This paper describes the changes made from Cal Poly’s initial railgun system, the Mk. 1 railgun, to the Mk. 1.1 system, as well as the design, fabrication, and testing of a newer and larger Mk. 2 railgun system. The Mk. 1.1 system is developed as a more efficient alteration of the original Mk. 1 system, but is found to be defective due to hardware deficiencies and failure, as well as unforeseen efficiency losses. A Mk. 2 system is developed and built around donated hardware from the Naval Postgraduate School. The Mk. 2 system strove to implement an efficient, augmented, electromagnetic railgun and projectile system capable of firing an approximate 1g aluminum projectile to speeds exceeding 2 km/s. A novel three part projectile is proposed to mitigate rail and projectile degradation. Projectile and sabot system kinematic equations are derived and the projectile is designed and tested along with Mk. 2 barrel. A numerical electromechanical model is developed to predict the performance of the Mk. 2 system and projectile assembly, and predicts a final velocity for the fabricated system exceeding 3.5 km/s and an efficiency as high as 24%. Testing of the Mk. 2 system showed catastrophic failure of the projectile during initial acceleration, resulting in very short acceleration times and distance, low velocity projectiles, and low efficiencies. During further testing of various projectile configurations, the barrel structure failed due to a large internal arc. Future work for the Mk. 2 system is discussed, a revised external barrel structure suggested, and a solid, more conventional solid chevron projectile design suggested.

Collision with orbital debris poses a serious threat to spacecraft and astronauts. Hypervelocity impacts resulting from collisions mean that objects with a mass less than 1g can cause mission-ending damage to spacecraft. A means of shielding spacecraft against collisions is necessary. A means of testing candidate shielding methods for their efficacy in mitigating hypervelocity impacts is therefore also necessary. Cal Poly’s Electromagnetic Railgun was designed with the goal of creating a laboratory system capable of simulating hypervelocity (≥ 3 km/s) impacts. Due to several factors, the system was not previously capable of high-velocity (≥ 1 km/s) tests. A deficient projectile design is revised, and a new design is tested. The new projectile design is demonstrated to enable far greater performance than the previous design, with a muzzle velocity ≥ 1 km/sbeing verified during testing, and an energy conversion efficiency of 2.7%. A method of improving contact and controlling wear at the projectile/rail interface using silver plating and conductive silver paste is validated. A mechanism explaining the problem of internal arcing within the railgun barrel is proposed, and design recommendations are made to eliminate arcing on the basis of the work done during testing. The primary structural members are found to be deficient for their application and a failure analysis of a failed member, loading analysis of the railgun barrel, and design of new structures is undertaken and presented.

The re-entry space vehicles encounter high temperatures when they enter the earth atmosphere and the high temperature air in the shock layer around the body undergoes partial dissociation. Also, the gas molecules injected into the shock layer from the ablative thermal protection system (TPS) undergo pyrolysis which helps in reducing the net heat flux to the vehicle surface. The chemical species due to the pyrolysis add complexity to the stagnation flow chemistry (52 chemical reactions) models which include species like NOx, CO and hydrocarbons (HCs). Although the ablative TPS is responsible for the safety of re-entry space vehicle, the induced chemical species result in variety of adverse effects on environment such as global warming, acid rain, green house effect etc. The well known three-way-catalyst (TWC) involves simultaneous removal of all the three gases (i.e, NOx, CO, Hydrocarbons) present in the shock layer. Interaction of such three-way-catalyst on the heat shield materials or on the wall of the re-entry space vehicle is to reduce the heat flux and to remove the gases in the shock layer, which is an important issue. For the re-entry vehicle the maximum aerodynamic heating occurs at an altitude ranging about 68 to 45 km during which the vehicle is surrounded by high temperature dissociated air. Then the simplest real gas model of air is the five species model which is based on N2, O2, O, NO and N. This five species model assumes no ionization and no pyrolysis gases are emitted from the heat shield materials. The experimental research work presented in this thesis is directed towards the