous and

MATERIALS AND METHODS

A 1.6 mm thick aluminum alloy AA2198-T851 sheet was utilized in this work. Its composition is shown in Table 1. The sheet was cut in coupons with 30 x 100 mm dimensions. For

Higashi, A.L.C., Lima, M.S.F.


utilized and its composition is shown in Table 2. Wires of 1.0 mm in diameter were cold rolled to ribbons of 1.6 mm of width and 0.2 mm thickness. The ribbons were placed between the sheets

as depicted in Fig. 1. Based on previous studies, the angle

Co. (USA) was used. The laser radiation is generated in a

is then connected to an Optoskand processing head. The focal length was 157 mm with a minimum spot diameter at the focus of 100 mm.

the surface against oxidation. The protection gas was delivered through a rounded copper tube of 2 mm internal diameter directly over the irradiated area (Fig. 1). A computer numerical control (CNC) table carries out the sample movement. Right before welding, the sample surface was grounded with a

SiC 600 paper to remove oxidation and then washed with distilled water and ethanol. Microstructural analyses were carried out using optical microscopy (OM) and scanning electron microscopy (SEM).

equipped with acquisition system and image processing

The equipment for mechanical tests was an MTS 810 tensile machine (USA) having a loading cell with 250 kN capacity. The mechanical testing was carried out in T-pull mode, as described in Fig. 2. The load is realized by pulling the stringer at constant speed of 1.0 mm/min. The sample dimensions for T-pull testing were: 3(W)x3(H)x2(L) cm. In order to understand the thermal and mechanical

As the physical properties of AlCuLi alloy are missing in the literature, the simulation had been performed using the constants of an AlSiMg alloy class AA6061. One and two-sided welding conditions were simulated in autogenous condition. In order to reproduce experimental conditions, the simulated sheets were rigidly attached to the borders, and the time between the start of each weld, for the two sides welds, was ten seconds. The mechanical formalism is based on Von Mises strain and stresses (Wikipedia, 2012).

bench. The white block is an alumina calibrated support, and the rounded nozzle is responsible for gas shielding.

stringer

skin

F

×

Figure 2. T-pull mechanical testing schematics. The skin part of the

The force (F) is applied parallel to the stringer direction.

Table 1. AA2198 alloy composition in weight percent (Al as the balance).

Element Cu Mg Li Ag Zr Mn Si Zi Ti Fe Other% wt. 3.50 0.80 1.10 0.50 0.18 0.50 0.08 0.35 0.10 0.10 0.15

Table 2. AA4047 alloy composition in weight percent.Element Al Cu Mg Mn Si Fe% wt. 87.83 0.0015 0.001 0.01 11.89 0.252


After a number of free trials, some experimental conditions have been retained for a detailed study. Table 3 presents the experimental conditions for the welds, where the run could be one or two depending if one or both sides were exposed to the beam. The heat input is the ratio of power per speed for each run.

RESULTS AND DISCUSSION

In the present work, the parameters affecting the heat input provided by the laser, speed (v) and power (P) were studied. In general, the combination of P and v of the welds generated weld beads of reasonably similar dimensions.

(Fig. 3a) or two sides (Fig. 3b), however the welded zones were always asymmetric. All samples presented epitaxial growth of grains from the base/molten metal interface

the base material promoted the coarser dendrite formation with columnar structure. Near the bead top, where cooling rate was higher, the grains had not preferential orientation

As can be seen in Fig. 3, a great quantity of pores is presented in all weld beads. Some of them presented large toes (Figs. 3a and 3b), which can be explained by the large number of pores within them. Microporosity as much as macro-porosity were present in the welded zone, indicating one or more mechanisms

the welds, however the volume fraction of pores does not have statistical meaning because of large variations in their density from one cross-section to the other. These pores are mainly linked to the lithium degassing during melting and are frequently associated with poor weldability of Al-Li alloys. The welds with both side seams, as presented in Fig. 3b, showed smaller pores, indicating that the Li vapor had more time to leave the molten pool. Other possible sources of porosity to be considered are surface and gas contaminants. The current careful control

contaminant, thus having a minor role in the porosity. Additionally, other types of aluminum alloys had been welded in the same experimental conditions (Siqueira et al., 2012), including use of the bench vise and the gas nozzle as presented in Fig. 1. Usually, these welds present only few small pores. Therefore, it is much likely that pores are due to Li degassing and, to an unknown extent, to hydrogen nucleation (ASM, 1993). Figure 4 presents a closer look of a separate pore cross-section using secondary scanning electron microscopy. It

(a) (b)Figure 3. Optical micrographic images of the weld seams. (a): one side beam, condition: one run – 1,400 W/3 m/min. (b): two runs – 1,200

W/4 m/min.


Table 3. Process parameters.Power(W)

Speed(m/min.)

Heat input(J/mm) Condition

1,200 2 36 Autogenous/two runs1,200 2 36 Autogenous/one run1,200 2 361,200 2 361,400 3 28 Autogenous/one run1,400 3 281,400 3 281,200 4 18


can be seen that the porosity is perfectly spherical with inner

with epitaxial growth from the disc to the surroundings. This observation indicates that pores inoculate the liquid with

Fig. 5a it could be seen a pore in the middle of an isolated grain (arrow). Page and Sear (2006) also showed that pores are preferred sites for heterogeneous nucleation of new phases. The same mechanism of metal nucleation around pores is also observed in metal foams. Duarte and Banhart

in the foam-processed aluminum alloys classes AlSi7 and

weld seams toughness, particularly in high temperatures in which hot cracks appears. However, the large pores observed here act as stress concentrators and probably hide the positive

in the fusion zone. The welds performed with laser power of

1,200 W and 2 m/min speed did not present hot cracking with

respectively. Thus, the HCS was reduced to this condition

current theories (Campbell, 2003), the chemical composition is

possible to accurately measure the actual composition of the welded zone using the energy-dispersive X-ray spectroscope of the scanning electron microscope (SEM-EDS), since many alloying elements were well-below 1 weight percent and the second most important alloying element (Li) was too light to be detected. Nevertheless, semiquantitative chemical analyses were performed, and are presented in Fig. 6 for two samples with Al-Si additions. It could be seen that silicon distribution is approximately homogeneous in all areas, but at the bottom region, next to the skin, called “2” in Fig. 6. The accumulation of Si in these regions could be explained due to the absence of

Through SEM-EDS imaging analyses, the composition of two side-welded beads was obtained. The content of silicon

eutectic Al-12% Si) was diluted in the bead during welding.

dilute out over the entire bead during welding.

composition on hot cracking is to compare the ratio of vulnerable to stress-relief times, as presented in Eq. 1. With regards to the same cooling conditions, one could compare the temperature interval, related to tv and tR, between an alloy composed Al-2.9%Cu-1.1%Si (Fig. 6) and another with Al-3.5%Cu. Thermocalc (1994) computations provided the results presented in Table 4, and as can be seen the HCS drops

one run.

(a) (b)

cracking, condition of 1,200 W/2 m/min.


on the actual melt composition and cooling conditions.

intervals (Table 4) and the mechanical strains during the

thermomechanical evolution during welding by computer

conditions to the experimental setup. One-side welding temperature and mechanical response were simulated with power of 1,400 W and a speed of 3 m/min. For the simulation of two-side weld, the chosen parameters were power of 1,200 W and speed of 4 m/min. The current thermal inputs were 28 and 18 J/mm per run (Table 3), for one and two-side welds, respectively. These simulation parameters were similar to those experimentally observed in Fig. 3. Figure 7 presents data plots as a function of processing time. For the two-side welds, the second curve begins at ten seconds because the second run started at this time.

quite similar at the beginning of the welding process. The second run, for the two-side weld, produces a second peak, which after ten seconds attains about 580 °C, approximately the solidusIndeed, the second run promotes a melt depth up to the oppo-site surface as shown in Fig. 3b. Since two melting periods are expected, the liquid had additional time for Li degassing in comparison to the one side run. Therefore, less porosity was obtained with two runs.


Condition 1.4 kW x 3 m/min. Condition 1.2 kW x 2 m/min.

Region % Al % Cu % Si

1 96.1 2.9 0.9

2 96.3 2.3 1.4

3 96.0 3.0 1.0

4 95.8 3.3 0.9

5 96.1 2.8 1.1

Average 96.1 2.9 1.0

Region % Al % Cu % Si

1 96.0 2.7 1.3

2 94.9 2.8 2.3

3 95.9 3.1 1.0

4 96.3 2.8 0.9

5 96.2 3.2 0.8

Average 95.9 2.9 1.2

picture indicate the chemical composition for each region.

Table 4. Calculation of temperature intervals in different compositions. T(fs) means temperature in Kelvin at a given solid fraction. HCS: hot cracking susceptibility.

Temperature(K) Al-2.9%Cu-1.1%Si

Welding without

T (fs=99%) 837.04 857.04T (fs=90%) 869.93 887.01T (fs=40%) 912.57 919.51

HCS(Equation 1) 0.77 0.92

HCS: hot cracking susceptibility.


The displacement (Fig. 7b) represents the shift in posi-tion during welding of a point at the centerline exactly at the interface between the sheets. The measurement position is

to the bench (Fig. 1), these movements were highly constrained leading to residual stresses. The rigid clamping had therefore

7d. The calculated strain during welding attained 5 x 10-5 for the -5 for the second. The most important

feature for cracking is the strain rate. A very high strain rate creates porosities at the root of dendrites, thus developing hot cracking (Rappaz et al., 1999). The value attained at the second run was 0.02 s-1. This value is very low and considered safe, at least for the AA6061 aluminum alloy (Drezet et al., 2008). The effect of different weld procedures on the mechanical stresses is presented in Fig. 7d. The low heat input of the two-side method compared to the one-side allowed a lower level of residual stress up to ten seconds. The residual stresses at ten seconds were 0.09 and 0.07 MPa, respectively. After the second run, the difference was even larger, 0.09 and 0.04 MPa.

These stress levels are very low compared to the elastic proper-ties of aluminum alloys and thus the distortion should be very small. Indeed, the T-sets did not show distortions after welding. All these simulation results had been developed using an

considered. Therefore, the results must be considered only in a qualitative way. Notwithstanding these results, it could be estimated that the T-joint with better chances to be used in

to understand if the observed massive porosity produces an unsuitable weld from the mechanical point of view. The mechanical characterizations of the welds were presented in Figs. 8 and 9. For clarity reasons, the stress is presented in logarithm scale. Figure 8 presents a direct strain-

T-joint, when welded from one side to the other. As can be seen, the curves were very similar with a plateau up to 3.2 mm

low stresses. It is easy to see in Fig. 2 that the skin sheet will bend creating a three-point load scheme at the beginning of

0100200300400500600700800900

1000

0 5 10 15 20time (s)

tem

pera

ture

(°C

)one sidetwo sides

one sidetwo sides

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0 5 10 15 20

time (s)

disp

lace

men

ts (1

0-3 c

m)

(a) (b)

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0 5 10 15 20

time (s)

Stra

in (1

0-3)

one sidetwo sides

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0 5 10 15 20

time (s)

stre

ss (M

Pa)

one sidetwo sides

(c) (d)

two sheets. (c) Von Mises strain. (d) Von Mises stresses.


the mechanical testing. The tensile stress and maximum elon-gation seems to be approximately the same, regardless the use

The mechanical behavior was completely different when welding from both sides (Fig. 9). Compared to the two-side

welding increased the tensile strength from 19 to 178 MPa, and the total elongation from 4.1 to 6.3 mm. The increased toughness, more than ten times, had been linked to the chemi-cal changing of the liquid bath, since the thermomechanical behavior (Fig. 7) was about the same.

0.1

1

10

100

1000

0 1 2 3 4 5 6 7

elongation (mm)

Str

ess

(MP

a)

autogeneousfiller

Figure 9. Comparison of the mechanical behavior between an

sides welded (two runs), P=1,200 W, v=2 m/min.

It is worthwhile to compare the best result obtained in the present work with the two cases. Firstly, the AA2198 sheet without welding as the maximum attainable value. Secondly,

another aerospace alloy, AA6013, welded in similar T-joint conditions, welded one-side and autogenously. These results are presented in Fig. 10.

700

600

500

400

300

200

100

00 2

(b)

(c)

(a)

4 6

elongation (mm)

Stre

ss (M

Pa)

8 10 12

Figure 10. Comparison of the tensile mechanical behavior.

Conditions: (a) unwelded AA2198 sheet (maximum attainable condition); (b) Welded on both sides (two runs), P=1,200 W, v=2 m/min; (c) AA6013 aluminum alloy autogenously welded on one side.

The AA2198 welded coupons presented lower tensile strength and total elongation in comparison to the AA2198 unwelded coupon. This is due to the stress concentrator factor caused by the weld bead. Comparing the best results obtained in T-joint welds for AA2198 and AA6013, it could be seen that the tensile strength was much higher in the first case. The AA2198 welded coupon attained 178 MPa, compared to only 46 MPa of the AA6013 case. On the other hand, the total elongations were 9.2 from AA6013 and 2.8 mm for AA2198, indicating a hardening effect of the filler material in the present case.

CONCLUSIONS

Even with a careful control of surface preparation, all the AA2198 T-joint welds presented pores, which were linked to the degassing of Li during melting.

parts to be joined could solve the hot cracking problem. The decrease of the vulnerable to stress relief time during

susceptibility for hot cracking. The results from thermomechanical and chemical analyses, and tensile T-pull strength testing indicated that welded by


0.1

1

10

100

1000

0 1 2 3 4 5 6 7

elongation (mm)

Stre

ss (M

Pa)

autogeneousfiller

Figure 8. Comparison of the mechanical behavior between an autogeneous and filler-added welded. Conditions: one-side welded (one run), P=1,200 W, v=2 m/min.


tougher joints. The welds at 2 m/min and 1,200 W under these conditions were showing most promising properties, even in comparison to T-joined autogeneous AA6013 alloy. Because of high pore density, it is safer to consider less critical applications than the aerospace one. Depending on other results, such as fatigue behavior, the AA2198 welded parts could be used, for example, in land transportation systems.

ACKNOWLEDGMENTS

The authors thank EMBRAER for providing the aluminum sheets, Financiadora de Fundos e Projetos (FINEP) and

(FAPESP) for partial funding.

ASM – American Society of Materials, 1993, “Metals Handbook – Volume 6: Welding, Brazing, and Soldering”, 10nd ed., Metals Park (Ohio), ASM International, pp. 1392-1393.

Bordesoules, I. et al., 2007, “Trends in developments of Aluminium solutions for aerospace applications”, In: Proceedings of the European Workshop on Short Distance Welding Concepts for Airframes frames – WEL-AIR, Hamburg, 13-15 june 2007, CD-Rom.

Campbell, J., 2003, “Castings”, 2nd ed., Oxford: Elsevier Pergamon, 332p.

in Aluminium-Magnesium Alloys”, The British Foundryman, Vol. 68, pp. 238-254.

Drezet, J.M. et al., 2008, “Crack-free aluminium alloy welds using a twin laser process, In: 61st International Conference

reliability of welded components in energy and processing

Duarte, I. and Banhart, J., 2000, “A study of aluminium foam formation-kinetics and microstructure”, Acta Materialia, Vol. 48, pp. 2349-2362.

King, D. et al., 2009, “Advanced aerospace materials: past, present and future, Aviation and The Environment”, Vol. 3, pp. 22-27.

Lima, M.S.F. et al., 2000, “Advanced laser welding process”, European patent no. 01810986.8.

Mangalgiri, P.D., 1999, “Composite materials for aerospace applications”, Bulletin of Materials Science, Vol. 22, pp. 657-664.

Page, A.J. and Sear, R.P., 2006, “Heterogeneous Nucleation in and out of Pores”, Physical Review Letters, Vol. 97, pp. 065701-1-065701-4.

Pallett, R.J. and Lark, R.J., 2001, “The use of tailored blanks in the manufacture of constuction components”, Journal of Materials Processing Technology, Vol. 117, pp. 249-254.

Piwonka, T.S. and Flemings M.C., 1966, “Pore formation

AIME, Vol. 236, pp. 1157-1165.

Rappaz, M. et al., 1999, “A new hot tearing criterion”, Metallurgical and Materials Transactions, Vol. 30A, pp. 449-455.

Rötzer I., 2005, “Laser-beam welding maker aircraft lighter”, Fraunhofer Magazine, Vol. 1, pp. 36-37.

Siqueira, R.H.M. et al., 2012, “Microstructural and Mechanical Characterization of Laser Welded and Heat-Treated AA6013 Aluminum Alloy”, In: Proceedings of XI Brazilian MRS Meeting, CD-Rom.

ThermoCalc thermodynamic database, 1994, version J, Stockholm Royal Institute, Sweden.

Wikipedia, the free encyclopedia, 2012, Von Mises yield criterion, Retrieved in June 25, 2012, from http://en.wikipedia.org/wiki/Von_Mises_ yield_criterion.


