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  • Spacecraft Propulsion and Power  (208)
  • Biology
  • Cell & Developmental Biology
  • Inorganic Chemistry
  • 2005-2009  (216)
  • 1935-1939
  • 2005  (216)
  • 1
    Publication Date: 2021-05-19
    Description: Etude réalisée dans le cadre des activités de recherche de l’Institut National des Sciences et Technologies de la Mer (INSTM)
    Description: Unpublished
    Keywords: Répartition sexuelle ; Thon rouge ; Biométrie morphologique ; Scombridés ; Ressources pélagiques ; Biologie ; Pelagic fish ; Biology ; Biometrics
    Repository Name: AquaDocs
    Type: Theses and Dissertations , Master thesis
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  • 2
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    Alexandria: NIOF
    Publication Date: 2021-05-19
    Description: In a second study of electromicroscopic analysis of ultrastructural organelles of coho salmon muscle showed a significant increase in lipid droplet density in the largest stages compared with the smallest fish. The opposite was observed with the mitochondria where its density decreased in the large sized fish. An apparent decrease in capillary density with increasing body mass was also observed. These finding may imply a possible additional role of lipid depot as fish grew from parr to smolt.
    Description: Published by NIOF, Alexandria
    Description: Published
    Keywords: Ultrastructure ; Cohosalmon ; Lipid droplets ; Salmon ; Biology ; Fish
    Repository Name: AquaDocs
    Type: Journal Contribution , Non-Refereed , Article
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  • 3
    Publication Date: 2021-05-19
    Description: Study of microscopic structural changes of red and white muscle of coho salmon, Oncorhynchus kisutch, through different developmental stages and through transition from parr to smolt revealed developmental variation in fibre size where white fibres were larger than red ones. The increase in fibre area, characterizing all stages of development, was associated with reduction in both fibre and capillary densities. In large fish the reduction in red muscle capillary density was 10 fold lesser than that of white muscle.
    Description: Published by NIOF, Alexandria
    Description: Published
    Keywords: Oncorhynchus kisutch ; White fibres ; Red fibres ; Biology ; Salmon ; Biology ; Fish
    Repository Name: AquaDocs
    Type: Journal Contribution , Non-Refereed , Article
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  • 4
    Publication Date: 2021-05-19
    Description: Il y a des annexes
    Description: Résumé : La biologie de la langouste rose des côtes mauritaniennes est aujourd’hui bien connue. Nous avons cependant rappelé les traits marquants. Sur le plan de la pêche, de 1963 à 1988, le stock de la langouste rose a connu trois phases d’exploitation. Après une phase de surexploitation entre 1963 et 1970-1971, le stock de langouste rose de Mauritanie a pu se reconstituer grâce à une réduction importante du nombre de langoustiers. Ceci c’est traduit à partir de 1971 par une augmentation sensible des captures et aussi de la PUE. Mais le dédoublement du nombre de ces derniers en 1987 et 1988 a fait chuter les P.U.E de 50% ce qui est un signe d’effondrement du stock à nouveau.
    Description: Abstract : Biology of Mauritanian pink lobster is known. However we presented some important aspects of specie’s biology. On fisherie’s plan, between 1963 and 1988, the stock of Mauritanian pink lobster went through three phases of exploitation. Like this, after an over exploitation phase between 1963 and 1970-1971. The stock of the pink lobster of Mauritania was reconstituted little by little owing to an important reduction enabled a significant increase in the number of catches and the C.P.U.E. But doubling in the number of beats in 1987 and 1988 reduced the C.P.U.E. by 50%, which again contributed to the breaking down of stock.