understanding of catalytic and non-catalytic surface reactions on high temperature materials in presence of strong shock heated test gas. We have also explored the possibility of using shock tube as a high enthalpy device for synthesis of new materials. In the first Chapter, we have presented an overview of re-entry space vehicles, thermal protection system (TPS) and importance of real gas effects in the shock layer. Literature survey on TPS, ablative materials and aerothermochemistry at the stagnation point of reentry capsule, in addition to catalytic and non-catalytic surface reactions between the wall and dissociated air in the shock layer are presented. In Chapters 2 and 3, we present the experimental techniques used to study surface reactions on high temperature materials. A brief description of HST2 shock tunnel is presented and this shock tunnel is capable of generating flow stagnation enthalpies ranging from 0.7 to 5 MJ/kg and has an effective test time of ~ 800 µs. High speed data acquisition system (National Instruments and Yokogawa) used to acquire data from shock tube experiments. The experimental methods like X-ray Photoelectron Spectroscopy (XPS), Scanning Electron Microscopy (SEM), X-ray diffraction (XRD), Raman and FTIR spectroscopy have been used to characterize the shock-exposed materials. Preliminary research work on surface nitridation of pure metals with shock heated nitrogen gas is discussed in Chapter 2. Surface nitridation of pure Al thin film with shock heated N2 is presented in Chapter 3. An XPS study shows that Al 2p peak at 74.2 eV is due to the formation AlN on the surface of Al thin film due to heterogeneous non-catalytic surface reaction. SEM results show changes in surface morphology of AlN film due to shock wave interaction. Thickness of AlN film on the surface increased with the increase in temperature of the shock heated nitrogen gas. However, HST2 did not produce sufficient temperature and pressure to carry out real conditions of re-entry. Therefore design and development of a new high enthalpy shock tunnel was taken up. In Chapter 4, we present the details of design and fabrication of free piston driven shock tunnel (FPST) to generate high enthalpy test gas along with the development of platinum (Pt) and thermocouple sensors for heat transfer measurement. A free piston driven shock tunnel consists of a high pressure gas reservoir, compression tube, shock tube, nozzle, test section and dump tank connected to a vacuum pumping system. Compression tube has a provision to fill helium gas and four ports, used to mount optical sensors to monitor the piston speed and pressure transducer to record pressure at the end of the compression tube when the piston is launched. Piston can attain a maximum speed of 150 m/s and compress the gas inside the compression tube. The compressed gas bursts the metal diaphragm and generates strong shock wave in the shock tube. This tunnel produces total pressure of about 300 bar and temperature of about 6000 K and is capable of producing a stagnation enthalpy up to 45 MJ/kg. The calibration of nozzle was carried out by measuring the pitot tube pressure in the dump tank. Experimentally recorded P5 pressure at end of the shock tube is compared with Numerical codes. Calibrated pressure P5 values are used to calculate the temperature T5 of the reflected shock waves. This high pressure and high temperature shock heated test gas interacts with the surface of the high temperature test materials. For the measurement of heat transfer rate, platinum thin film sensors are developed using DC magnetron sputtering unit. Hard protective layer of aluminum nitride (AlN) on Pt thin film was deposited by reactive DC magnetron sputtering to measure heat transfer rate in high enthalpy tunnel. After the calibration studies, FPST is used to study the heat transfer rate and to investigate catalytic/non-catalytic surface reaction on high temperature materials. In Chapter 5, an experimental investigation of non-catalytic surface reactions on pure carbon material is presented. The pure carbon C60 films and conducting carbon films are deposited on Macor substrate in the laboratory to perform shock tube experiments. These carbon films were exposed to strong shock heated N2 gas in the shock tube portion of the FPST tunnel. The typical shock Mach number obtained is about 7 with the corresponding pressure and temperature jumps of about 110 bar and 5400 K after reflection at end of the shock tube. Shock exposed carbon films were examined by different experimental techniques. XPS spectra of C(1s) peak at 285.8 eV is attributed to sp2 (C=N) and 287.3 eV peak is attributed to sp3 (C-N) bond in CNx due to carbon nitride. Similarly, N(1s) core level peak at 398.6 