INTRODUCTION

At the end of the 1990s, among the Brazilian sounding rockets, the VS-40 was presented as one that provides the best conditions for experiments in microgravity (Ribeiro, 1999). Space systems are complex, i.e., their behavior is governed by many distinct but interacting physical phenomena, and multidisciplinary, requiring balance among competing objectives related to safety, reliability, performance, operability, and cost (Rowell and Korte, 2003). Over time, advances in the engineering of complex systems have allowed to more quickly identify feasible solutions and exploit the synergy among the design disciplines (Rowell and Korte,

advances yet. The interactions between the design disciplines of the VS-40 were processed in a sequential order, in which those disciplines that act early in the conceptual design establish constraints on the others that follow later, leading to a concept without regarding the trade-offs that may exist between the design objectives. The plausible consequence of such sequential methodology is a suboptimal design with

respect to the entire project, promoted by low synergy between the design disciplines.

methodology that allows exploiting the synergy between its design disciplines has not been used yet for Brazilian sounding rockets. A methodology called multidisciplinary design optimization (MDO) replaces the traditional sequential methodology by synergic interactions between the design disciplines, promoting the overall gain in product’s performance, decreasing the design time (Floudas and Pardalos, 2009). Why should the VS-40 be revised? It promises the best conditions for microgravity experiments, but not widely launched yet such as the VSB-30, also a Brazilian sounding rocket, so that it could be more studied, and perhaps improved

it was not originally designed for carrying a payload with exposed canards, indicating that its design can be altered, if

of complex systems, and it may have some subsystems that could be improved regarding its next launches at Brazilian

doi: 10.5028/jatm.2012.04044412

Multidisciplinary Design Optimization of Sounding Rocket Fins Shape Using a Tool Called MDO-SONDAAlexandre Nogueira Barbosa1*, Lamartine Nogueira Frutuoso Guimarães2

1

2

Abstract:

Keywords:

Received: 31/07/12 Accepted: 08/10/12*author for correspondence: [email protected]


territory carrying the Sub-orbital SARA, a Brazilian platform for microgravity experiments. Motivated by a search for VS-40 improvements, the use of the MDO was introduced in Brazilian sounding rockets. Therefore, the objective of this paper was to provide a perspective of the MDO application in this context based on a case study of the VS-40. As case study, the shape optimization

increasing the chances of adverse effects that could lead to unstable behaviors. To perform the optimization, a computer tool called MDO-SONDA (MDO of Sounding Rockets), which was developed by Alexandre Nogueira Barbosa, was

introduced by this paper.

SOUNDING ROCKETS AND MICROGRAVITY ENVIRONMENT

Sounding rockets, such as the VS-40, are characterized by

(2001), such rockets are constituted of solid fueled motors and a payload that carries instruments to take measurements

Thus, the sounding term means taking measurements. In comparison with the VSB-30, the VS-40 bi-stage can provide a wide exposure to the microgravity environment, characterized by a condition where an object is subjected

achieved by moving in free fall, where there are no forces other than gravity acting on the object. Payloads carried by rockets achieve the microgravity environment after the burnout of the rocket when the thrust force is zero and the payload is above the atmosphere. It is assumed that the Kármán line, at 100 km above the seawater surface, might be used as a reference for microgravity

atmosphere and the outer space, from which the atmosphere becomes so thin that the drag force could be neglected.

FACTS ABOUT THE VS-40

In spite of the fact that the VS-40 provides more exposure to microgravity than the VSB-30, since the 21st century began, rather than the VS-40, the VSB-30 has been most frequently

, 2011).

for microgravity experiments with an advantage, the payload recovery operation associated with the VSB-30 is less costly

times more distant from the continent-ocean boundary than the VSB-30, demanding more autonomy for the recovery means. From 2004 to 2010, ten VSB-30 campaigns were successfully performed, three of them in the Brazilian territory

, 2011). In contrast to the VSB-30, three VS-40 campaigns has occurred so far, two of them in the Brazilian

carrying the Sharp Edge Flight Experiment (SHEFEX) II (Weihs launched at the Andøya Rocket Range in Northern Norway

compensate for the aerodynamic effects of the small canards at the payload, as can be seen in Fig. 1c (Weihs , 2008). In 1997, a recovery orbital platform called SARA for supporting short-orbital experiments in microgravity environment was proposed (Moraes and Pilchowski, 1997).

few minutes of microgravity conditions, an orbital one can provide more than ten days before reentering the Earth’s

(a) (b) (c)Sources: (a) and (b) Institute of Aeronautics and Space, Brazil;

Figure 1. VS-40 launches.


similar example of such a kind of platform (Reddy, 2007).

for application of SHEFEX derived technology, which is a reusable orbital return vehicle for experiments under microgravity conditions (Weihs , 2008b). Thereafter, a platform called Sub-orbital SARA, which is part of the road map to achieve the orbital mission purpose of this platform, has been constructed to be launched by a VS-40, supporting an experimental module to be exposed to

of the S44 motor, which constitutes the fourth stage of the

methodology had recently been presented.

as a case study using such methodology to demonstrate its application in the context of Brazilian sounding rockets. However, before presenting the results of the optimization, the main aspects of the MDO-SONDA will be further depicted.

MULTIDISCIPLINARY DESIGN OPTIMIZATION OF SOUNDING ROCKETS

The MDO-SONDA was conceived to exploit the synergy between the design disciplines of sounding rockets. Among them, those that use physics-based engineering models are: propulsion, aerodynamics, heating, structures, controls, and trajectory. Its current version interacts in batch mode with

and another of trajectory. Thus, it can exploit the synergy between these two disciplines. Interacting with at least two disciplines makes the MDO-SONDA able to demonstrate the MDO methodology. Besides, it can support multiobjective problems. It can also investigate the trade-offs between the design objectives.

The current version is only prepared for optimization of

to structure proper interfaces for further studies, including the shape optimization of other rocket subsystems, such as

transitions between rocket stages of different diameters. The main aspects of the MDO-SONDA are architecture, inputs, outputs, optimization algorithm, and how to proceed with the optimization.

Architecture

The architecture of the MDO-SONDA is described in two parts: the interaction between the objective function and

another of trajectory (Fig. 2a); and, the interaction between the optimization algorithm and the objective function (Fig. 2b).

rocket simulation (ROSI). The missile datcom is a widely used semi-empirical aerodynamic prediction code, which estimates aerodynamic forces, moments, and stability derivatives for a wide range

descriptors: Mach number, altitude, and angle of attack (Sooy and Schmidt, 2005). Its original version was developed in

the FORTRAN 90 version was documented by the U.S. Air Force (Blake, 1998). The ROSI is also a FORTRAN code. It computes the motion of a rigid body in a three-dimensional space, considering also its rotation in yaw, pitch, and roll axes (Ziegltrum,

successfully used for the trajectory calculation of Brazilian sounding rockets.

Objective function

ROSI software

Trajectory discipline

Missile DATCOM software

Aerodynamic discipline

interacts with interacts with

2nd1st step

Designvariables

Evaluationof each cost

function

Optimizationalgorithm

assigns values to feeds

feed

calls

returnsObjective function

2nd

3rd

4th

5th

1st step

(a) (b)Figure 2. Interactions of the MDO-SONDA.


The MDO-SONDA calls the executable codes in batch mode, which means to run to completion without manual intervention. The missile datcom provides to ROSI the

D), ),

Mq

lp), and center of pressure (Xcp).

lp

rate of the rocket. Unfortunately, the missile datcom does not

l). To use missile datcom calculation indirectly, it is assumed that

l

(Eq. 1):

.C

C C0 01l

l 1

22

c,

d=d (1)

The MDO-SONDA manages the process of each executable code, writes their inputs, and reads their outputs, coordinating their interaction. During the optimization loop, if they freeze for any reason, their processes are, automatically, killed and restarted but with different inputs. First, the MDO-SONDA interacts with missile datcom, obtaining the aerodynamic

of ROSI, which also receives the mass and inertia properties of the rocket, i.e., the changes of mass, center of gravity, moment of inertia and product of inertia, computed by the MDO-SONDA due to spent stage separations, system releases

user-friendly interface to insert input values and to check, graphically, outputs of both missile datcom and ROSI. It also

code, which is automatically generated to make sure that there is not any apparent mistake.

Inputs

The MDO-SONDA inputs can be grouped in three parts.

The second are the elements of the optimization problem:

ones are the optimization algorithm settings. With respect to

spent stage separation, nose fairing ejection, and system release. Such events divide the trajectory calculation into phases, since

of Mach and altitude for each change in rocket geometry, due to the separation of its parts, and jet plume, due to switching a motor on and off. Each phase is characterized by rocket

of the body, propulsion data, and mass and inertia properties of each subsystem that still remains in the rocket during the

computes the total mass and inertia properties of each phase

Outputs

The MDO-SONDA provides an output interface for each executable code and for the optimization results. Using such interfaces, the user can save and analyze later the Pareto-optimal solutions by using the features of the output interface for missile datcom and ROSI in order to verify and validate

Optimization algorithm

Since it is expected that the objective functions have many local minima and maxima and unknown function’s gradient, the appropriate methods are, traditionally, genetic algorithms and simulated annealing, according to the logic decision for choosing MDO, which was proposed by Rowell

MDO-SONDA is based on a multiobjective nongenerational

nongenerational approach is adequate for multiobjective issues, since it preserves individuals that are closer to the Pareto front

this genetic algorithm approach, which was used in this work, is based on the proposal of Borges and Barbosa (2000). The nongenerational algorithm starts generating and assessing the

quantity of iterations is started, which will be satisfactory if all individuals become nondominated at the completion of the optimization. Each iteration consists of selecting two individuals, denoted by parents, generating their offspring that


individual to decide on his/her inclusion into the population. Despite the denomination given to this genetic algorithm, nongenerational, each iteration denotes a generation, since a new individual can be introduced into the population. In the version used in this work, new individuals are accepted only if they are not bad than the worst individual in the population

of the search for optimal solutions instead of the original binary operators, since the optimization problem of current interest is based on continuous objective functions. The proposed real operator works on a normalized search space. Firstly, appropriate values are assigned to its parameters:

c), lower bound of mutation ( inf), and upper bound of mutation ( sup

is a real number and the last two are integers. Secondly, the operator visits each solution that were previously chosen to

chosen solution, a variable ( ) of it that is a design variable is randomly chosen to suffer mutation. Thirdly, an integer ( ) is randomly generated between inf and sup, and a real value (p) is randomly generated between zero and one. Finally, the new value of derives from the old one plus an increment (m), which is given by Eq. 2:

( ),

;m

tv

t v

if p

otherwise1

0=

-

-

=) (2)

where

t c ek2$= -

(3)

c increases,

it is important to establish a compromise between both

How to proceed with the optimization

The optimization is a trial process. It consists of choosing the preliminary intervals for the design variables. The output interface for optimization results uses a method for analyzing multivariate data, which is called parallel coordinates. This

method consists of parallel lines, vertical and equally spaced, where each line corresponds to a design variable and the maximum and minimum values of each variable are usually scaled to the upper and lower boundaries on their respective

graphically, whether the promising region of the search space is reaching the lower and upper bounds or not. Then, if it does, it suggests that the bounds should be extended. Otherwise, it may suggest that the bounds should be more restrictive. Furthermore, the analyses of the optimization results may expose unfeasible conditions that were not considered before in the optimization problem. Thus, the optimization is also a learning process on the self-optimization problem.

CASE STUDY

This section presents the case study of the VS-40 by using the MDO-SONDA. Firstly, the elements of the optimization

nongenerational genetic algorithm will be presented, and

Finally, a mission analysis considering a hypothetical payload mass to microgravity experiment will be presented on the point of view of the trajectory discipline to evaluate the gain

Design problem statement

original design of the VS-40 with a payload of 240 kg, and assuming that this mass is the minimum acceptable for this

To achieve such a goal, two design objectives were pursued:

and maximization of the shortest interval between critical

pressure, minimum static margin, and pitch-roll crossing. The second objective is commonly pursued to avoid subjecting the rocket to severe conditions that could induce an unstable behavior. The transonic speed refers to the range of Mach 0.8 to 1.4, in which severe instability can occur due to oscillating shock waves and large acoustic energy release. The maximum dynamic pressure is often related to the point of maximum

on the instants of both the transonic speed and the maximum


rocket propulsion. The static margin is the position of the center of pressure, where the aerodynamic forces act, minus the position of the center of gravity, both measured with respect to the nose tip as referential and positive in the direction of the rocket tail. If the static margin is negative, that is, the center of pressure is ahead of the center of gravity, the rocket is aerodynamically unstable. If it is positive but too small, it increases the rocket oscillations, which can affect the rocket performance. The pitch-roll crossing, that is, the crossing between the pitch and the roll rates, can lead to a physical phenomenon called roll resonance followed by the roll lock-in, where the roll rate deviates from its

, 1979). These two latter critical

Before proceeding with the comments on the solutions to

second objective suffer as a result? It is also demonstrated that the MDO methodology can be used to investigate whether design objectives are competing or not, leading to a more comprehensive understanding of the system’s trade-offs. Figure 3 describes the design variables. The VS-40 is a

airfoil geometry and two segments. In this case study, only the second segment was subjected to optimization (Fig. 3). Still, the variation of mass and inertia properties related to the shape

for optimization.

The optimization was subjected to the following side

Such constraints are necessary because excessive roll rate affects the structure, and too small static margin increases oscillations. Both situations can affect rocket performance.

Optimization settings and results

Table 2 presents the settings of the multiobjective nongenerational genetic algorithm used in MDO-SONDA. It also shows that the neighborhood radius and the graduation

the distribution of solutions along the Pareto front (Borges and Barbosa, 2000). Despite the small number of design variables, this case study showed that computational cost could become an issue. A single simulation involving interactions between aerodynamics and trajectory calculations takes 12 seconds

objective function were required for seven design variables, the optimization took four hours.

Var-4 1 Var-5 Var-6Var-4 Var-5 Var-6

where0 Var-5 10 Var-6 1

ab

(*)

Var-3

Var-1 (deflection angle)

Var-4

ba

Span station at (*)

Var-2

Var-7 Side (i)

Side (ii)

Note: (ii) is the mirror of (i).

First segment(area = 0.2279 m2)

Second segment

1.2513 m

Fin panel

Table 1. Bounds of the design variables.Design variable Nominal Upper bound

1 (degrees) 0.6 0.42 0.62 (m) 0 0 2.48433 (m) 0.7095 0.7095 0.90954 (m) 1.2513 1 1.25135 0.348038 0.348038 0.4176466 0.799168 0.719 0.9590027 (m) 0.016783 0.011748 0.016783


We have found a Pareto front, demonstrating that the

competing objectives (Fig. 4).

and Barbosa (2000), gave well-distributed points along the Pareto front (Fig. 4). However, despite the fact that population

points. Indeed, in some of these points, there is more than one solution with slight differences between them. Optimization results seem to be coherent. The interval between the transonic speed and the maximum dynamic

The optimization could not lead to solutions that exceed such

reduced up to 29% without increasing the chances of adverse effects that could lead to unstable behaviors (Fig. 4). There are some chances that adverse effects increase when two or more

on -axis the total drag minus its value without computing the

Thus, in terms of the total drag, the reduction was up to 5%. Regarding the parallel coordinates graph, the promising area of the search space has reached the limits of almost the totality of the design variables (Fig. 5). In Fig. 5, regarding the line of Var-7, which is related to the

they also want to avoid unfeasible solutions. Therefore, the lower bound of Var-7 is kept, assuming that its reduction can lead to structural issues. Table 3 presents a Pareto-optimal solution associated with each point in Fig. 4. It is worth noting, based on Var-3 and Var-4 values in Table 3 and the chord at the base of the

Nose tip

Fin profile

-8 -7.5 -7 -6.5 -6 -5.5 -5 -4.50

1

2

3

4

5

6

7

8

9

Drag force caused by fins (kN)

Shor

test

inte

rval

bet

wee

n cr

itica

l flig

ht e

vent

s (s)

123

4

56

78

910

11

Original fins

Pareto-optimalsolutions

Figure 4. Optimization results.

Solution* Variable 1 (m) Variable 2 (m) Variable 3 (m) Variable 4 (m) a (m) b (m) Variable 7 (m)Original 0.600 0.000 0.710 1.251 0.652 0.348 0.0168

1 0.587 2.429 0.793 1.167 0.584 0.382 0.01172 0.589 1.995 0.787 1.084 0.608 0.426 0.01183 0.550 1.995 0.787 1.084 0.597 0.418 0.01184 0.526 1.995 0.787 1.084 0.584 0.409 0.01185 0.427 1.995 0.787 1.084 0.577 0.409 0.01186 0.420 1.766 0.793 1.095 0.610 0.440 0.01177 0.420 1.995 0.812 1.095 0.626 0.423 0.01178 0.420 1.995 0.842 1.066 0.603 0.416 0.01189 0.421 1.995 0.862 1.089 0.622 0.421 0.0118

10 0.420 2.029 0.862 1.212 0.690 0.470 0.011711 0.422 1.855 0.905 1.196 0.684 0.461 0.0118

*solutions are ordered as in Fig. 4.

Table 2. Optimization algorithm settings.