    Description: IMROP
    Description: Published
    Keywords: Pêche ; Etat des stocks ; Mauritanie ; Biologie ; Fishing ; Langouste rose ; Biology ; Palinurus mauritanicus ; Mauritania ; Status of stocks ; Pink lobster ; Lobster culture
    Repository Name: AquaDocs
    Type: Journal Contribution , Non-Refereed , Article
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  • 5
    Publication Date: 2021-05-19
    Description: Thèse pour l’obtention du grade de Docteur de l’Université de Bretagne Occidentale spécialité océanologie biologique
    Description: RESUME: Les eaux océaniques mauritaniennes sont au carrefour d'eaux froides et salées provenant du Nord, et d'autres chaudes et moins salées du Sud. L'interaction de ces eaux est à l'origine d'un régime hydrologique à 4 saisons: une saison froide de janvier à mai, une saison de transition froide-chaude de juin à juillet, une saison chaude d'août à octobre et enfin une saison de transition chaude – froide de novembre à décembre. La présente étude met en relation ces variations hydrologiques saisonnières, la distribution et la biologie de Mustelus mustelus. La distribution spatiale et temporelle de M. mustelus sur le plateau continental mauritanien a été suivie en analysant les données récoltées au cours de 20 campagnes de prospections scientifiques, 13 hauturières et 7 côtières, réparties sur les 4 saisons hydrologiques qui caractérisent cette zone géographique. La population d'émissoles est concentrée au Nord du Cap Timiris. L'espèce ne semble pas effectuer de migration latitudinale, mais un déplacement côte - large et inversement a été mis en évidence. Durant la saison froide, cette espèce à affinité tropicale est repoussée vers la côte, dans la Baie du Lévrier et le Nord Est du Banc d'Arguin par les basses températures venant du Nord qui persistent sur le plateau continental mauritanien de janvier à mai. Ce sont surtout les mâles matures qui rejoignent alors, pour l'accouplement, les femelles, de distribution plus côtière. Entre juin et octobre, avec le réchauffement des eaux, les mâles commencent à se déplacer vers des eaux plus profondes. Mais dès le début du refroidissement, un nouveau mouvement vers la côte s'amorce. Un échantillonnage mensuel des débarquements de la Pêche Artisanale a permis de collecter, en 2 ans, 2510 séries de données individuelles exploitées pour les études de l'alimentation, de la reproduction et de la croissance. L'étude des contenus stomacaux n'a pas montré de différence significative entre les femelles et les mâles, l'analyse a été faite sans distinction des sexes. Le nombre d'estomacs vides est de 10,3 % des estomacs examinés. La distribution Nord de M. mustelus pourrait être liée à l'abondance des principales proies de l'espèce, les bernard-l'hermites (Anomoures). En effet, ces proies qui seraient abondantes dans sa zone de répartition sont dominantes aussi bien en termes de nombre que de poids et d'occurrence dans les contenus des estomacs examinés. Les proies secondaires sont des Poissons, des Mollusques et des Annélides. Ainsi, l'émissole serait une espèce opportuniste qui se nourrit principalement de proies les plus vulnérables et accessoirement de proies de capture plus difficiles. L'étude du régime alimentaire chez l'émissole lisse en Mauritanie témoigne d'un comportement en rapport étroit avec le fond. La distribution différentielle des deux sexes influence le sex ratio qui est en faveur des femelles dans la zone côtière et des mâles dans la zone hauturière. La taille de première maturité sexuelle de M. mustelus en Mauritanie est de 67 cm pour les mâles et 72 cm pour les femelles. Chez les femelles, l'activité vitellogénique est continue et dure toute l'année chez les femelles matures (y compris gestantes) sauf en période de fécondation, période pendant laquelle elle est arrêtée. Le nombre d'ovocytes vitellogèniques de grand diamètre (supérieur ou égal à 10 mm) atteint un maximum en mai dans l'unique ovaire droit présent chez les émissoles lisses. La chute de leur nombre dans l'ovaire en juin et juillet annonce l'ovulation. Les mâles s'accouplent avec les femelles entre janvier et mai; les spermatozoïdes sont alors stockés dans le tiers inférieur des glandes nidamentaires jusqu'à la période de fécondation (juillet-août). L'organogenèse dure jusqu'aux mois d'octobre – novembre; elle aboutit alors à des embryons qui ressemblent morphologiquement aux adultes, au-delà de ces mois, le développement embryonnaire se limite à une augmentation de taille et de poids. Les embryons sortent de leurs capsules, les réserves vitellines se résorbent et sont remplacées par le placenta. Chez les femelles, la durée de la gestation est de 7 à 10 mois. Les femelles commencent à mettre bas à partir de février, la parturition se poursuit jusqu'en juin, mois pendant lequel les rares femelles encore gestantes mettent bas. A la naissance les juvéniles ont des tailles comprises entre 240 et 320 mm. La fécondité utérine maximale observée au cours de cette étude est 13 embryons; la moyenne est de 4 embryons par portée. Les bandes de croissances observées dans des coupes de vertèbres de femelles de 45 à 99 cm et de mâles de 50 à 85 cm de longueur totale ont été utilisées pour l'estimation de l'âge des poissons. Les données d'âge et de longueur ont permis d'établir les équations de croissances selon les modèles de Von Bertalanffy et de Holden. Ces modèles conduisant à des résultats différents; celui de Von Bertalanffy a été retenu en raison de sa flexibilité qui rend son application plus courante dans les pêcheries. Les paramètres de ce modèle sont, pour les femelles K=0,21, L∞=113,4 et t0=-2,03, pour les mâles K=0,26; L∞=91,3 et t0= -2,43. Les femelles ont donc des croissances plus rapides que les mâles et les âges à la première maturité sexuelle sont de 2,6 ans pour les mâles et de 2,8 pour les femelles. #