eV and 400.1 eV observed are attributed to sp3-C-N and sp2-C=N of carbon nitride, respectively. SEM study shows the formation of carbon nitride crystals. Carbon C60 had melted and undergone non-catalytic surface reaction with N2 while forming carbon nitride. Similar observations were made with conducting carbon films but the crystals were spherical in shape. Micro Raman and FTIR study gave further evidence on the formation of carbon nitride film. This experimental investigation confirms the formation of carbon nitride in presence of shock-heated nitrogen gas by non-catalytic surface reaction. In Chapters 6 and 7, we present a novel method to understand fully catalytic surface reactions after exposure to shock heated N2, O2 and Ar test gas with high temperature materials. We have employed nano ZrO2 and nano Ce0.5Zr0.5O2 ceramic high temperature materials to investigate surface catalytic reactions in presence of shock heated test gases. These nano crystalline oxides are synthesized by a single step solution combustion method. Catalytic reaction was confirmed for both powder and film samples of ZrO2. As per the theoretical model, it is known that the catalytic recombination reaction produces maximum heating on the surface of re-entry space vehicles. This was demonstrated in this experiment when a metastable cubic ZrO2 changed to stable monoclinic ZrO2 phase after exposure to shock waves. The change of crystal structure was seen using XRD studies and needle type monoclinic crystal growth with aspect ratio (L/D) more than 15 was confirmed by SEM studies. XPS of Zr(3d) core level spectra show no change in binding energy before and after exposure to shock waves, confirming that ZrO2 does not change its chemical nature, which is the signature of catalytic surface reaction. When a shock heated argon gas interacted with Ce0.5Zr0.5O2 compound, there was a change in colour from pale yellow to black due to reduction of the compound, which is the effect of heat transfer from the shock wave to the compound in presence of argon gas. The reduction reaction shows the release of oxygen from the compound due to high temperature interaction. The XPS of Ce(3d) and Zr(3d) spectra confirm the reduction of both Ce and Zr to lower valent states. The oxygen storage and release capacity of the Ce0.5Zr0.5O2 compound was confirmed by analyzing the reduction of Ce4+ and Zr4+ with high temperature gas interaction. When Ce0.5Zr0.5O2 (which is same as Ce2Zr2O8) in cubic fluorite structure was subjected to strong shock, it changed to pyrochlore (Ce2Zr2O7) structure by releasing oxygen and on further heating it changed to Ce2Zr2O6.3 which is also crystallized in pyrochlore structure by further releasing oxygen. If this heating is carried out in presence of argon test gas, fluorite structure can easily change to pyrochlore Ce2Zr2O6.3 structure, which is a good electrical conductor. Due to its oxygen storage capability (OSC) and redox (Ce4+/Ce3+) properties, Ce0.5Zr0.5O2 had been used as oxygen storage material in three-way-catalyst. Importance of these reactions is that the O2 gas released from the compound will react with gas released from the heat shield materials, like NOx, CO and hydrocarbon (HCs) species which results in reduction of temperature in the shock layer of the re-entry space vehicle. The compound Ce0.5Zr0.5O2 changes its crystal structure from fluorite to pyrochlore phase in presence of shock heated test gas. The results presented in these two Chapters are first of their kind, which demonstrates the surface catalytic reactions. In Chapter 8, we present preliminary results of the oxygen recombination on the surface of heat shield material procured from Indian Space Research Organization (ISRO) used as TPS in re-entry space capsule (Space capsule Recovery Experiment SRE-1) and on thin film SiO2 deposited on silicon substrate. The formation of SiO between the junctions of SiO2/Si was confirmed using XPS study when shock exposed oxygen reacted on these materials. The surface morphology of the ablated SiO2 film was studied using SEM. The damage induced due to impact of shock wave in presence of oxygen gas was analyzed using Focused Ion Beam (FIB) microscope. The results reveal the damage on the surface of SiO2 film and also in the cross-section of the film. We are further investigating use of FIB, particularly related to residual stress developed on thin films due to high pressure and high temperature shock wave interaction. In Chapter 9, conclusions on the performance of FPST, synthesis of high temperature materials, catalytic and non-catalytic surface reactions on the high temperature material due to shock-heated test gases are presented. Possible scope for future studies is also addressed in this Chapter.