Parameter ValueSize of the population 20Number of generations 600Neighborhood radius 2

0.51.41

Upper bound of mutation 6Float-point precision 0.001


showed in Fig. 3, the Pareto-optimal solutions have from 2.8 to 19.6% more surface than the original panel. Surface area often has more impact than geometry, increasing the drag, despite any attempts to reduce it by choosing an adequate geometry. However, the extended surface area of the Pareto-optimal solutions does not seem to cause any disadvantage in

one that provides more drag in supersonic speed, based on equal surface area and span between the geometries (Fleeman,

causes more drag than the Pareto-optimal solution number 11, which has the largest surface area. Among the Pareto-optimal solutions, the drag increases as

of the surface area (Fig. 6). However, the solution number 1 is an outlier, since it causes less drag than solutions from 2 to 7 but it has an area slightly extended with similar geometry (Fig. 4). Solutions are ordered as in Fig. 4.

Despite the fact that solutions providing the shortest

seconds are those safer than the solution number 1, for

the Pareto-optimal solution that causes the largest reduction of the drag, increasing the rocket’s performance.

Mission analysis

The proposed mission to be analyzed is characterized by a hypothetical payload of 240 kg, which is carried by the VS-40 to be exposed to microgravity environment. If one suppose the mission is scheduled for December, corresponding to the

and Fisch, 2007), when wind surface reduces gradually with

goal is to evaluate what is the gain in the performance of the

ones considering this hypothetical mission. The maximum expected gain can be estimated without performing any optimization. The trajectory simulation

provides an expected gain of 2.9% (Fig. 7). Despite the small

seen that the conditions of a mission analysis can affect the

gain in microgravity of 1.6% (Fig. 7). However, since the VS-40 is an unguided rocket, wind effects and dispersion factors should be considered. The mission analysis consists of taking into account these factors in the evaluation of the

{2,3,4,5}

Original solution

67

9

1110

8

1

1.1 1.15 1.2 1.25 1.3 1.355

5.5

6

6.5

7

7.5

8

Area of the fin panel (m2)

Mag

nitu

de o

f the

dra

g du

e to

fins

(kN

)

Design variablesVar-7Var-6Var-5Var-4Var-3Var-2Var-1

Nor

mal

ized

Inte

rval

1

0

Promisingregionboundary

DominatedsolutionsOriginalfinsPareto-optimalsolutions

Figure 5. Vizualization of the parallel coordinates.


elapsed time in microgravity is estimated by considering the dispersion factors of the rocket. Finally, the gain in the

based on the average value of the elapsed time in microgravity,

Brazilian territory, to compensate for the wind effect, it is necessary to adjust the launch azimuth and elevation based on wind data, which are collected few moments before liftoff. Two types of wind sensing devices are provided, rawinsondes to high altitudes and anemometer measurements

azimuth and elevation adjustments for sounding rockets, still adopted by the Brazilian launch centers, is based on Hennigh (1964). It consists of determining, for a range of launch elevations, the wind weighting as a function of the altitude, and the splashdown displacements caused by a unit range- and cross-wind, respectively. Such displacements are determined by considering the wind up to an upper limit of the effective atmosphere. The range-wind azimuth is given in the direction of the rocket launch tower, while

input, the procedure consists of evaluating the ballistic wind, combining data provided by the wind sensing devices with

the wind weighting function, which had been previously calculated. The ballistic wind is hypothetical and constant in

upper limit of the effective atmosphere. In practice, the upper limit of the effective atmosphere is roughly 25 km (Hennigh, 1964). Finally, considering the ballistic wind, the splashdown displacement caused by a unit wind, and the assumption that the response of the rocket is linear with the wind velocity, the launch azimuth and elevation are adjusted. However, due to stochastic behavior of the wind, dispersion factors of the rocket, structural issues, geographical constraints, and rocket assumption of the linear response to make the adjustments,

ElAA El -ElR A- R ElA

and A are, respectively, the adjusted elevation and azimuth; and, ElR and R are, respectively, the reference elevation and azimuth.

December 2008, obtained with sensors, we have estimated the probability of not violating such constraints for a range of launch azimuth and elevation values, given to one attempt of launch (Fig. 8). Suppose the hypothetical mission cannot exceed two attempts of launch, given that the probability for one attempt (P) can be expressed by Eq. 4:

p P1 1 nn

1

= - -^ h (4)

where, Pn is the probability, between 0 and 1, for n attempts of launch.

Pn at 0.98, for instance, the probability of not violating constraints of launch azimuth and elevation can be at least 0.9 (90%). As the elapsed time in microgravity increases with the launch elevation (Fig. 7), let us select the maximum launch

of nonviolation of the constraints. Based on Fig. 8, the VS-40

The theoretical deviation of the elapsed time in microgravity

of varying the dispersion factors, and computing their results on the trajectory of the rocket, assuming a normal distribution

error. The aerodynamic coefficients are, for instance,

80 80.5 81 81.5 82 82.5 83 83.5 84 84.5 85880

900

920

940

960

980

1000

1020

Launch elevation (degrees)

Elap

sed

fligh

t tim

e in

mic

rogr

avity

(s)

The VS-40 with its original finsThe VS-40 with improved finsThe VS-40 without computing fins drag

Gain of 1.6%

Gain of 2.9%


dispersion factors to be considered. Studies that evaluate the accuracy of the missile datcom compared to experimental wind tunnel data shows that the results for aerodynamic drag are predicted by missile datcom with an error, whose magnitude is less than 20% for a variety of rocket geometries (Sooy and Schmidt, 2005). At transonic speeds, where boundary layer shock interaction takes place, missile datcom does not have the capability to accurately represent such kind of interaction. Table 4 presents the dispersion factors that were assumed to calculate the deviation of the elapsed time in microgravity.

No predominant wind speed and direction have been considered in the calculation of the deviation of the elapsed time in microgravity. Table 5 presents the deviation of the elapsed time in microgravity.

increase from 1.6 to 2.9% (Table 5). As previously discussed, the expected gain does not seem to justify any attempt of

the other hand, it was demonstrated that the factors associated with the mission analysis could affect the gain evaluation. It is expected that, by involving more subsystems and design

Brazilian sounding rockets can be demonstrated regarding different applications, besides their application in microgravity experiments.

FUTURE WORKS

In future works, at least four lines of development should be considered. First, new functionalities may be added to the MDO-SONDA. Interfaces might be created for graphical

plots of the trajectory parameters. The user should be able to customize the optimization problem and to set the interaction

Table 5. Average value and error of the elapsed time in microgravity.

(degrees)Elapsed time in microgravity (s)

81.5 922±138 939±13482 932±138 949±129


Laun

ch a

zim

uth

(deg

rees

)

80 80.5 81 81.5 82 82.5 83 83.5 84 84.5 8530

35

40

45

50

55

60

65

70

75

10

10

10

20

20

20

30

30

30

40

40

40

50

50

50

60

60

60

70

70

70

80

80

80

90

90

1010

10

20

20

20

30

30

30

40

40

40

50

50

50

60

60

60

70

70

70

80

80

80

90

90

90

100

1010

10

20

20

20

30

30

30

40

40

40

50

50

50

60

60

60

70

70

70

80

80

80

90

90

90

100


Laun

ch a

zim

uth

(deg

rees

)

80 80.5 81 81.5 82 82.5 83 83.5 84 84.5 8530

35

40

45

50

55

60

65

70

75

(a) (b)Figure 8. Probability of not violating constraints of launch azimuth and elevation adjustment to compensate for the wind effect (%), given to

Table 4. Dispersion factors error for each rocket stage.

Dispersion factorError

First stage Second stage±0.5 –±3.0 –

Head and cross wind (m/s) ±2.0 –Thrust variation (%) ±3.0 ±3.0Thrust misalignment in pitch and yaw (degrees) ±0.1 ±0.1

Aerodynamic drag (%) ±20.0 ±20.0Weight variation (%) ±1.0 ±1.0Fin misalignment (degrees) ±0.01 ±0.01Ignition time variation (s) – ±2.0


MDO-SONDA. This latter should be able to recalculate the mass and inertia properties of the rocket considering the change of the shape that is being optimized. Data-mining methods might be included in the future to assist the user on searching for trade-offs, when the number of design variables and objectives are such that the traditional methods of data visualization are not enough to make them explicit. Also, the MDO-SONDA should be compared with other codes.

MDO-SONDA, involving more design disciplines. For instance, teamwork involving experts in propulsion and

generate the thrust curve from the propellant variables and to estimate the structural resistance of the rocket against

Third, the optimization mechanisms may be more

combination of two or more metaheuristics, cooperating or competing with each other, and surrogate models might improve the overall performance of the optimization by reducing the number of objective function evaluations. Parallel computing might be used together with such approaches for large-scale optimization problems. The search for appropriate parameter values related to the optimization mechanisms are an issue for future works. Finally, with respect to the last line of development to be seen in future, two or more subsystems may be redesigned, simultaneously, to improve the rocket, for instance, two or more

can be executed to investigate the impact of any variations of the design variables on the its objectives. In addition, two or more missions with respect to the same rocket may be simultaneously considered at the same optimization process.

CONCLUSIONS

In this paper, a MDO application in the context of Brazilian sounding rockets was demonstrated. As case study, the shape

next launches at the Brazilian territory to perform microgravity experiments. This paper began by introducing the concepts of sounding rockets and the microgravity environment, which was followed by presenting facts about the VS-40, and explaining

why it should be revised. Before commenting the results of the optimization, the main aspects of the MDO-SONDA were depicted. It was found that the minimization of the drag due

comprehensive understanding of the VS-40 trade-offs. The drag

in order to avoid adverse effects that could lead to unstable behaviors. However, in terms of the total drag, the reduction was

factors of the rocket. Despite the small gain, it was demonstrated that the factors associated with the mission analysis could affect the gain evaluation. Finally, four lines of development for future works were suggested: the addition of new functionalities to MDO-SONDA; the participation of more design disciplines,

the optimization mechanisms, adding sophisticated methods, such as surrogate models; and the simultaneous optimization of two or more subsystems of the rocket.

REFERENCES

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Engineering Optimization, Rio de Janeiro.


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AIAA, Reston, 468 p.

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Management, Vol. 3, No. 3, pp. 325-330.

www.dlr.de/dlr/en/desktopdefault.aspx/tabid-10081/151_read-4100/.

ASE-RT-010-2004, Institute of Aeronautics and Space, Brazil.

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Oslo, Norway.

Montenbruck, O. l

1st ESA Workshop on Satellite Navigation User Equipment Technologies, Noordwijk, Holland.

Mecânica, Bauru-SP, Brazil, Vol. 1, 9 p.

Brazil. Retrieved in July 6, 2012, from http://pintassilgo2.

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Space Administration (NASA), Hampton, Virginia, NASA/TM-2003-212654, 29 p.

Rockets, Vol. 42, No. 2, pp. 257-265.

Weihs, H. l

of the AIAA 15th Space Planes and Hypersonic Systems and

Weihs, H. l


INTRODUCTION

Composite propellants are being used in several missile applications, which basically contain ammonium perchlorate – AP (from 65 to 70%), a metallic fuel like aluminium powder (15 to 20%), and a liquid binder such as hydroxyl terminated polybutadiene – HTPB (10 to 15%) along with certain process aids and diisocyanate based curatives (Boyars and Klager, 1969). Due to the presence of polymeric binder, propellants are viscoelastic in nature. Vibrational methods are used in order to determine the dynamic mechanical properties of such materials. These vibrational tests measure the deformation of the material to periodic forces. From these dynamic mechanical tests, different variables are obtained, such as: storage modulus

modulus is related with the energy stored during deformation, and loss one is associated with the dissipation of the energy as heat. From the ratio of the loss to storage moduli, loss factor is obtained, and it represents the damping capacity of the material. Thus, dynamic mechanical analysis is a technique that measures the modulus and damping of materials as they are deformed under periodic strain or stress (Ferry, 1980; Groves et al., 1992; Foreman, 1997; Morton et al., 1969). Propellants, which are viscoelastic in nature, are subjected to time, temperature, and frequency effects during the analysis to determine their dynamic mechanical property data (Tod, 1987; Hanus, 2001). The material properties that can be measured by this technique in addition to storage and loss moduli, and

of cure, creep, stress relaxation, and so on. Exhaustive literature survey reveals that a number of studies have been carried out to evaluate the effects of binders,

doi: 10.5028/jatm.2012.04044012

I T DT C RU D

High Energy Materials Research Laboratory – Pune – India

Abstract: Dynamic mechanical analysis is a unique technique that measures the modulus and damping of materials as they are deformed under periodic stress. Propellants, which are viscoelastic in nature, are subjected to time, temperature, and frequency effects during the analysis to determine their dynamic and transient properties. The choice of parameters during the experiments like temperature, frequency, strain (%), and stress level is very crucial to the results obtained since the propellant behaves differently under different conditions. A series of experiments like strain and temperature ramp/frequency sweeps, creep, stress relaxation, etc. have been conducted using high burning rate composite propellant (burn rate ~20 mm/s at 7,000 kPa), in order to determine the precise effects of such parameters on the results obtained. The evaluated data revealed that as the temperature increases the storage modulus, loss modulus, and tan delta curves with respect to the frequency shift towards the lower side. Moreover, there is equivalency between the increase in the tempera-ture and the decrease in the frequency, which can be used for the time-temperature superposition principles. Further, in transient tests, the relaxation modulus has been found to decrease when increasing strain levels in the given time range. Also, relaxation modulus versus time curves were found to shift towards the lower side with increasing temperature while creep compliance decreases with the increase in stress and decrease in temperature. The glass transition value of the composite propellant increases when there is an increase in the heating rate.

Keywords: Glass transition temperature, Storage modulus, Loss modulus, Polybutadiene, Viscoelastic properties.

Received: 06/07/12 Accepted: 08/08/12*author for correspondence: [email protected]


humidity, and composition on the dynamic mechanical properties of propellants, viz. Bhagawan et al. (1995) studied the dynamic mechanical properties of different binders and corresponding propellants in terms of storage modulus and

had the highest one (~ -20 °C) and the propellants had higher moduli than their corresponding binders at any temperature. Cogmez et al. (1999) also attempted to compare the dynamic data of two HTPB-based propellants with different solid compositions, viz., one with 87% solid loading having 16% Al as metallic fuel, and the other with 86% solid loading without metallic fuel; the former propellant was found to be less stiffer and more dissipative than the latter at higher temperatures.

like creep and stress relaxation by Mohandas et al. (2000), who studied the effect of humidity on the transient properties of propellants having the standard composition of HTPB/Al/AP. It was found that when propellant is exposed to high relative humidity (RH) levels, the creep strain increases and equilibrium stress during stress relaxation decreases by a factor of two. Further to this, Musanic (2002) studied double-based propellants and the effect of testing parameters like frequency, heating rate, length to thickness ratio, etc. on their dynamic

However, little work has been carried out to study the effect of various testing parameters on the dynamic as well as transient properties of composite propellants, which are

used during the experiment, i.e., heating rate, frequency, stress/strain level applied, temperature, and so on. Therefore, in the present study, an exhaustive data set was generated to determine the effect of various parameters like frequency, heating rate, strain (%), stress level, temperature on the dynamic and transient properties of composite propellants using different DMA test methods such as:

frequency sweep;

frequency sweep;

In the following section, the effect of the previously mentioned parameters will be reported.

RI NT

The experiments were carried out using high burning rate composite solid propellants having the following composition: HTPB, AP with tetra-modal distribution, Al and other additives with toluene diisocyanate (TDI) as the curative. The testing samples were cut from the propellant block in the form of rectangular bars containing the following dimensions: 60 x 12.5 x 3 mm. All dynamic mechanical measurements were carried out on TA Instruments Dynamic Mechanical Analyser

performed on dual cantilever clamp varying the frequency, temperature, stress, and strain levels.

35 ºC with amplitude increasing linearly from 0.5 to 50

storage modulus was determined by plotting a graph of storage modulus versus strain.

frequency sweep: the sample was given a series of strains at three frequencies viz., 3.5, 11, 35 Hz at 35 ºC and the effect of strain levels was evaluated on modulus by plotting a graph of storage modulus versus frequency.

frequency sweep: the sample was given a constant strain of 0.5% at three frequencies (3.5, 11 and 35 Hz), while varying the temperatures for subsequent tests to determine the effect

sweep: samples were given a constant strain of 0.01%, and temperature increased from 35 to 85 ºC at the heating rates of 1, 2, and 10 ºC/minutes, at the same time the frequencies were varied to determine the frequency effect on storage modulus by plotting a curve of storage modulus versus temperature.

under various strain levels at 35 ºC for 30 minutes, and their relaxation moduli were determined.

Wani,V. et al.


loaded under 0.1% strain level at a series of temperatures ranging from 35 to 85 ºC for 30 minutes, and their relaxation moduli were also determined.

stress levels at 35 ºC for a ten-minute creep time with 20 minutes of recovery, being their creep compliances compared.

ten-minute creep time and 20-minute recovery time: samples were loaded under 1 MPa stress level at a series of temperatures ranging from 35 to 85 ºC for 30 minutes (10-minute creep time and 20-minute recovery time), and their creep compliances were determined.