    Description: ABSTRACT: Mauritanian coastal waters are crossroads between cold and salted northern waters, and warmer and less salted southern waters. The interaction of these waters is at the origin of four hydrological seasons: cold (January to May), cold-to-warm transition (June-July), warm (August to October) and warm-to-cold transition (November-December). The present study connects these seasonal hydrological variations with the distribution and biology of Mustelus mustelus. Spatial and temporal distribution of M. mustelus on the Mauritanian continental shelf was followed while analyzing data collected during 20 scientific campaigns, 13 deep-sea and 7 coastal, distributed over the 4 hydrological seasons which characterize this geographical area. Smoothound sharks population is concentrated North of Cape Timiris. The species does not seem to carry out of latitudinal migration, but a coast to deep sea displacement - and conversely a deep sea to coast - was found. During the cold season, this species with tropical affinity is pushed back towards the coast, in Baie du Lévrier and Northeast of the Banc d’Arguin by cold water temperatures coming from the North which persist on the Mauritanian continental shelf from January to May. Mature males then join, for the coupling, females, which are of more coastal distribution. Between June and October, with the reheating of water, males start to move towards deeper water. But from the very start of cooling, a new movement towards the coast starts. A monthly sampling of the artisanal fishery catches made it possible to collect, in 2 years, 2,510 individual specimen data for the study of food, reproduction and growth. The study of stomach contents did not show a significant differences between females and males. The analysis was made regardless of gender. The number of empty stomachs is 10,3 % of the examined stomachs. The Northern distribution of M. mustelus could be related to the abundance of the principal prey species, the hermit crab (Anomoura). Indeed, these preys which would be abundant in its zone of distribution are dominant in terms of number, biomass and occurrence in the contents of the examined stomachs. The secondary preys are of fish, molluscs and annelids. Thus, smoothound shark would be an opportunist species which feeds mainly on the most vulnerable preys and incidentally on more difficult preys of capture. The study of the smoothound shark diet in Mauritania testifies to a behavior in close connection with the bottom. The differential distribution of the two sexes influences the sex ratio which is in favor of females in the coastal zone and males in the deep-sea zone. Size at first sexual maturity of M. mustelus in Mauritania is 67 and 72 cm for males and females, respectively. For females, the vitellogenic activity is continuous year-round for mature females (including gestating ones), except during the period of fecundation. The number of vitellogenic ovocytes of large diameter (greater or equal to 10 mm) reaches a maximum in May in the single right ovary of the smoothound shark. The fall of their number in June and July announces ovulation. Males couple themselves with females between January and May; spermatozoids are then stored in the nidamental glands lower third until the period of fecundation (July-August). Organogenesis lasts until October - November; it then leads to embryos which resemble to the adults morphologically. Afterwards, the embryonic development is limited to an increase in size and weight. The embryos leave their capsules, vitellogenic reserves reabsorb and are replaced by the placenta. Gestation period is 7 to 10 months for females. The females start to put low as from February, parturition continues until June, month during which the rare still gestating females give birth. With birth the youthful ones have sizes ranging between 240 and 320 mm. The maximum observed fecundity during this study is 13 embryos; the average is of 4 embryos per litter. Growth bands observed in cuts of vertebrae on females and males of total length ranging from 45 to 99 cm and from 50 to 85 cm, respectively, were used to estimate fish age. Age and length data made it possible to establish growth equations according to Von Bertalanffy and Holden’s models. These models lead to different results; Von Bertalanffy model was retained because of its flexibility which makes its application more current in fisheries. The parameters of this model are K=0.21, L∞=113.4 for females and t0= -2.03, and K=0.26 L∞=91.3 and t0 = -2.43 for males. The females tend to grow faster than males and the age at the first sexual maturity is 2.6 and 2.8 years for males and females, respectively.