Thesis (M.S.)--Air Force Institute of Technology, 2003. / Title from title screen (viewed July 1, 2004). "March 2003." Vita. "AFIT/GCA/ENV/03-04." "ADA415257"--URL. Includes bibliographical references (p. 71-74). Also issued in paper format.

The present research is an experimental and numerical study of aeroacoustic phenomena occurring in large solid rocket motors (SRM) as the Ariane 5 boosters. The emphasis is given to aeroacoustic instabilities that may lead to pressure and thrust oscillations which reduce the rocket motor performance and could damage the payload. The study is carried out within the framework of a CNES (Centre National d'Etudes Spatiales) research program. <p><p>Large SRM are composed of a submerged nozzle and segmented propellant grains separated by inhibitors. During propellant combustion, a cavity appears around the nozzle. Vortical flow structures may be formed from the inhibitor (Obstacle Vortex Shedding OVS) or from natural instability of the radial flow resulting from the propellant combustion (Surface Vortex Shedding SVS). Such hydrodynamic manifestations drive pressure oscillations in the confined flow established in the motor. When the vortex shedding frequency synchronizes acoustic modes of the motor chamber, resonance may occur and sound pressure can be amplified by vortex nozzle interaction.<p><p>Original analytical models, in particular based on vortex sound theory, point out the parameters controlling the flow-acoustic coupling and the effect of the nozzle design on sound production. They allow the appropriate definition of experimental tests.<p><p>The experiments are conducted on axisymmetric cold flow models respecting the Mach number similarity with the Ariane 5 SRM. The test section includes only one inhibitor and a submerged nozzle. The flow is either created by an axial air injection at the forward end or by a radial injection uniformly distributed along chamber porous walls. The internal Mach number can be varied continuously by means of a movable needle placed in the nozzle throat. Acoustic pressure measurements are taken by means of PCB piezoelectric transducers. A particle image velocimetry technique (PIV) is used to analyse the effect of the acoustic resonance on the mean flow field and vortex properties. An active control loop is exploited to obtain resonant and non resonant conditions for the same operating point.<p><p>Finally, numerical simulations are performed using a time dependent Navier Stokes solver. The analysis of the unsteady simulations provides pressure spectra, sequence of vorticity fields and average flow field. Comparison to experimental data is conducted.<p><p>The OVS and SVS instabilities are identified. The inhibitor parameters, the chamber Mach number and length, and the nozzle geometry are varied to analyse their effect on the flow acoustic coupling.<p><p>The conclusions state that flow acoustic coupling is mainly observed for nozzles including cavity. The nozzle geometry has an effect on the pressure oscillations through a coupling between the acoustic fluctuations induced by the cavity volume and the vortices travelling in front of the cavity entrance. When resonance occurs, the sound pressure level increases linearly with the chamber Mach number, the frequency and the cavity volume. In absence of cavity, the pressure fluctuations are damped.<p><p> / Doctorat en sciences appliquées / info:eu-repo/semantics/nonPublished

Tato diplomová práce je zaměřena na sestavení přehledu konstrukcí a teplených cest kosmických lodí a kosmických vozidel, které se v současné době používají. Na základě specifických požadavků a standardů jsou vypracovány koncepční návrhy zástavby tepelného spínače do konstrukce družice. V jednotlivých kapitolách jsou popsány určité členy, které se na vedení tepla podílí a jsou důležité pro navrhované koncepty z pohledu konstrukce. Diplomová práce popisuje postupy výpočtu tepelných vodivostí a rozložení působící tíhové síly do míst uchycení pro jednotlivé koncepty. Na závěr provádí hodnocení a výběr potenciálně nejvhodnějšího návrhu.