R U T ND DI CU ION

All the analyses were carried out on high burning rate composite solid propellant (burn rate ~20 mm/s at 7,000 kPa) using different test methods of DMA and varying parameters like temperature, frequency, heating rate, and stress/strain levels. A typical DMA curve of high burning rate composite solid propellant is shown in Fig. 1, wherein the composite propellant was given an oscillation strain of 0.01% with 2 ºC / minutes heating rate at 11 Hz frequency. It is clear from Fig. 1 that the tan delta maximum is at -62.1 ºC, which is taken to be the Tg temperature. The effects of various parameters like temperature, frequency, strain, stress and heating rate on such sample being tested for various dynamic and transient properties like storage modulus, loss

are described in details.

I

The dynamic properties of the high burning rate composite propellant were studied using dual cantilever clamp at an oscillatory strain of 0.5% at a heating rate of 2 ºC/minutes, with temperatures ranging from 35 to 80 ºC at several frequencies. The results for the variation of storage modulus, loss modulus,

Figs. 2 to 4, respectively. It is clear from Figs. 2 to 4 that as the temperature increases the storage modulus versus frequency, loss modulus versus versus frequency curves shift towards the lower side since the temperature decreases the chains become stiffer and less mobile leading to an increase in the modulus. This also supports the fact that an increase in the temperature is equivalent to a decrease in the frequency. The transient tests of creep were also carried out for the high burning rate composite propellants using dual

temperatures, viz., 65 and 75 ºC. The results obtained are presented in Fig. 5, which reveals that the creep compliance/strain increases with the increase in the temperature at a given time range. This might be due to the higher strains induced in

leading to a decrease in the modulus and hence increase in the compliance that is the reciprocal of modulus.

Figure 1. DMA result for a standard sample at 11 Hz with 0.01% oscillatory strain at heating rate of 2 ºC/minutes.


Wani,V. et al.

Figure 4. Tan delta versus frequency for different temperatures at 0.5% strain.

versus frequency for different temperatures at 0.5% strain.

Figure 3. Loss modulus versus frequency for different temperatures at 0.5% strain.


The transient tests of stress relaxation were also carried out

at different temperatures, viz. 35, 40, 45 and 50 ºC. The results obtained are shown in Fig. 6, which shows that relaxation modulus versus time curves shift towards the lower side (the relaxation modulus decreases) as temperature increases. This might be due to the fact that at higher temperatures the

to bear the strain applied resulting in rapid decrease in stress and modulus.

I

rate composite propellant was analysed using dual cantilever clamp at different frequencies, viz, 0.1, 0.2, 1, 2, 3.5, 4.6,

11,35 Hz and so on, at a heating rate of 3 ºC/minutes with an oscillatory strain of 0.01%, and the results obtained are presented in Figs. 7 and 8, respectively. It is clear from Fig. 7 that as the frequency increases, the storage modulus versus temperature curves shifts upwards indicating an increase in the storage modulus with the increase in the frequency. It is also clear from Fig. 8 that as the frequency increases, the storage modulus versus strain curves shift towards the upper side, that is, the storage modulus values increase with the frequency. This may be due to the fact that increase in the frequency (equivalent to decrease times) freezes the chain movements resisting intermolecular slippage, and leading to

et al., 2009; Young and Lovell, 1991). It should be noted that storage modulus versus temperature curves at various frequencies can be shifted using time-

Figure 5. Creep compliance versus time at 1 MPa stress at different temperatures.

Figure 6. Relaxation modulus versus time at different temperatures at 0.1% strain.


master curve at a single reference temperature, thus increasing the frequency range in which the sample properties can be known beyond the frequency range, where the sample was

has the same effect on the measured viscoelastic property as decreases in temperature or in time. The amount of shifting along the horizontal (x-axis) of each curve to align with the reference

by the Williams-Landel-Ferry (WLF) equation (Eq. 1):

/loga C T T C T T1T 0 2 0=- - + -^ ^h h6 @ (1)

whereC1 and C2 are constants,T0 is the reference temperature (K),T is the measurement temperature (K), andaT is the shift factor.

The WLF equation is typically used to describe the time/temperature behaviour of polymers in the Tg region, and it has been reported in the literature to predict the performance of polymers (Foreman, 1997).

I

the high burning rate composite propellant was tested using dual-cantilever clamp at 35 ºC at frequencies from 3.5 to 35 Hz at various strains, ranging from 0.001 to 3% strain, and the results obtained are shown in Fig. 9. It is clear from Fig. 9 that as the strain applied on the sample increases the storage modulus versus frequency curves shift downwards, that is, the storage modulus decreases on increasing the oscillatory strain. This is well-supported by the fact that the

Wani,V. et al.

versus temperature at different frequencies at 0.01% strain.

versus strain in different frequencies at 35 ºC.


elastic component of a material is obtained by the stress ratio from each strain, therefore, if the strain increases, the storage modulus will drop. The effect of strain level was also studied on the stress relaxation behaviour of the high burning rate composite propellant using dual cantilever clamp at 35 ºC with strains varying from 0.01 to 2% for 40 minutes each. The results obtained can be seen in Fig. 10, which shows that the relaxation modulus decreases with increasing strain levels in the given time range, a quite obvious fact since the modulus is obtained by stress ratio from strain, as the strain increases the modulus decreases.

I

The effect of stress applied on the composite propellant when it is subjected to creep was determined by testing the

samples in dual cantilever clamp at 35 ºC for a ten-minute period with the stress applied varying from 0.1 to 3 MPa and measuring the corresponding creep compliances. A plot of creep compliance versus time for high burning rate composite propellant under creep subjected to different stress levels at 35 ºC is shown in Fig. 11, which infers that as the stress level increases the creep compliance versus time curve shifts downwards, that is, the creep compliance decreases

reciprocal of modulus and this is the ratio of stress by strain, therefore, as the stress level increases the modulus increases accordingly, leading to decrease in the compliance. Fig. 11 also reveals that the difference between creep compliance for stress values around 0.1 and 0.5 MPa is more than the difference between the creep compliance for stress values around 2 and 3 MPa. This may be accounted to the fact that at higher stress values the material is strained beyond

versus frequency for different strains at 35 ºC.

Figure 10. Relaxation modulus versus time for different strains at 35 ºC.


its viscoelastic limit, and a permanent set appears in the material so there is only marginal enhancement in modulus for higher stress values.

I

tests were carried out using different heating rates, i.e., 1, 2, 5 and 10 (ºC/minutes) at three frequencies (3.5, 11, 35 Hz) from -80 to 80 ºC at a sinusoidal strain of 0.01%. The results obtained for the heating rates 1, 2 and 10 ºC/minutes are

as the heating rate increases, the curves shift towards the higher temperature side. The values for E’, E”, and tan delta obtained are higher at higher heating rates. Also, the value of

Tg increases as the heating rate increases, as shown in Fig. 14. Tg at 1 ºC/minute heating rate is around -65 ºC, while at

peak (for the value of Tg) starts diminishing at a 5 ºC/minute heating rate. It is clear from Fig. 14 that at a 10 ºC/minute

versus the temperature curve. This may be because the heat transfer from the furnace to the sample is not instantaneous, but depends on the conduction, convection, and radiation that can occur within the DMA instrument. Thus, a thermal lag is present between the sample and the furnace, and as higher the rate of heating, the greater this lag is likely to be present. Therefore, at a 10 ºC/minute heating rate, the sample is not able to acquire the required temperature in such a short term, thereby no peak is observed. Hence, lower heating rates (up to 3 ºC/minutes) are preferred to get accurate results.

Wani,V. et al.

Figure 11. Creep compliance versus time in different stress levels at 35 ºC.


CONC U ION

frequency, strain/stress levels, heating rate on the dynamic, and transient properties of high burning rate composite propellant was studied successfully. The results revealed that

results. Data also showed that increase in the frequency has the same effect on the measured viscoelastic property as decrease in temperature or decrease in the time. An increase in the stress or a decrease in the temperature leads to decrease in the creep compliance, while an increase in the strain or increase in the temperature directs to decrease in the relaxation modulus. Also, an increase in the heating rate or in the frequency shifts

DMA curves to higher temperatures. Very high heating rates (~10 ºC/minutes) get inaccurate results. Therefore, to obtain accurate results, lower heating rates, which cannot be higher than 3 ºC/minutes, are preferred. Moreover, dynamic and transient properties determined at different parameters may be used to: characterize the propellant material, get the shift factors (aT) from multifrequency strain curves at different temperatures using WLF model, develop the master curve for the propellant at the required reference temperature and be used to predict the performance of the propellant over a lifetime of its application.


C NO D NT

parameters.

REFERENCES

et al.

Boyars, C. and Klager, K., 1969, “Propellants, Manufacturing,

Washington D.C, 88p.

Cogmez, A. et al., 1999, “Comparison of two HTPB based composite propellants by dynamic mechanical analysis”, Proceedings of the 30th International Conference of ICT, Karlsruhe, V29 pp. 1-12 or 29-1 to 29-12.

Ferry, J.D., 1980, “Viscoelastic Properties of Polymers”, 3rd

Foreman, J., 1997, “Dynamic mechanical analysis of polymers”, American Laboratory, pp. 198-206.

Groves, I.F. et al., 1992, “Dynamic mechanical analysis – A versatile technique for the viscoelastic characterization of materials”, International Labmate, Vol. 17, Issue 2, TA070.

Hanus, M., 2001, “Dynamic mechanical analysis of composite

trends in research of energetic materials, Czech Republic, pp. 112-121.

Mohandas, C.V. et al.

based propellant”, 3rd

Materials Conference and Exhibit, Thiruvananthapuram, India.

Morton, M. et al., 1969, “Dynamic Response and Damping behaviour of heterogeneous polymers”, Technical report AFML-TR-67-408 Part – II.

results of dynamic mechanical analysis of double base rocket

research of energetic materials”, Czech Republic.

et al., 2009, “Dynamic Mechanical Analysis of

Butadiene Rubber Blends”, Journal of Applied Polymer

Tod, D.A., 1987, “Dynamic mechanical analysis of propellants”, Proceedings of 18th international conference of ICT, Karlsruhe, V 44 pp. 1-14 or 44-1 to 44-14.

Young, R.J. and Lovell, P.A., 1991, Introduction to Polymers, 2nd

Wani,V. et al.


LIST OF SYMBOLS

T Kinetic energy of the system;X Generalized coordinate vector;M Inertia or mass matrix;

First order time derivative of X;

Q Vector that includes external and velocity dependent inertia force;

i Geometrical constrain equations;A Jacobian matrix of constrain equations;

h1, h2, …, hpp independent displacement modes of rigid-body movement;Row vector consists on these combined

Length density of the active cable;T( ) Internal force vector;f( ) Friction force vectors in the contact point;N( ) Pressure force vectors in the contact point;

, Velocity and acceleration of cable length variety;

r Radius of the pulley;

Half angle between the two active cable element nearby the point;Direction cosine of the cable element;

k , k–1 Start and end angle according to contact region;

ij

INTRODUCTION

Deployable truss structures have been applied in many applications, such as solar arrays, masts and antennas (Meguro et al., 2003) that have small-stowed volumes during launch, and are deployed by certain means to assume

a deployable truss structure consists of a large number of struts and kinematic pairs, which are simple, such as revolute

and pantograph struts (Cherniavsky et al.deployable structure has many advantages, including lighter weight, higher precision, smaller launch volume, and higher

on building the dynamic model, solving the differential

Kinematic Analysis of the Deployable Truss Structures for Space ApplicationsXu Yan1*, Guan Fu-ling1, Zheng Yao1, Zhao Mengliang1,2

Zhejiang University, Hangzhou – China2Shanghai JiangNan Architectural Design Institute – China

Abstract: Deployable structure technology has been used in aerospace and civil engineering structures very popularly. This paper reported on a recent development of numerical approaches for the kinematic analysis of the deployable truss structures. The dynamic equations of the constrained system and the computational procedures were summarized. The driving force vectors of the active cables considering the friction force were also formulated. Three types of macroele-ments used in deployable structures were described, including linear scissor-link element, multiangular scissor element, and rigid-plate element. The corresponding constraint equations and the Jacobian matrices of these macroelements were formulated. The accuracy and ef ciency of the proposed approach are illustrated with numerical e amples, including a double-ring deployable truss and a deployable solar array.

Keywords: Kinematic analysis, Deployable truss structures, Macroelements, Deployable solar array.

Zheda Road – Xihu – Hangzhou, Zhejiang – China


only addressed simple beams and rigid bodies, which are

the coordinate partitioning method of constrained Jacobian

tangent space method for constrained multibody systems

basis vector of the tangent space of constrained surface

investigated the mechanism characteristics of deployable truss and tensegrity structures in their literatures (Calladine

et al. (2000)

Newton chord method was employed to solve the equations

approach of deployable structures based on the straight- and

control methods of the hoop truss deployable antenna were

systematically addressed a kinematic analysis method of deployable truss structure based on macroelements, in which

EQUATIONS OF MOTION CONSIDERING DRIVING CONDITIONS

Dynamic equations for the constrained system

For the deployable spatial structure, the dependent Cartesian coordinates are used as generalized ones for the dynamic

two revolute joints at the two ends, and the kinetic energy of

T X MX21 T= o o

whereX: is the generalized coordinate vector, andM

dXdtd

XT

XT

Q 0T

22

22- - =o^ ch m (2)

whereQ: is the vector that includes the external and velocity dependent inertia force,

X

dX MX Q 0T - =p^ h (3)

deployable truss structures, the vector dX, according to the

types of constrains of the entire structures, the geometrical

; , , ,X i s0 1 2i gU = =^ h Since all constrains of the deployable structure are constant with time t during the deployment process, the derivative of

AX 0=o (5)

whereA: is the corresponding Jacobian matrix of constraint

X h h h Hp p1 1 2 2 ga a a a= + + =o o o o o

where

h , h2, … , hp: are p independent displacement modes of the rigid-body movement,

: is a row vector that consists of these combined

et al


X H A AXa= - +p p o o

X H

H MH H MA AH H Q 0T T T

a

a a

=

- - =+

o o

p o o

With the initial condition:

,X X X Xt t0 0 0 0= == =o o

When the initial displacement and velocity vectors are known, the vector t=0

and acceleration of the deployable process in each time step

simple and can be summarized as follows:

simulation, such as the coordinates of the joints, structural topology, constraint conditions, driving mechanisms, boundary conditions and the length of time step, and so on, are provided;

n, the mass matrix and driving force vectors are formed;

Jacobian matrix are formulated;

space of are determined; has full column, the rank will be estimated, if

so go to the ninth step;

displacement, velocity, and acceleration of all joints are obtained;

not, go to the second step; otherwise, go to the ninth step;

step, otherwise, the analysis should be stopped;

Active cable driving and friction

force vectors of the active cables, which forms the term Q in

cables are driven by the motor, the cable length becomes

the active cables will become smaller after it loops over the pulley, so the Coulomb friction law is employed to consider

is assumed as Tactive cable and the pulley in the joints, driving forces of the cables in each deployable element are T2, …, Tk, …, Tn

T > … > Tk > … > Tn.

and a microarc element ds in the contact point between the

length density of the active cable is and the internal force vector is denoted as T(vectors in the contact point are denoted as f ( ) and N ( ),

(a) contact domain

(b) microarc element


*sinT d d N ds

ri i i i+ = +^ ^h h l2o

cosT d d T fi i i i i+ - - =^ ^ ^h h h dslp

where and : are the velocity and acceleration of cable length variety,r: is the radius of the pulley,

: is the half angle between the two active cable elements

and ij in a

X X X X li jT

i j21

- - =^ ^h h6 @

l X XX

Xj i

i

j

m m m= - = -o o oo

o^ h 6 @) 3

where

= l1 (Xj – Xi)

into equations of motion, being the second one employed:

f Ni i=^ ^h hwhere

the result:

d and the limit is gotten by d .

ddT

T prl pl2i

n n= + -p o^ h

provided as k and k

T

T T rl ldT

dk 1

k

k

k2

1

n n

i

ii

+ -=

- -

p o^ h# #

T rl l

T rl le

k

k

2

12

k k1

n n

n n

+ -

+ -= n i i- --

p o

p o

^^ ^h

h h

k – k = – 2 , and the relation of the internal force of cable in two adjacent deployable elements can be

(20)T rl l

T rl le

k

k

2

12

2

n n

n n

+ -

+ -= n r b- -

p o

p o

^^ ^h

h h

cable and the pulley in the joint j

f T Tk k k1 1= -- -

Q Q Qce

ie

je T= " , (22)

Such force of the vector Q of the entire structure is obtained

et al


JACOBIAN MATRICES OF THE MACROELEMENTS

In this section, some deployable macroelements are investigated, and the corresponding Jacobian matrices are

matrices for a constant distance constraint on the members, the position constraint of the sleeve element, the angle constraint of the revolute joints and of synchronize gears have

the formulations are not fully developed here, therefore see Nagaraj et al.

Linear scissor-link element

macroelement in the deployable truss structures, where two pairs

to rotate freely around the axis perpendicular to their common

il and oi is

uniplets ij and lk, and the Jacobian matrix is formulated as

A0 0

0 0e ijT

ijT

klT

klT1

m m

m m=

-

-= G (23)

where the direction cosine of the uniplets ij is

ij = Lij

1 (Xj – Xi), like the kl

It is assumed that the length of oi, ok are a and the length of oj, ol are k*aother at the point orelative position of the connection point o

X Xa k

aX X X

a ka

X1 1

j i k l li -+

+ = -+

+^ ^ ^ ^h h h h

kX X kX X 0i j k l+ - - = (25)

Differentiating it with respect to X, therefore the Jacobian matrix is obtained:A kI I kI Ie

2 3 3 3 3 3 3 3 3= - -# # # #" ,

deployment process, the planar equation is the constraint

r r r 0ij ik il# : =^ h When differences are compared with respect to X, the

A r r r r r r r r r reil jk ij ik ik il ij il ij ik3 1 3 1 3 1 3# # # # #= - # # #^ ^ ^ ^h h h h" ,

Ae = Ae3

Ae3

Ae3

Planar multiangular scissor element

Planar multiangular scissor element, as illustrated in

truss structures, in which the uniplets ij and lk are not aligned at an intermediate point O

constraint equations of the planar multiangular scissor


element include:io, jo, ko and lo of the macroelement are

considered and there are four constant distance constraint equations;

ij and lk are added to the macroelement; thus, there are two constant distance constraint equations;

constraint, which can be formulated by the same method

For the planar multiangular scissor element, the row number of the Jacobian matrix Ae

Rigid-plate element

Planar rigid-plate element is a type of macroelement used ijkl is shown in

degree of the element is analyzed: the degree of freedom of four spatial points i, j, k, lsix struts are appended, the total degree of freedom becomes

Six struts ij, jk, kl, li, ik and jl of the element are considered,

the planar-plate element, the row number of the Jacobian matrix Ae

NUMERICAL EXAMPLES

Double-ring deployable truss

A type of double-ring deployable truss based on quadrilateral elements is investigated for large-size mesh

is deployed by a motor and becomes shorter, the diagonal

Several planar truss elements can make a closed loop by

topology is determined and the major design parameters of the

simulated by the program developed based on numerical

et al


between the active cables and the

When the lengths of the diagonal sleeves are equal to a designed value, the locked constraint conditions of the sleeves work and

(a) x axes coordinate (b) y axes coordinate


which are obtained from the geometric equation of this

Deployable sail arrays

macroelements are dynamic machines during deployment and

each kink) and, after passing over a pulley at joint 2, is connected

= =

dt s is used in

are equal to a designed value, the locked constraint conditions

et al


x-co

ordi

nate

(m)

Time (s)

4.5

4

3.5

3

2.5

2

1.5

1

0.5

00 20 40 60 80 100 120 140

y-co

ordi

nate

(m)

Time (s)

0.6

0.5

0.4

0.3

0.2

0.1

0

-0.10 20 40 60 80 100 120 140

(a) x axes coordinate (b) y axes coordinate

CONCLUSIONS

forces of active cables are combined with the equations of

equations and Jacobian matrices of the macroelements are

deployment process and dynamic parameters at each time step can be simulated for evaluating the deployment behaviors of

the capabilities of this method in the motion analysis are of

error in the time step of the simulation is too large to stop

works are suggested: the reliability analysis of the deployment process can be researched; and deployment control of the

REFERENCES

et al., 2000, “An explicit integration method for real time simulation of multibody vehicle models”,

mass-orthogonal projection method for constrained multibody

for constrained multibody systems”, Computer Methods In

et al., 2005, “Large deployable space antennas based on usage of polygonal pantograph”, Journal of

of multibody system dynamic equations using the coordinate partitioning method in an implict Newmark scheme”,


for state space representation of constrained mechanical

space method for constrained dynamic analysis”, Journal

et al., 2003, “Key technologies for high-accuracy

et al.

et al.approach for static analysis of pantograph masts”, Computers

decomposition for constrained dynamical systems”, Journal

concepts for large antenna structures or solar concentrators”,

method”, Proceedings of the Institution of Mechanical

deployable toroidal spatial truss structures for large mesh antenna”, Journal of the International Association for Shell

et al


LIST OF SYMBOLS AND NOMENCLATURES

CLA Centro de Lançamento de AlcântaraABL Atmospheric Boundary Layer

Boundary layer thicknessH Wind tunnel heightIBL Internal boundary layerIu Turbulence intensityl Coastal cliff heightMIT Mobile integration towerPIV Particle image velocimetry

TA-2 Aeronautic Wind Tunnel of Institute of Aeronau-tics and Space

u(z) LU(z),U(zr ) Mean velocities corresponding to heights z and zr

Uinf Free stream velocityu* Friction velocityzr Reference heightW Wind tunnel width

INTRODUCTION

The majority of the Brazilian rockets are launched from the Centro de Lançamento de Alcântara (CLA), which has a privileged geographical location, 2º 18’S that enables the operation of suborbital vehicles and satellites with safety launchings in several directions over the Atlantic Ocean (Pires et al., 2008; Avelar et al., 2010; Fisch et al., 2010, Pires et al. 2010). An effective use of the launch opportunities at CLA is possible due to the climate conditions with a

demographical density allows the displacement of several sites for launching or logistic support. However, despite the many favorable aspects, mainly because of its proximity to the Equator, the launching center has a peculiar topography due to the existence of a coastal cliff with 40m height (Fig. 1), which can modify the atmospheric boundary layer (ABL) characteristics and consequently affect the safety of rocket launching operations, since the rockets launching pad and the place where the space vehicles are assembled, i.e., mobile integration tower (MIT), are located around 150 to 200m from the border, respectively. Another important physical feature occurrence at the CLA is the formation of an internal boundary layer (IBL) as a consequence of the surface

doi: 10.5028/jatm.2012.04044912

Wind Tunnel Simulation of the Atmospheric Boundary Layer for Studying the Wind Pattern at Centro de Lançamento de AlcântaraAna Cristina Avelar 1*, Fabrício Lamosa Carneiro Brasileiro2, Adolfo Gomes Marto1, Edson R. Marciotto1, Gilberto Fisch1, Amanda Fellipe Faria1

1Instituto de Aeronáutica e Espaço – São José dos Campos/SP – Brazil2Universidade Paulista – São José dos Campos/SP – Brazil

Abstract: Centro de Lançamento de Alcântara is the main Brazilian launching center. In spite of presenting several desirable aspects, due to its proximity to the Equator, it has a peculiar topography because of the existence of a coastal cliff, hich modi es the characteristics of the atmospheric boundary layer. his may affect roc et launching operations, especially hen associated ith safety procedures. his or is a continuation of previous experimental studies about the air o pattern at this launching center. An improved ay of simulating the atmospheric boundary layer in a short test section ind tunnel using passive methods is presented here. It is also presented a preliminary analysis of the air o pattern in Centro de Lançamento de Alcântara, at speci c positions as the edge of cliff and around the mobile integration to er, from ind tunnel measurements using particle image velocimetry. hree values of eynolds number, based on the coastal cliff height, l, ranging from 6.8×105 to 2.0×106, were considered.

Keywords: Atmospheric Flow, Wind unnel, Boundary Layer, Centro de Lançamento de Alcântara.

Received: 04/09/12 Accepted: 10/10/12*author for correspondence: [email protected]ça Marechal Eduardo Gomes, 50 – Vila das AcáciasCEP 12228-904 – São José dos Campos/SP – Brazil


roughness variation, from ocean surface to continental terrain. The wind blowing from the oceanic smooth surface interacts with the low woodland vegetation modifying itself with the formation of an IBL (Pires, 2009), which makes the study of

region even more important.

Figure 1. A general view of Centro de Lançamento de Alcântara.

The simulation of an ABL in a wind tunnel with short-test section is quite complicated and there are several methods for this purpose discussed in the literature (Counihan, 1969). A simple way of generating thick boundary layers is by using passive methods (Barbosa et al., 2002; Loredo-Souza et al.a combination of spires, wedges or grids together with roughness elements distributed on the wall. Ten possible ways of simulating neutral, stable, and unstable atmospheric conditions in different wind tunnel types were described in Hunt and Fernholz (1975). A short review of the techniques used to thicken the boundary layer was presented by Barbosa et al. (2002). Besides, thickening devices with sophisticated geometry were described by Ligrani et al. (1979 and 1983).

designs, which have motivated researchers to choose satisfactory geometries by trial and error. ABL physics is very complex, and the main reason for

the surface, which occurs primarily through mechanical and thermal mechanisms. The mechanical interaction arises from the friction caused by the wind against the ground surface,

and associated turbulence. In the absence of thermal process, the ABL is said to be neutral, and a logarithmic velocity

u(z), characterized by the friction velocity u* and the terrain roughness height zo, is expected to be found (Loredo-Souza et al., 2004). According to Barbosa et al. (2000), for wind speeds higher than 10m/s, the turbulence produced

buoyancy, therefore thermal effects become negligible. This is the case of CLA, where strong winds are observed during the dry season, from July to December. The ABL and atmospheric

from observations, numerical simulations, and wind tunnel measurements (Pires et al., 2008; Avelar et al., 2010; Fisch et al., 2010; Marciotto et al., 2012) The present work is an extended version of a paper recently presented at the fourth AIAA Atmospheric and Space Environments Conference, in New Orleans, from 25 to 28 June 2012, Avelar et al. (2012), and it is also a continuation of a previous study (Avelar et al., 2010), in which the procedures for a boundary layer simulation in a short-test section wind tunnel (TA-2) were described and some preliminary results

CLA region, were presented. Herein, the ABL was simulated using a combination of spires, barrier, and bottom wall surface roughness. The results

wind tunnel, TA-2, of the Instituto de Aeronáutica e Espaço, in Brazil, without using screens downstream of the spires, as in a previous work (Avelar et al., 2010). Three values of Reynolds number ( el) based on the coastal cliff height, l, ranging from 6.8×105 to 2.0×106 were considered. The

sensitive to small Reynolds number variations. In addition, turbulence measurements from hot-wire techniques have been conducted. Some stereo PIV velocity measurements for the values of Reynolds number considered were also conducted, showing strong recirculation regions behind the TMI, and it

small variations of this parameter. METHODOLOGY

inside the ABL, for example, the logarithmic and power law equations (Arya, 2001). According to the logarithmic law,

Avelar, A.C. et al.


the vertical variation of the horizontal wind velocity, U, from the surface up to 100 to 150 m, which corresponds to the

U zku

ln zz r

0

)=^ ` ch j m (1)

where,u*: is the friction velocity,: is the Von Kármán constant,

z0: is the mean terrain roughness, andzr: is assumed to be 10m, which is the height suggested by

the World Meteorology Organization to represent the horizontal wind surface.

The friction velocity, u*, is dependent on the wall shear stress, w, consequently being a measure of the logarithmic declivity close to the wall (Loredo-Souza et al., 2004). Such

to the surface, however it is extensively employed also in the surface layer up to about 100m above sea level (Garratt, 1994).

U zU z

zz

ref ref

ra

=^^ chh m (2)

where,U(zref ): is the mean velocity correspondent to a reference height zref .

The exponent is a characteristic of the type of terrain. It varies from 0.11 for smooth surface as lakes and the ocean to 0.34 for cities with high density of buildings. For the ocean surface, some studies consider between 0.11 (Hsu et al., 1994; Barbosa et al., 2002) and 0.15 (Blessmann, 1973).

Although commonly used, the power law equation has some drawbacks, which were pointed out by Loredo-Souza et al. (2004). Since this equation is valid for any value of zr, the top of the ABL is not recognized in this model. The second issue is that in spite of providing a good representation of the

adjustment in Ekman’s layer, but not into the surface layer. In the present work, the power law equation was used

obtaining z0 and u*. In fact, according to Hsu (1988), in situ measurements of the aerodynamic roughness length are not

always possible since it is related to both the wind speed and the wave characteristics of the ocean. The value of 0.11 for the exponent was assumed in the power law equation.

Wind tunnel atmospheric boundary layer modeling

The experiments were conducted in TA-2, which is a closed-circuit aeronautic subsonic wind tunnel. Its test section has a 2.10m height, H, and 3.00m width, W. A 1,600 HP motor produces a maximum speed of 120m/s through the test section. Spires, roughness elements, and a barrier positioned downstream of spires were used for simulation of a thick boundary layer. The

entrance. The combination of these elements generates the

depend on the desired boundary layer characteristics and on the wind tunnel size, and they were calculated following the methodology proposed by Blessmann (1973). For the boundary layer formation, initially, a set of 180 small blocks with 80×80×20mm was displaced on the wind tunnel bottom wall separated by 150mm. A 200mm high barrier was positioned 350mm downstream of the spires.

the barrier and spires, or changing the density of the roughness elements were tried as well. Two multi-manometers, with Pitot tubes for dynamic pressure measurements installed along its

in the boundary layer. The tallest multi-manometer (rake 1) has 15 Pitot tubes equally distributed along its extension and spaced by 13mm. The smallest one (rake 2) has 16 Pitot tubes. The 11 lowest Pitot tubes are spaced by 5mm and the

Wind Tunnel Simulation of the Atmospheric Boundary Layer for Studying the Wind Pattern at Centro de Lançamento de Alcântara

Figure 2. Multi-manometer with Pitot tubes.

Rake 1

Rake 2


a=0.11, which was assumed to be the closest of what is found over the ocean (Hsu et al., 1994). The positions where dynamic pressure measurements were carried out are represented in Fig. 3.

The circle in Fig. 3 is located in the middle of the test section. The distance between the spires and wind tunnel central line was of 7,860mm.

Turbulence measurements

Turbulence measurements were performed for the free-stream velocity of, approximately, 40m/s. Mean velocity

a constant temperature hot-wire anemometer, from Dantec Dynamics. These measurements were conducted only in the middle of the wind tunnel test section, in the location indicated as P1 in Fig. 3, after the simulation of the atmospheric boundary layer. It was used a straight golden-plated wire probe (55P01). For data collection, a sample rate of 10kHz was used. The measurements were conducted in several vertical positions. A manually controlled device (Fig. 4), which allowed the vertical displacement of the hot-wire probe

during the experiment, was also used. Because of a physical limitation of this device, the highest vertical position where turbulence measurements were conducted was 765mm.

Particle Image Velocimetry measurements

topography was installed in the TA-2 test section, and PIV measurements were conducted at the edge of the coastal cliff and around the MIT. In the present study, the coastal cliff slope angle was assumed as 70º with the horizontal plane, and this value was then reproduced in the model. However, since this inclination angle is not constant along the coastal cliff length, as a continuation of the present analysis, other inclinations will be further considered.

Dynamics two-dimensional PIV system (Fig. 5). The system was a double-cavity pulsed laser, Nd:Yag, 15Hz, with an output power of 200mJ per pulse at the wavelength of 532 nm (New Wave Research, Inc.) and two HiSense 4M CCD camera, built by Hamamatsu Photonics, Inc. with acquisition rate of 11Hz,

A Nikon f# 2.8 lenses with 105mm of focal length was used. The laser sheet was shot from the wind tunnel top wall, which was replaced by a glass window, and such sheet was produced using cylindrical lens placed at the end of an articulated optical arm, which transmits the laser from its source to the region of focus (ROF). This arm was used to allow the laser sheet displacement over the model. The red circles in Fig. 5 indicate locations where PIV measurements were conducted, at the edge of the cliff and around the TMI.

Figure 5. Particle image velocimetry measurements.

glycol water-solution) generated by a Rosco Fog Generator placed inside the wind tunnel diffuser. The digital camera

Avelar, A.C. et al.

Mobile IntegrationTower MIT

Roughnesselements

Spires

Barrier

350 mm

P7

710

P5

500 P4

530 P1 P61030

P2

P3Coastal CliffPosition

Wind

Figure 3. Positions in test section in which dynamic pressures values were measured with the multi-manometer.

Figure 4. Hot-wire probe in the TA-2 test section.


on an aluminum trail supported by a three-axis-positioning device. The number of image pairs captured per second was 5.6, and around 200 image pairs, from each camera, were averaged for one measurement condition. The instantaneous images were processed using the adaptive correlation option of the commercial software Dynamic Studio, developed by Dantec Dynamics. A 32×32- pixel interrogation window with 50% overlap and moving average validation was used.