Contrairement aux autres biens meubles, les navires, les aéronefs et les objets spatiaux affectés à la navigation internationale sont rattachés à un Etat. Le lien de droit public établi entre ces engins et l’Etat est communément appelé « nationalité ». Mais ce terme n’exprime pas à leur propos une institution à tous égards identique à la nationalité des personnes. Le rattachement examiné ne repose en effet pas sur des éléments de fait (naissance, ascendance etc.), mais uniquement sur un acte administratif interne, l’immatriculation. L’étude de la pratique, notamment des conventions internationales et des législations nationales, montre clairement que – contrairement à ce qu’on soutient souvent – il n’y a pas lieu de subordonner ce rattachement à un lien effectif. Ce qui importe, compte tenu notamment du fait que ces engins évoluent dans des espaces soustraits à toute compétence territoriale, est d’identifier l’Etat qui est seul compétent à l’égard de l’« ensemble organisé » formé par le véhicule, les personnes et la cargaison à bord, et qui est responsable de ses activités. Le droit international interdit dès lors la double immatriculation, mais il laisse aux Etats le pouvoir discrétionnaire de déterminer les conditions d’attribution de leur « nationalité », sans subordonner l’opposabilité internationale de celle-ci à quelque autre exigence que ce soit. Le danger est toutefois que cela favorise un certain laxisme de l’Etat d’immatriculation, ce qui exposerait au risque que des dommages graves soient causés aux personnes impliquées dans les activités de ces engins et – surtout – aux tiers. Mais ce sont les obligations internationales imposées et les droits corrélatifs reconnus dans le chef de l’Etat d’immatriculation qui sont déterminants à cet égard et non quelque mystérieuse « effectivité » du rattachement. Autrement dit, s’il n’est pas nécessaire d’imposer à l’Etat d’immatriculation des conditions internationales limitant sa liberté dans l’attribution de sa « nationalité » aux engins, il est indispensable d’exiger que celui-ci respecte ses obligations, c’est-à-dire exerce effectivement son contrôle et sa juridiction. Cette constatation se vérifie quel que soit l’engin en cause. Le rattachement créé par l’immatriculation constitue donc une institution "sui generis", commune aux navires, aéronefs et objets spatiaux et dont le régime juridique est encadré par le droit international. / Unlike any other movable property, ships, aircraft and space objects that are engaged in international navigation are linked to a State. The legal connection established between these craft/vessels and the State is commonly referred to as “nationality”. However, in this case the term does not represent an institution identical in all respects to the nationality of persons. With regard to vessels, the legal connection to a State is not based on factual elements (such as birth, descent etc.), but merely on the internal administrative act of registration. The study of State practice, notably international conventions and national laws, clearly shows that – contrary to what is often argued – there is no need to make this connection dependent on a pre-existing effective link. What matters most, given that these craft navigate in international space beyond the territorial jurisdiction of sovereign States, is to identify the State that holds sole jurisdiction over said “organized entity” consisting of the vehicle, the persons and the cargo on board and that is responsible for its activities. Public international law therefore prohibits dual registration, but leaves States free to determine the conditions under which they will confer their “nationality”, without imposing any other requirement for the opposability of this legal bond to third States. The danger is that this situation encourages laxity on the part of the States of registry and therefore creates the potential for serious damage incurred by persons involved in these vessels’ activities and – mostly – by third persons. In this regard, it is the international obligations and corresponding rights of the States of registry which are critical, and not a mysterious “effectiveness” of the legal bond. In other words, it is not necessary to impose on the State of registry any international conditions which would limit its freedom with regard to the conferral of its “nationality” upon vessels. It is however indispensable to require that said State complies with its obligations, meaning that it has to effectively exercise its jurisdiction and control over those craft. This statement holds true regardless of the craft concerned. The legal bond created by the registration therefore constitutes a "sui generis" institution, common to ships, aircraft and space objects, and whose legal regime is governed by international law.