The configurations tested for the boundary layer

spires, the barrier and the roughness elements were only

tested to illustrate the role of these devices for an appropriated

As can be noticed from Figs. 6 to 10, the spires have a major

the roughness element, the generation of a thick boundary layer is not possible. The barrier has the purpose of generating

180 wood blocks were used. These results seem to indicate

With this purpose, wood strips perpendicularly to the spires were added, as shown in Fig. 14. By adding the three horizontal strips, as observed in

with the exponent 0.11. Comparing Fig. 15 and 17 and observing the correspondent


1.20

1.20

1.00

0.80

0.60

0.40

0.20

0.000.00 0.20 0.40 0.60 0.80 1.00

U = 20 m/s

U = 30 m/s

U = 40 m/s

Power Law = 0.11

y /

U / Uinf

0.000.00 0.20 0.40 0.60 0.80 1.00 1.20

0.20

0.40

0.60

0.80

U = 20 m/s

y /

U / Uinf

U = 30 m/s

U = 40 m/s

Power Law - 0,11


Figures 18 to 21 were included to show some velocity

shows that the wind tunnel lateral walls do not affect

study, Avelar et. al, 2010, for the boundary layer formation in the same wind tunnel, and Fig. 23 presents the velocity

Avelar, A.C. et al.

1.20

1.20

1.00

0.80

0.60

0.40

0.20

0.000.00 0.20 0.40 0.60 0.80 1.00

y /

U / Uinf

U = 20 m/s

U = 30 m/sU = 40 m/sPower Law - 0,11

1.20

1.20

1.00

0.80

0.60

0.40

0.20

0.000.00 0.20 0.40 0.60 0.80 1.00

y /

U / Uinf

U = 20 m/s

U = 30 m/sU = 40 m/sPower Law - 0.11

1.20

1.20

1.00

0.80

0.60

0.40

0.20

0.000.00 0.20 0.40 0.60 0.80 1.00

y /

U / Uinf

U = 10 m/s

U = 20 m/s

U = 30 m/s

U = 40 m/s

Power Law - 0.11


It can be observed that whenever the power-law is well the range of speed studied (from 20 to 40m/s).


1.20

1.20

1.00

0.80

0.60

0.40

0.20

0.000.00 0.20 0.40 0.60 0.80 1.00

y /

U / Uinf

U = 20 m/s

U = 30 m/sU = 40 m/sPower Law - 0,11

Measurement close to the lateral wall of the wind tunnel.

Measurement close to the mobile integration tower site.

1.20

1.20

1.00

0.80

0.60

0.40

0.20

0.000.00 0.20 0.40 0.60 0.80 1.00

y /

U / Uinf

U = 20 m/s

U = 30 m/sU = 40 m/sPower Law - 0.11

1.20

1.20

1.00

0.80

0.60

0.40

0.20

0.000.00 0.20 0.40 0.60 0.80 1.00

y /

U / Uinf

U = 20 m/s

U = 30 m/sU = 40 m/sPower Law - 0.11


Turbulence measurement results

Table 1 shows the intensity turbulence, Iu, measured for various vertical positions and associated h ratio in the central position of the TA-2 test section, where h is the distance from the

the wind tunnel velocity of 40m/s. The turbulence measurements

presented in Table 1 is presented in Fig. 24.

Table 1. Turbulence intensity measurements.Measurement position y (mm) Iu (%) y

P1 765 4.6 0.66P2 665 4.9 0.57P3 565 5.3 0.48P4 465 6.1 0.40P5 365 6.7 0.31P6 315 7.9 0.23P8 215 9.1 0.18P9 165 10.0 0.14

The turbulence intensity values measured in the generated boundary layer, represented in Fig. 24, are in agreement with the values encountered by Wittwer et al. (2012), who experimentally studied CLA small scale models, 1:400 in the wind tunnel “Joaquim Blessmann” of the laboratory LAC / UF , in Porto Alegre, Brazil. In this study, mean

hot-wire anemometer technique. From Table 1 and Fig. 24, it can be observed that the

frequency spectrums, for each vertical position where turbulence measurement were conducted, are shown in Fig. 25. From Fig. 25, it can be observed that in the inertial range the -5/3 Kolmogorov’s law is followed by all curves.

Avelar, A.C. et al.

14

0.7

0.6

0.5

0.4

0.3

0.2

0.12 6 10

y /

Iu (%)

Figure 25. Turbulence spectrum for P1 to P10.

et al. (2010).

1.2

1600

1400

1200

1000

800

600

400

200

00 0.2 0.4 0.6 0.8 1

y /

U / Uinf

TheoreticalU = 27 m/sExpon. (U=27m/s)


Particle Image Velocimetry results

A schematic representation of the CLA wind tunnel model is shown in Fig. 26. The squares numbers 1 and 2 indicate the positions over the model surface, in which the PIV measurements

for the cliff slope of 70º and wind incidence direction of 0º.


1

2

Figure 26. Schematic representation of the Centro de Lançamento de Âlcantara physical model.

(a) el =6.8x105

(b) el =1.4x106

Figure 27. Particle image velocimetry results for the edge of cliff, square number 1, for different el values.

(a) el =6.8×105

(b) el =1.4×106

(c) el =2.0×106

Figure 28. Particle image velocimetry results around the mobile integration tower for different el values.


The PIV measurements were carried out with the

in CLA region. From Figs. 27 and 28, it can be observed an

already separated, the Reynolds number seems not to play an important role. In fact, according to Larose and D’Auteuil (2006), it is expected that bluff bodies with sharp edges, which is the case of MTI, the aerodynamics characteristics are almost insensitive to Reynolds Number as long as this parameter reaches 10,000. It can be pointed out also that the IBL seems to grow asymptotically.

CONCLUSIONS

Following a previous study on the simulation of the ABL in a short-test section wind tunnel, a combination of passive turbulence generators were tested in the present work. Good

horizontal strips were added perpendicularly to the spires in the conventional setup (roughness, barrier, and spires). Whenever the power-law is well-followed, the dimensionless wind speed

regime change for the range of speed studied (from 20 to 40m/s).

around the step corner, representing the coastal cliff, and around the MIT. For the range of Reynolds number tested,

pattern. In both cases, a very turbulent wake was downstream observed. A future analysis of this research will compare wind

ACKNOWLEDGMENTS

The authors would like to thank the technicians José Rogério Banhara and José Ricardo Carvalho de Oliveira, the Engineers Alfredo Canhoto, Wellington dos Santos and Matsuo Chisaki, Ana Clara Dias Barbosa and Tailine Corrêa for their valuable help to this research. Also, to the Agência Espacial Brasileira (AEB), the Conselho acional de esenvolvimento Cient co e ecnol gico (CNPq) under the Grants 559949/2010-3, PQ 303720/2010-7 (Fisch), Universal 471143/2011-1 (Marciotto), and the Fundaç o de Amparo esquisa do Estado de o Paulo

REFERENCES

Arya, S.P., 2001, “Introduction to Micrometeorology”, Academic Press, USA, 2001, 2nd edition.

Avelar, A.C. et al., 2012, “Atmospheric Boundary Layer Simulation in a Wind Tunnel for Analysis of the Wind Flow at the Centro de Lançamentode Alcântara”, 4th AIAA Atmospheric and Space Environments Conference 25-28, New Orleans, Louisiana, AIAA paper 2012-2930.

Avelar, A.C. et al., 2010, “Simulation of the Atmospheric Boundary Layer in a Closed Circuit Wind Tunnel with Short Test Section”, 27th AIAA Aerodynamic Measurement Tecnology and Ground Testing Conference, Chicago, AIAA paper AIAA-2010-4343.

Barbosa, P.H.A. et al., 2000, “Simulation of atmospheric

do XI CBMET 2000, Rio de Janeiro, Brazil.

Barbosa, P.H.A. et al., 2002, “Wind Tunnel Simulation of Atmospheric Boundary Layer Flows”, Journal of the Brazilian Society of Mechanical Sciences, Vol. 24, No. 3, pp. 177-185.

Blessmann, J., 1973, “Simulação da estrutura do vento natural em um túnel de vento aerodinâmico”, Tese (Doutor em Ciências), Instituto Tecnológico da Aeronaútica – ITA, São José dos Campos, Brazil, 169 p.

Counihan, J., 1969, “An improved method of simulating an atmospheric boundary layer in a wind tunnel”, Atmospheric Environment, Vol. 3, pp. 197-214.

Fisch, G. et al., 2010, “The Internal Boundary Layer at the Alcântara Space Center: Winds Measurements, Wind Tunnel Experiments and Numeric Simulations,” Proceedings of the Fifth International Symposium on Computational Wind Engineering (CWE2010) Chapel Hill, North Carolina, USA May 23-27.

Garratt, J.R., 1994, “The Atmospheric Boundary Layer”, Cambridge University Press, Cambridge, USA, 316 p.

H’su, S.A., 1988, “Coastal Meteorology”, Academic Press, San Diego, 260 p.

Avelar, A.C. et al.


Hsu, A.S. et al., 1994, “Determining the Power-Law Wind-

Sea”, Journal of Applied Meteorology, Vol. 33, No. 6, pp. 757-765.

Hunt, J.C.R. and Fernholz, H., 1975, “Wind-tunnel simulation of the atmospheric boundary layer: a report on Euromech 50”, The Journal of Fluid Mechanics, Vol. 70, pp. 543-559. Vol. 7, pp. 361-366.

Larose, G. and D’Auteuil, A. 2006, “On the Reynolds number sensitivity of the aerodynamics of bluff bodies with Sharp edges”, Journal of Wind Engineering and Industrial Aerodynamics, Vol. 94, pp. 365-376.

Ligrani, P.M. et al., 1979, “The Thermal and Hydrodynamic Behavior of Thick Rough-Wall Turbulent Boundary Layers”, Report No HMT-29, Stanford University.

Ligrani, P.M. et al.Boundary Layers for Studying Heat Transfer and Skin-Friction on Rough Surfaces”, Journal of Fluids Engineering, Vol. 105, pp. 146-153.

Loredo-Souza, A.C. et al., 2004, “Simulação da Camada Limite Atmosférica em Túnel de Vento,” Turbulência, Vol. 4, pp. 137-160.

Marciotto, E.R. et al., 2012, “Characterization of Surface Level Wind at the Centro de Lançamento de Alcântara for Use in Rocket Structure Loading and Dispersion Studies”, Journal of Aerospace Technology and Management, Vol. 4, No. 1, pp. 69-79.

Pires, L.M.B. et al., 2008, “Experimentos em Túnel de Vento da Camada Limite Interna no Centro de Lançamento de Alcântara”, Proceedings of Escola de Primavera de Transição e Turbulência, EPTT 2008, São Carlos, São Paulo, Brazil.

Pires, L.B.M., 2009, “Estudo da Camada Limite Interna Desenvolvida em Falésias com Aplicação para o Centro de Lançamento de Alcântara”, Tese (Doutorado em Meteorologia), National Institute for Space Research, São José dos Campos, São Paulo, Brazil, 150 p.

Pires, L.B.M. et al., 2010, “Atmospheric Flow Measurements Using the PIV and HWA Techniques”, Journal of Aerospace Technology and Management, Vol. 2, No. 2, pp. 127-136.

Wittwer, A. R., et al., 2012, “Avaliação Experimental do Escoamento Atmosférico no Centro de Lançamento

Reduzida”, Proceedings of the VIII Escola de Primavera de Transição e Turbulência, 24 a 28 de setembro de 2012, São Paulo – SP, Brazil.



INTRODUCTION

The demand for customized small-scale unmanned air

such platforms in Taylor and Thomas Cord (2003), Green and

span from zero to substantially high advance ratios. In an

et al.

prop-stream effects over the broader range of advance ratios

Stone (2002) comprises numerical modeling and is based on superimposition of prop-stream in numerical panel methods. The prop-stream effects for the small/mini UAV have

d/b) is relatively

that the diameter of the propeller is often comparable to the

Propeller-induced Effects on the Aerodynamics of a Small Unmanned Aerial VehicleAdnan Maqsood*1, Foong Herng Huei2, Tiauw Hiong Go2

1

2Nanyang Technological University – Singapore

Abstract: The present paper has discussed investigations about the propeller slipstream effects on the aerodynamics of a generic unmanned air vehicle platform in the wind tunnel for a broad advance ratio range. The propeller-induced effects

propeller-diameter-to-wing-span-ratio. The stall angle of attack of the small unmanned air vehicle is generally delayed under slipstream effects. The study evaluated the shift in stall angle of attack as a function of propeller-diameter-to-wing-span and advance ratios of the propeller. The aerodynamics of the unmanned air vehicle platform is estimated through wind-tunnel experiments. The study reported in this paper is part of an effort to develop the framework for the analysis

effects on the aerodynamics of a generic small unmanned air vehicle are studied in the wind tunnel for the shift in the aircraft stall angle of attack. The lift-curve slope of the aircraft is independent from the variation of advance ratio. The

using inverse-quadratic relationship. This empirical trend of the stall behavior with advance ratio can be useful in the

Keywords:


d/b

et al.decreases as the advance ratio of the propeller increased. The

d/b in the range of 0.16 to 0.3.

diminished as the advance ratio of the propeller is decreased et al.

as part of the aircraft design and development. The study reported in this paper is part of an effort to

interaction for small/micro UAVs at an early design stage.

tunnel for a broad advance ratio range. The propeller effects

investigations involving various design parameters.

EXPERIMENTAL SETUP AND TESTING

d/b

to maneuver in tight spaces and can hover over the target area

T6061 at Nanyang Technological University (NTU).

et al. (1999). The

electric motor through the Tahmazo®

electronic speed controller (ESC). The control signal from

Pololu®

during testing. The schematic of the propulsion system setup

The relationship of the propeller driven propulsion system is developed from the velocity of the aircraft, the blade-pitch

et al.


pitch propeller, the advance ratio is the primary entity to describe the relationship. The advance ratio (J) of the propeller

JN DV#

= 3 (1)

V : is the true airspeed of the aircraft in m/s,N: is the number of revolutions per second of the propeller, andD: is the diameter of the propeller.

The change in the free-stream velocity causes variation in

throttle setting to delineate the maximum effects of propeller-

behavior (Ralston and Hultberg, 2010), and this is also

observed that the lift-curve slope is practically independent

constant across the complete Reynolds number variation from

Propeller-induced Effects on the Aerodynamics of a Small Unmanned Aerial Vehicle

Coe

ffic

ient

of L

ift

Angle of Attack, degrees

00,10,20,30,40,50,60,70,80,9

1

0 2 4 6 8 10 12

Re = 0.05 MillionRe = 0.08Re = 0.113Re = 0.128Re = 0.147Re =0.156Re = 0.166

Reynolds number range.


by itself is small enough to introduce any effect of Reynolds number on lift curve slope and stall angle of the aircraft.

that the static thrust of the aircraft is approximately 4 N. As the advance ratio is varied, the thrust eventually becomes negative. This means that such high advance ratios can only

et al.

It should be noted that Dprop is negative and represents the propeller thrust.

L L L

D D D

aero total prop

aero total prop

= -

= + (2)


aerodynamic characteristics on both sides of the fuselage

CL

CL varies from

et al.

et al.

in advance ratio. The investigations by Null et al.

slope variation in the current experiment might be caused by

CL

CL behavior can be observed clearly

et al.

Coe

ffic

ient

of L

ift

Angle of Attack, degrees

0

0,2

0,4

0,6

0,8

1

1,2

1,4

0 2 4 6 8 10 12 14

J = 0.39J = 0.56J = 0.63J = 0.73J = 0.78J = 0.82Unpowered


to advance ratio:

,pJa

,stall stall up2a a= + (3)

,pstalla

,upstalla

a ,upstalla = 9.6 and a(.): in degrees.

R2

d/b

this relationship governs the overall shift in the stall behavior

for practical purposes.

CONCLUSIONS

aerodynamic characteristics of small UAV/micro air vehicle

d/b values. The effects are more

against the advance ratio, and such dependence can be

d/bcan be very useful during early aircraft design stage.

REFERENCES

et al.

et al.

AIAA Guidance, Navigation and Control Conference and Exhibit, South Carolina, USA.

California.

et al.

Stone, H. et al.

Propeller-induced Effects on the Aerodynamics of a Small Unmanned Aerial Vehicle


International Air and Space Symposium and Exposition, The

et al.

et al.

et al.


INTRODUCTION

The aerodynamic plasma actuator is a particular

PREVIOUS WORK

et al.

et al.

E D D C

E D D T I

Abstract: Fluids in a dielectric barrier glow discharge at atmospheric pressure have attracted interest from communi-ties of thermo and uid d namics as well as control. his wor investigated the effects of a plasma actuator operating at . and . on the pressure distribution around a pol vin l chloride circular c linder in low-velocit air ow. he e periment was repeated with the actuator at various angles. he results show an acceleration of the ow demonstrated b a lower pressure coef cient at the actuator area. he ow separation was also dela ed b an angle of as much as . hese effects were shown to be greater when the actuator was positioned at an angle closer to the separation area.

Keywords: low discharge ielectric barrier discharge Flow control lasma actuator nthetic plasma et.


et al.et al.

et al.et al.

et al.

et al.et al.

et al.

et al.

PLASMA GENERATION

et al.


thermal plasmas have electrons and heavy particles at the

nonthermal plasmas on the other hand have the ions and

e K e is the electron

e KK and e K


METHODOLOGY

E

CU

P P

P P

P P

21p

T2t=

-=

-

-

3

3

3

3i i^^^h

hh

–

Cvv

1p

2

= -3

` j

V V C1 p= -3

Cp

3.92 1 1.5 6.2 /m sV = + =

Cp

3.92 1 .5 .2 /m sV 2 7= + =

EXPERIMENTAL OBSERVATIONS



Cp curve

Cp

Cp

Cp

Cp

Cp

0 30 60 90 120 150 180 210 240 270 300 330 360

-3.0

-2.5

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5 Actuator at 0°

Cp

Angle

Off On

Cp

0 30 60 90 120 150 180 210 240 270 300 330 360

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0Cp

Angle

Actuator at 60°

Off On

Cp


Cp

Cp

Off On

-30 0 30 60 90 120 150 180 210 240 270 300 330 360

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

Cp

Angle

Actuator at 100°

Cp

Cp

0 30 60 90 120 150 180 210 240 270 300 330 360

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

Cp

Angle

Actuator at 110°

Off On

Cp

Off On

0 30 60 90 120 150 180 210 240 270 300 330 360

-2.5

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

Cp

Angle

Actuator at 80°

Cp

0 30 60 90 120 150 180 210 240 270 300 330 360

-3.0

-2.5

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

Cp

Angle

Actuator at 90°

Off On

Cp

Off On

0 30 60 90 120 150 180 210 240 270 300 330 360

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

Cp

Angle

Actuator at 70°

Cp


0 30 60 90 120 150 180 210 240 270 300 330 360

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

Cp

Angle

Actuator at 120°

Off On

Cp

Cp

Off On

0 30 60 90 120 150 180 210 240 270 300 330 360

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

Cp

Angle

Actuator at 180°

Cp

CONCLUSIONS

Cp

REFERENCES

et al.

et al.

et al.


et al.

et al.

et al.

et al.

et al.

et al.

et al.

et al.

et al.

et al.


INTRODUCTION

PLAnetary Transits and Oscillations of stars (PLATO) is a project from the Cosmic Vision program by the European Space Agency (ESA), which is currently in phase M3 (launch scheduled for 2022). A large-scale photometer composed of 34 cameras is used, whose objective is to detect and characterize extrasolar planets and their host stars by using the transit method (Catala and PLATO consortium, 2008). This multi-camera approach will jointly observe the same region of the sky. There is a front-end electronic system (FEE) associated with each camera, which is responsible for image digitization and for their sending to the digital processing unit (DPU) of the satellite. The satellite has 16 normal DPU (N-DPUs) for mining scienti c data from the images and two fast ones (F-DPUs) for interfacing with fast cameras. For the FEE and DPU communication, the SpaceWire and remote memory access protocols (RMAP) were selected (ESA, 2008), both of which were widely used in several space missions (Parkes and Ferrer, 2010). The project involves the creation of a dedicated device, which is capable of simulating the behavior of a normal FEE (NFEE) with the main objective of testing the proposed architecture for the PLATO satellite and of validating, in a near future, the DPU s ight software. The system was implemented in VHSIC hardware description language (VHDL) and embedded into a ilinx C VL 0 eld-programmable gate array (FPGA) (Xilinx, 2011).

PLATO ARCHITECTURE

The architecture proposed for the PLATO satellite (Larque and Plasson, 2011) has 34 independent telescopes, each made up of an optical unit and a camera. From the 34 cameras, 32 are normal ones (N-Cameras), meant for scienti c use and delivery of high-resolution photometry (81 Mpx/Camera, 1 px = 16 bits) at 25-second intervals. The two remaining cameras (F-Cameras) are used in the satellite’s control loop at 2.25-second intervals (41 Mpx/camera). The 32 scienti c-purpose N-Cameras were arranged in the satellite in four subgroups of eight cameras (Fig. 1). The two fast cameras work independently and have each a dedicated processing unit. The charge-coupled device (CCDs) used in the mission have two zones: the image zone, which is made up of detec-tors, and the memory zone, used for the storage of temporary images. The memory zone has two access points, thus enabling simultaneous readings in different areas. Each camera is equipped with its own four-CCD-wide matrix, each containing 4,520 lines by 4,535 pixels. A FEE is coupled to each camera and is responsible for controlling CCD operation and communicating with the DPU, sending images and status information (housekeeping). The FEE also propagates time information to the DPU, which uses these data for image processing. Image delivery occurs every 6.25 seconds, alternating the CCDs at every cycle. The process is handled using two SpaceWire links that operate each at 100 Mbps. Each link transfers half of a CCD’s image (right and left sides) simultaneously, increasing, thus, the system’s general data signaling rate to 200 Mbps. The RMAP was chosen for image transmission. This protocol allows writing and reading from a memory unit

doi: 10.5028/jatm.2012. 04042712

Electronic Simulator of the PLATO Satellite Imaging SystemRafael Corsi Ferrão*, Sergio Ribeiro Augusto, Tiago Sanches da Silva, Vanderlei Cunha ParroInstituto Mauá de Tecnologia – São Caetano do Sul/SP – Brazil

Abstract: This paper described an architecture that is able to emulate the behavior of the imaging transfer system proposed for the PLATO satellite – PLAnetary transits and oscillations of stars. It was conceived to accurately repre-sent ight operation as to validate the satellite digital processing unit on its development phase. etails related to the PLATO mission, its architecture, and the implementation technical details are presented in this article.

Keywords: SpaceWire, Remote Memory Access Protocol, PLAnetary Transits and Oscillations of stars, Ground support equipment, Hardware description languages.

Received: 08/05/12 Accepted: 06/09/12*author for correspondence: [email protected]ça Mauá 1 – CEP: 09580 900 – São Caetano do Sul/SP – Brazil


through a SpaceWire node. Image transmission takes place in a manner which is transparent to both nodes, that is, without demanding processor intervention. Delivery is carried out by a RMAP command that encap-sulates half of a CCD’s line. This command is sent without data veri cation or acknowledgement. Control commands sent from the DPU to the FEE, on the other hand, are provided with both of these capabilities. A total of 16 Normal-DPUs execute onboard of the satel-lite routines for mining scienti c data from the images stored in the memories. These routines run in parallel with image reception. Each N-DPU is responsible for data processing sent by two N-FEEs. The complete transfer of a quarter of an image (one CCD) must last at most 3.3 seconds, leaving enough spare time for the complete execution of the image data mining. Image transfer happens as depicted in Fig. 2.

ARCHITECTURE PROPOSED FOR THE SIMULATOR

The proposed platform (SimuCam) emulates a NFEE capable of testing half of a Normal-DPU (each NDPU takes

care of two NFEE), sending new images at each sync signal. The objectives are to have a system capable of dynamically test the DPU’s embedded software and to validate the archi-tecture proposed for the mission. The architecture shown in Fig. 3 has as its cornerstone two volatile memories used for temporary storage of the image to be transmitted. These memories work complementarily and are used as data buffers. Their operation may be described as follows: while a memory is being loaded with a new image that will be transferred, the other one is being read and its data sent by the simulator to the DPU. At the next sync signal, the

Ferrão, R.C. et al.

Figure 1. Optic bank.

Figure 2. One cycle of the Plato charge-coupled device transfer timing.


memories are switched, and a new picture is loaded into the memory that was recently read. Figure 4 illustrates the process. Image loading into memory is performed through a high-speed communication path (USB) between the workstation and the memory controller (Fig. 5), being the latter responsible for managing memory access and access to the simulator’s internal register. The SpW/RMAP block is responsible for the interface with the SpaceWire codec and for the creation and interpretation

of RMAP commands. Implemented according to the ECSS-E-ST-50 standard (ESA, 2008), the block allows the DPU to read and modify the simulator’s internal registers, changing in this way its behavior during operation. It also creates and automatically sends RMAP writing commands containing the image that will be transferred, in addition to extra information in each command (information about the line and the CCD). Housekeeping data are sent separately, after the end of trans-mission of each CCD. The RMAP handling has two internal interfaces, read and write, which take care of generating and propagating RMAP packets, as well as of interpreting RMAP commands sent by the DPU. These interfaces run in parallel, allowing a writing command to be executed in local memory, while images are being transferred by the write block. If a write block interprets a command with acknowledgemen t or a read command, the answer is placed on a response queue (with a priority level inferior to that of the images). Only after the end of the image data transmission, the answers are dispatched. The read block is a Mealy state machine that interprets, reads or writes commands (with or without veri cation and/or answer). Communication to memory and to the codec is direct, making it entirely transparent to other blocks. The read block restrains the access to some data slots in memory, which can only be accessed by the DPU in some operation modes (Housekeeping, for example). Data concerned with the interpreted packets are recorded in the simulator’s register. The writing block is responsible for creating headers and for sending data to the DPU. At each new sync signal, the writing block reads from the simulator’s general register the initial con guration information of the SpaceWire link, DPU, and simulator. Using this information for image transmission,

Electronic Simulator of the PLATO Satellite Imaging System

Figure 4. Data ow in memory Figure 5. Memory allocation data.

Figure 3. Simulator architecture (SimuCam).


the RMAP header con gurations are automatically changed at each new command, without the need for external control. Data sent to the DPU are previously loaded into an autono-mous rst in rst out (FIFO) device (Haywood, 2004) by the data charge block, which addresses the memory controller by taking into account memory-image allocation. This charging procedure is necessary so that the memory controller can be shared with more than one RMAP writing block. Figure 6 illustrates the blocks interfaces of a single RMAP handling. The pattern generation block interprets addresses sent by data charge and creates patterns that are used for tests and validation. The simulator has an internal register bank that contains all of its con gurations. These registers are shared by all blocks, and the con gurations can be changed both by the USB interface and by the RMAP writing command sent by the DPU to the simulator. A TimeCode (ESA, 2008) is used for the propagation of time and is sent automatically whenever a sync signal is received, ensuring time propagation required by the applica-tion. The TimeCode points out to the DPU the beginning of a new transmission. Besides, it is also used to indicate the CCD (zero, one, two, three) being transferred. The sync block propagates or creates reference signals (6.25 and 25 seconds) used by all other blocks for synchronism and internal update.

The FEE control is a state machine that controls memory access to both the RMAP reading and loading blocks. It also oversees the memory exchange operation and takes care of arbitrating reading accesses, ensuring that both RMAP memo-ries always have their internal buffers suf ciently full so that the image transmission does not remain idle, hence raising the system’s ef ciency. The control logic takes into consideration that the reading operation executed by the RMAP block is performed in bursts (of con gurable size), raising the ef -ciency of the access to the system’s SDRAM memory. A codec provided by Commissariat à l’Énergie Atom-ique (CEA) was used for applications in Xilinx Virtex5 chips (Frédéric and CEA, 2008). It implements all protocol speci cations with low-chip resource usage. Furthermore, it supports a maximum transmission frequency of 320 Mbps and has a 32-byte reception FIFO. Several system and simulator characteristics can be modi ed, enabling the simulator to be as generic as possible. The following con gurations can be made:

simulator parameters: they con gure the simulator’s operation and its operation mode. Among others, CCD reading direction (clockwise or counter-clockwise) can be selected. Likewise, error insertions into RMAP packets and internal and external synchronisms (and synchronism period) can be con gured

housekeeping: related to SDRAM memory space allocation, data quantity, and RMAP protocol con gurations

CCD description: it occupies itself with CCD physical information, such as size, pre-scan columns, semi-ring rows, analog to digital converter (ADC) speed and resolution, delays between lines, and RMAP protocol con gurations used for sending the image to the DPU.

IMPLEMENTATION AND TESTS

The GR-PCI-XC5V board from Aero ex Gaisler (Gaisler, 2011) was used for the implementation of the simulator SimuCam. This board has a FPGA Virtex5 XC5VLX50 core from Xilinx, as well as FLASH, SRAM, and SDRAM memories. The hardware proposed for the application was described in VHDL in order to run operations in parallel, aiming at the global increase in performance of the simulator. All blocks having an interface between FIFOs were implemented using the concept of autonomous FIFO, dispensing data ow control between blocks. The chip’s internal RAM was used as memory for the FIFOs.


Figure 6. RMAP handling interfaces.


All blocks within this project had their logic and functionality tested with test benches exclusively created for each block. A global system simulation was performed verifying system integrity and timing. Tests were carried out in the Paris Observatory, in Meudon jointly with LESIA (Laboratoire d’Études Spatiales et d’Instrumentation en Astrophysique), which is responsible for the speci cation of the satellite’s architecture. The test scenario can be seen in Fig. 7. The simulator was connected via two SpaceWire links to a GR-RASTA system (Gaisler, 2011) enclosing a LEON2 processor, left in charge of simulating a half normal DPU, and a Debug System Unit (DSU) to perform rmware debugging. The tests involved sending data (patterns) from the simula-tor to LEON2’s memory zone, which was responsible for verifying data integrity. Timing imposed by the speci cation was tested with a SpaceWire link analyzer (STAR-Dundee, 2012). A Software Ground Support Equipment (SGSE) was developed to control the SimuCam. The sent patterns were prede ned and were generated inside the simulator with the pattern generation block (Fig. 6).

Figure 7. Test bench.

CONCLUSIONS

With the aid of the proposed system, it was possible to send an entire image from the SimuCam in only 2.93 seconds, better than the originally time proposed, i.e., 3.3 seconds (Table 1). The SpaceWire links attained a 99.93% utilization rate, being held idle for 240 ns for every half line sent (between RMAP commands) as shown in Fig. 8. The use of the platform at a higher speed of the SpaceWire link (200 Mbps) was also possible without incurring ef ciency loss in the link.

The entire implementation of the architecture (SpaceWire codec RMAP handling memory controller USB handling FEE control, and internal register) used 45% (13314) of the

Electronic Simulator of the PLATO Satellite Imaging System

40 ns 20 ns NCHAR0 ns 40 ns EOP20 ns 20 ns 80 ns80 ns 60 ns NULL100 ns 20 ns 80 ns140 ns 40 ns 40 ns160 ns 20 ns NULL220 ns 60 ns 80 ns240 ns 20 ns NULL300 ns 60 ns 80 ns340 ns 40 ns NCHAR [FE]380 ns 40 ns 80 ns440 ns 60 ns NCHAR [01]

LinkIdle

EOP – End of PackageNCHAR – Normal character

Figure 8. Transfer analysis of the RMAP protocol with the least time between two lines using Stardundee’s SpaceWire analyzer.

Table 1. Comparative table between the real NFEE and the obtained timing from implemented SimuCam.

Con guration NFEE SimuCam Unit

CCD width 4,510 4,510 Pixels

CCD height 4,510 4,510 Pixels

Smearing rows 10 10 Rows

Prescan per CCD output 25 25 Pixels

Pixel coding 16 16 Bits

ADC Speed 4 6.97 MHz

SpaceWire links 2 2 Number

SpaceWire bit rates 100 100 Mbps

Full CCD total transfer time 3.3 2.93 seconds

Package size + 16 bytes RMAP header 4,576 4,576 Bytes

Instantaneous data rate for 1 link 80 100 Mbps

Averaged data rate over full CCD transfer

70.5 99.93 Mbps

NFEE: normal front end electronicADC: analog to digital converter.


FPGA’s lookup table (LUTs) and 22% (6,568) of its slice registers, enabling the implementation of two simulators in the same chip. The proposed architecture from the system is exible and suitable to be used again in other projects. Since it was devel-oped with emphasis on recon gurability, this system allows the dynamic modi cation of its behavior through changes in several simulator values, such as CCD size, ADC rate, RMAP protocol con gurations, reading direction, among others.

ACKNOWLEDGEMENTS

The authors thank the National Council for Scienti c and Technological Development – Brazil (CNPq) for the nancial support, Mauá Institute of Technology (IMT) and LESIA for the collaboration opportunity.

REFERENCES

Catala, C. and PLATO consortium, 2008, “PLATO PLAnetary Transits and Oscillations of stars – A study of exoplanetary systems, proposal”, Vol. 1, Retrieved in 20 Feb. 2011, from http://www.lesia.obspm.fr/perso/claude-catala/plato_web.html.

ESA Requirements and Standards Division, 2008, “SpaceWire links, nodes, routers and networks ECSS-E-ST-50-11C”.

Frédéric, P. and CEA, 2008, “Commissariat à l’Énergie Atomique et aux énergies alternatives”, IP SpaceWire Data Sheet.

Gaisler, A., 2011, “Aero ex Gaisler”, Retrieved in 21 Feb. 2011, from http://www.gaisler.com/cms/.

Haywood, S., “Autonomous Cascadable Dual Port FIFO”, Retrieved in 2004, from http://www.spacewire.co.uk/auto_ fo.html.

Larque, T. and Plasson, P., 2011, “DPU FEE interface requirement document”, PLATO DPS TS 138 THALES, Vol. 11.

Parkes, S. and Ferrer, A., 2010, “SpaceWire-D: Deterministic Data Delivery with SpaceWire”, In: Proceedings of the 3rd International SpaceWire Conference, St. Petersburg.

STAR-Dundee, 2011, “SpaceWire Link Analyser Mk2”, Retrieved in 20 Mar. 2012, from http://www.star-dundee.com/products.

Xilinx, 2011, “FPGA, CPLD, and EPP Solutions”, Retrieved in 1 Set. 2011, from www.xilinx.com/.



A Comprehensive Investigation of Retrodirective Cross-Eye Jamming

Warren Paul du PlessisUniversity of PretoriaPretoria – South [email protected]

Thesis submitted for PhD degree in Electronic Engineering at the University of Pretoria, Pretoria, South Africa, in 2010.

Advisors: Professors J. Wimpie Odendaal and Johan Joubert

Keywords: Electronic countermeasures, Jamming, Monopulse radar, Tracking radar, Electronic warfare.

Abstract: Cross-eye jamming is an electronic attack technique that induces an angular error in the radar being jammed. The

monopulse tracking radars, which are largely immune to other forms of jamming. The objective of this research was to gain a complete understanding of cross-eye jamming so that systems

The main contribution of this work is a comprehensive mathematical and experimental study of retrodirective cross-eye jamming. The mathematical analysis considers all aspects of an isolated, single-loop, retrodirective cross-eye jamming engagement, thereby avoiding the approximations inherent in other cross-eye jamming analyses. Laboratory experiments that accurately represent reality, by using the radar for both transmission and reception, and simulating a true retrodirective cross-eye jammer, were performed to validate the theoretical analysis. Lastly, the relationship between the angular error induced in the radar being jammed and the matching required from a cross-eye jammer system was explored. The most important conclusion of this work is that the traditional analyses of cross-eye jamming are inaccurate for the conditions under which cross-eye jammers operate. These inaccuracies mean that the traditional analyses are overly conservative, particularly at short ranges and for high cross-eye gains, suggesting that practical cross-eye jammers can be realized more easily than is generally believed.

Ant Colony Optimization Applied to Laminated Composite Materials

Rubem Matimoto KoideUniversidade Tecnológica Federal do ParanáCuritiba/PR – [email protected]

Thesis submitted for Masters in Mechanical and Materials Engineering at Universidade Tecnológica Federal do Paraná (UTFPR), Curitiba, Paraná, Brazil, in 2010.

Advisor: PhD Marco Antonio Luersen

Keywords: Ant colony optimization, Meta-heuristic, Laminated composite materials.

Abstract: The ant colony algorithm is a heuristic that was formulated in the 1990s by Marco Dorigo. The idea was inspired by the behavior of real ants, related to their ability to

was running by exploiting the pheromone trails, a chemical substance deposited by the ants during their journeys. Due to this cooperative behavior and effective search, the ants build

then simulated in optimization algorithms, called ant colony optimization. Thus, this dissertation aimed at studying and applying the ant colony method to the optimization of laminated composite materials. This kind of material is made by stacking plies, in which each ply is composed by a matrix, usually

related to the best settings of the orientation angles of the plies,

implemented and applied to laminated composite plate issues, such as the maximization of the strength, the minimization of the cost, and the maximization of the fundamental frequency. This last issue was also solved using an interface with the

of problems without an analytical solution for the structural response. The numerical tests carried out indicate that the method is competitive compared to other techniques found in the literature for the optimization of composite laminates materials.

Thesis abstractsThis section presents the abstract of most recent Master or PhD thesis related to aerospace technology and management


Classifying Low Probability of Intercept Radar Using Fuzzy ARTMAP

Pieter Frederick PotgieterUniversity of PretoriaPretoria – South [email protected]

Dissertation submitted for Masters in Electronic Engineering at the University of Pretoria, Pretoria, South Africa, in 2011.

Advisor: Professor Jan Corné Olivier

Keywords:Electronic support, Estimation, Fuzzy ARTMAP, Intercept receiver, Low probability of intercept, Parameters, Performance, Radar.

Abstract: Electronic support operations concern themselves with the ability to search for, intercept, track, and classify threat emitters. Modern radar systems in turn aim at operating undetected by intercept receivers. These radar systems maintain low probability of intercept by using low power emissions, coded waveforms, wideband operation, narrow beam widths, and evasive scan patterns without compromising accuracy and resolution. The term low probability of intercept refers to the small chance or likelihood of intercept actually occurring. The complexity and degrees of freedom available to modern radar place a high demand on electronic support systems to provide detailed and accurate real-time information. Intercept alone

(using Fuzzy ARTMAP) of the Pilot Mk3 low probability of intercept radar. Fuzzy ARTMAP is a cognitive neural method combining fuzzy logic and adaptive resonance theory to

ARTMAP systems are formed by self-organizing neural architectures that are able to rapidly learn and classify both

discreet and continuous input patterns. To evaluate the suitability of a given electronic support intercept receiver against a particular low probability of intercept radar, the

by combining the radar range, intercept receiver range, and sensitivity equations. The radar wants to force an opposing intercept receiver into its range envelope. On the contrary, the intercept receiver would ideally want to operate outside

the radar. The maximum likelihood detector developed for this study was capable of detecting the Pilot Mk3 radar, as it allowed enough integration gain for detection beyond the radar maximum range. The accuracy of parameter estimation in an intercept receiver is of great importance, as it has a

Among the various potentially useful radar parameters, antenna rotation rate, transmit frequency, frequency sweep, and sweep repetition frequency were used to classify the Pilot Mk3 radar. Estimation of these parameters resulted in very clear clustering of parameter data that distinguish the Pilot Mk3 radar. The estimated radar signal parameters are well separated to the point that there is no overlap of features. If the detector is able to detect an intercepted signal, it will be able to make accurate estimates of these parameters. The

modes of the Pilot Mk3 low probability of intercept radar.

output spaces is achieved from a training dataset comprising only 5% of the total dataset. If any radar has a low probability of intercept, there must be a consideration for the radar as well as the opposing intercept receiver. Calculating the low probability of intercept performance factor is a useful tool for such an evaluation. The claim that a particular radar has low probability of intercept against any intercept receiver is too broad to be insightful. This also holds for an intercept receiver claiming to have 100% probability of intercept against any radar.


Besides the participation of Editorial Board, the Journal of Aerospace Technology and Management had a collaboration of specialists as reviewers to evaluate the manuscripts. To them, the JATM thanks for the contribution in Vol. 4 (2012)

Abel de Lima Nepomuceno - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Adilson Walter Chinatto Jr. - Universidade de Campinas - Campinas/SP - Brazil

Adolfo Gomes Marto - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Adrián Roberto Wittwer - Universidad Nacional del Nordeste – Resistencia/ Argentina

Alberto W. S. Mello Junior - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Alessandro Teixeira Neto - Escola de Engenharia de São Carlos - São Carlos/SP - Brazil

Alex da Silva Sirqueira - Fundação Centro Universitário Est. Zona Oeste - Rio de Janeiro/RJ - Brazil

Alexandre Nogueira Barbosa - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Amaury Caruzzo - Instituto Tecnológico de Aeronáutica - São José dos Campos/SP - Brazil

Ana Cristina Figueiredo M. Costa - Universidade Fed. de Campina Grande - Campina Grande/PB - Brazil

Anderson Zigiotto - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Andrea Fukue - Embraer - São José dos Campos/SP - Brazil

Arcanjo Lenzi - Universidade Federal de Santa Catarina - Florianópolis - Brazil

Argemiro Soares S. Sobrinho - Instituto Tecnológico de Aeronáutica - São José dos Campos/SP - Brazil

Ariovaldo Felix Palmério - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Bernardo Barbosa da Silva - Universidade Federal de Pernambuco - Recife/PE - Brasil

Carlos Alberto Gurgel Veras - Universidade de Brasília- Brasília/DF - Brazil

Carlos Alberto Rocha Pimentel - Universidade Federal do ABC - Santo André/SP - Brazil

Carlos R. Ilário da Silva - Embraer - São José dos Campos/SP - Brazil

Cayo Prado Fernandes Francisco - Universidade Federal do ABC - Santo André/SP - Brazil

Célio Costa Vaz - Orbital Engenharia - São José dos Campos/SP - Brazil

Cesar J. Deschamps - Universidade Federal de Santa Catarina - Florianópolis - Brazil

Cícero R. de Lima - Universidade Federal do ABC - Santo André/SP - Brazil

Clovis Sansigolo - Instituto Nacional de Pesquisa Espaciais - São José dos Campos/SP - Brazil

Daniel Soares de Almeida - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Daniela Buske - Universidade Federal de Pelotas- Pelotas/RS - Brazil

Dimitri Mavris - Georgia Institute of Technology - Atlanta/GE - USA

Elcio Jeronimo de Oliveira - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

AD HOC REFEREES


Elizangela Camilo - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Evaldo Simões da Fonseca - Instituto Nacional de Pesquisa Espaciais - São José dos Campos/SP - Brazil

Fernando de Souza Costa - Instituto Nacional de Pesquisas Espaciais - Cachoeira Paulista/SP- Brasil

Francisco K. Arakaki – Embraer - São José dos Campos/SP - Brazil

Furio Damiani - Universidade Estadual de Campinas - Campinas/SP - Brazil

Gustavo Bono - Universidade Federal de Pernambuco - Recife/PE - Brazil

Gustavo Ripper - Inst. Nac. Metrologia, Normalização e Qualidade Ind.- Rio de Janeiro/RJ - Brazil

Gustavo Trapp - Embraer - São José dos Campos/SP - Brazil

Heidi Korzenowski - VALE Soluções em Energia - São José dos Campos/SP - Brazil

Helcio Francisco Villa Nova -Universidade Federal do ABC - Santo André/SP - Brazil

Heraldo Silva da Costa Mattos - Universidade Federal Fluminense - Niterói/RJ - Brazil

Hilton Cleber Pietrobom - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Hossein Bonyan Khamseh - Shahid Beheshti University - Irã

Humberto Araújo Machado - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Ijar M. Fonseca - Instituto Nacional de Pesquisa Espaciais - São José dos Campos/SP - Brazil

Jesuíno Takachi Tomita - Instituto Tecnológico de Aeronáutica - São José dos Campos/SP - Brazil

Joana Ribeiro - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Joao Batista de Aguiar - Universidade Federal do ABC - Santo André/SP - Brazil

João Batista P. Falcão Filho - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

João Marcelo Vedovoto - Universidade Federal de Uberlândia - Uberlândia/MG - Brazil

Jonas Gentina - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Jorge Rady de Almeida Junior - Escola Politécnica da USP -São Paulo/SP - Brazil

Jose Carlos Gois - Universidade de Coimbra/Coimbra - Portugal

Jose Maria Fernandes Marlet - Embraer - São José dos Campos/SP - Brazil

Katia Lucchesi - Universidade de Campinas - Campinas/SP - Brazil

Leonel Fernando Perondi - Instituto Nacional de Pesquisa Espaciais - São José dos Campos/SP - Brazil

Leopoldo P.R. de Oliveira - Escola de Engenharia de São Carlos - São Carlos/SP - Brazil

Ligia Maria Soto Urbina - Instituto Tecnológico de Aeronáutica - São José dos Campos/SP - Brazil

Luciano Kiyoshi Araki - Universidade Federal do Paraná - Curitiba/PR - Brazil

Luciene Dias Villar - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil


Luis Cláudio Rezende - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Luiz Carlos Gadelha - Instituto Nacional de Pesquisa Espaciais - São José dos Campos/SP - Brazil

Luiz Carlos S. Goes - Instituto Tecnológico de Aeronáutica - São José dos Campos/SP - Brazil

Marat Rafkov - Universidade Federal do ABC - Santo André/SP - Brazil

Marcelo Araujo Silva - RM Soluções Engenharia - São Paulo/SP - Brazil

Marcelo Curvo - Embraer - São José dos Campos/SP - Brazil

Marcelo José Ruv Lemes - Embraer - São José dos Campos/SP - Brazil

Marcio Teixera de Mendonça- Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Marco Antônio Ferraz - Embraer- São José dos Campos/SP - Brazil

Marcos Massi- Instituto Tecnológico de Aeronáutica- São José dos Campos/SP - Brazil

Maria Cecília Zanardi - Faculdade de Engenharia de Guaratinguetá - Guaratinguetá/SP - Brazil

Maurício Vicente Donadon- Instituto Tecnológico de Aeronáutica- São José dos Campos/SP - Brazil

Miriam Kasumi - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Nicelio José Lourenço - Instituto de Aeronáutica e Espaço- São José dos Campos/SP - Brazil

Nilson C Cruz - Universidade Estadual Paulista - Sorocaba/SP - Brazil

Odylio Denys de Aguiar - Instituto Nacional de Pesquisa Espaciais - São José dos Campos/SP - Brazil

Paulo César Pellanda - Instituto Militar de Engenharia - Rio de Janeiro/RJ - Brazil

Paulo Gilberto de Paula Toro - Instituto de Estudos Avançados - São José dos Campos/SP - Brazil

Paulo Roberto Bergamaschi - Universidade Federal de Goiás - Catalão/GO - Brazil

Paulo Zavala - Universidade Estadual de Campinas - Campinas/SP - Brazil

Pierre Kaufmann - Universidade Mackenzie - São Paulo/SP - Brazil

Ricardo Elgul Samad - Instituto de Pesquisas Energéticas e Nucleares - São Paulo/SP - Brazil

Robero Ramos - Universidade Federal do ABC - Santo André/SP - Brazil

Roberto Gil Annes da Silva - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Roberto Lopes - Instituto Nacional de Pesquisa Espaciais - Cachoeira Paulista/SP - Brazil

Roberto Mendes Finzi Neto- Universidade Federal de Uberlândia - Uberlândia/MG - Brazil

Roberto Vasconcelos - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Rogéria Cristiane Gratão de Souza- Universidade Estadual Paulista - São José do Rio Preto/SP - Brazil

Rogeria Eller- Instituto Tecnológico de Aeronáutica- São José dos Campos/SP - Brazil

Rogério Mota- Faculdade de Engenharia de Guaratinguetá- Guaratinguetá/SP - Brazil


Sonia A. Goulart de Oliveira - Universidade Federal de Uberlândia- Uberlândia/MG - Brazil

Tony Springer - NASA Headquarters - Washington DC - USA

Valdemir Carrara - Instituto Nacional de Pesquisa Espaciais - São José dos Campos/SP - Brazil

Vinicius André R.Henriques - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil

Wesley Gois - Universidade Federal do ABC - Santo André/SP - Brazil

William W. Vaughan - National Space Science and Technology Center- Huntsville/AL - USA

Wilson F. N. Santos - Instituto Nacional de Pesquisa Espaciais - Cachoeira Paulista/SP - Brazil

Wilson Shimote - Instituto de Aeronáutica e Espaço - São José dos Campos/SP - Brazil


INSTRUCTIONS TO AUTHORS(Revised in March, 2012)

SCOPE AND EDITORIAL POLICY

Departamento de Ciência e Tecnologia Aeroespacial (DCTA), in São José dos Campos, São Paulo State, Brazil. The journal is quarterly

selected from the following topics:

PAPER SUBMISSION

of the manuscript, the author must send a declaration indicating the reasons so that the review process occurs fairly.

allowed, it is mandatory to mention the JATM. The author also declares that the manuscript is an original paper, its content is not

Instituto de Aeronáutica e Espaço

articles from other journals.


PEER REVIEW

approve after verifying in the new version the adherence to the reviewers’ suggestions or will send to another evaluation round

MANUSCRIPT CATEGORIES

Editorial:

Review articles:

Original papers:

Communications: They should report previous results of ongoing research and should not exceed eight pages.

Thesis abstracts:

MANUSCRIPT STRUCTURE

usual format.

Title and names of authors:

Abstract:

not permitted.


Keywords:

Introduction:

Methodology:

Results:

Discussion:

Acknowledgements:

References:

and completeness of their references.

References in text:

names, use the form “Smith et al

they do not need to appear in the reference list.

Reference list:

are as the following ones:


Tables:

Figures:

Equations:


Correspondence

All correspondence should be sent to:JOURNAL OF AEROSPACE TECHNOLOGY AND MANAGEMENT

Instituto de Aeronáutica e EspaçoPraça Mal. Eduardo Gomes, 50- Vila das Acácias

CEP 12228-901São José dos Campos/São Paulo/Brazil

Contact

Phone: (55) 12-3947- 6493/5122E-mail: [email protected]: http://www.jatm.com.br

Published by: Departamento de Ciência e Tecnologia Aeroespacial Distributed by: Instituto de Aeronáutica e Espaço

Proofreading and standardization: Zeppelini EditorialEditing: TL Publicidade e Assessoria Ltda

Cover: Created by Gus Morais-based on the original by Paulo R. Porphirio MoreiraEdition: 750

São José dos Campos, SP, BrazilISSN 1984-9648

Historical Note: JATM was created in 2009 after the iniciative of the diretor of Instituto de Aeronáutica e Espaço (IAE), Brigadeiro Engenheiro Francisco Carlos Melo Pantoja. In order to reach the goal of becoming a journal that could represent

from the beginning. From September 2011, it has been edited by the Departamento de Ciência e Tecnologia Aeroespacial

The copyright on all published material belongs toDepartamento de Ciência e Tecnologia Aeroespacial (DCTA)

Journal of Aerospace Technology and Managementwww.jatm.com.br

vol. 4 n.4 - journal of aerospace technology and management

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