    Description: IMROP
    Description: Unpublished
    Keywords: Mauritania ; Croissance ; Biologie ; Biology ; Physical environment ; Mauritanie ; Ecology ; Fecundity ; Growth ; Emissole lisse ; Alimentation ; Smouthound shark ; Mustelus mustelus ; Fécondité ; Reproduction ; Milieu physique ; Ecologie ; Distribution ; Shark fisheries
    Repository Name: AquaDocs
    Type: Theses and Dissertations , Master thesis
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  • 6
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    CNROP - AtlantNIRO -MEIPP-sa, 2001 plaquette de vulgarisation
    Publication Date: 2021-05-19
    Description: Le présent prospectus donne une courte caractéristique biologique des principales espèces pêchées dans la Zone Economique Exclusive Mauritanienne et des engins utilisés pour la pêche, ainsi que les informations relatives à leur composition chimique, à la valeur nutritive de la matière première, et à ses caractéristiques avec les possibilités de sa transformation en différents produits. Les informations présentées pourront être utiles aux producteurs pour le choix de la matière première et des possibilités de sa valorisation
    Description: IMROP / AtlantNIRO / MEIPP-sa
    Keywords: Fishery resources ; Valorisation produits ; Composition chimique ; Biologie ; Ressources halieutiques ; Biology ; Valeur nutritive ; Fish processing ; Transformation des produits ; Engins de pêche ; Food value ; Fishing gear ; Fish species ; Processing fishery products
    Repository Name: AquaDocs
    Type: Other
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  • 7
    Publication Date: 2021-05-19
    Description: L’étude porte sur l’exploitation, la biologie et la dynamique du poulpe (Octopus vulgaris, Cuvier) des eaux mauritaniennes. La répartition spatio-temporelle dans la région du Cap Blanc ne semble être liée ni à la taille, ni au sexe, mais varie au cours de l’année. L’analyse de l’état actuel de la pêcherie basée sur les congélateurs céphalopodiers met en évidence une augmentation de l’effort de pêche et la poursuite de chute des p.u.e. La composition des captures révèle une progression du pourcentage des individus de poids inférieur à 200 grammes (« poulpo »). L’étude biologique montre qu’il existe deux saisons de reproduction par an (mai-juillet et septembre-novembre) qui engendrent deux recrutements, l’un en mai et l’autre à partir de septembre. La croissance pondérale a été obtenue sur les des distributions de fréquences de poids. Une transformation des poids en longueurs a permis de déterminer les paramètres de l’équation de VON BERTALANFFY pour chaque cohorte. Après détermination des coefficients de mortalité naturelle et par pêche, les résultats obtenus par le modèle global et le modèle de RICKER sont discutés. La mise en œuvre d’un modèle global montre que le niveau actuel des captures est proche de la prise maximale équilibrée et que l’effort de pêche ne doit pas être augmenté. //
    Description: This study deals with the exploitation, the biology and the dynamics of common octopus (Octopus vulgaris, Cuvier) in Mauritanian waters. In the Cap Blanc region, the spatio-temporal distribution appears to be independent of the length and sex but variable through the year. The analysis of the present state of the fishery reveals an increase of the fishing effort and continuous decrease of the catch per unit effort. An increase of the small individuals, with weight under 200 g (“poulpo”), can be seen from catches composition. The biological study shows that there are two spawning seasons (May-July and September-November) which allow two recruitments in May and September. The growth parameters of the VON BERTALANFFY equation are calculated from the transformed weights in lengths. The calculated mortality coefficients are used in a RICKER’s model of yield per recruit and the results are discussed. The use of a generalized production model shows that the present catch is very close to the maximum sustainable yield. Therefore, one would aware to prevent an increase of the fishing effort.
    Description: IMROP
    Description: Unpublished
    Keywords: Reproduction ; Common octopus ; Population dynamics ; Environnement ; Dynamique de population ; Exploitation ; Environment ; Poulpe ; ZEE Mauritanienne ; Distribution ; Mauritanian Exclusive Economic Zone ; Biologie ; Croissance ; Growth ; Octopus vulgaris ; Biology ; Octopus fisheries
    Repository Name: AquaDocs
    Type: Theses and Dissertations , Master thesis
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  • 8
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    American Association for the Advancement of Science (AAAS)
    Publication Date: 2005-12-24
    Description: 〈br /〉〈span class="detail_caption"〉Notes: 〈/span〉New York, N.Y. -- Science. 2005 Dec 23;310(5756):1880-5.〈br /〉〈span class="detail_caption"〉Record origin:〈/span〉 〈a href="http://www.ncbi.nlm.nih.gov/pubmed/16373539" target="_blank"〉PubMed〈/a〉
    Keywords: Astronomical Phenomena ; Astronomy ; Biology ; Brain/growth & development ; Brain Diseases/genetics ; Climate ; Earth (Planet) ; Evolution, Planetary ; Humans ; Nuclear Reactors ; Plant Development ; Plants/genetics ; Potassium Channels ; *Research ; Systems Theory
    Print ISSN: 0036-8075
    Electronic ISSN: 1095-9203
    Topics: Biology , Chemistry and Pharmacology , Computer Science , Medicine , Natural Sciences in General , Physics
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  • 9
    Publication Date: 2005-11-02
    Description: The ground testing of a Rocket Based Combined Cycle engine implementing the Simultaneous Mixing and Combustion scheme was performed at the direct-connect facility of Purdue University's High Pressure Laboratory. The fuel-rich exhaust of a JP-8/H2O2 thruster was mixed with compressed, metered air in a constant area, axisymmetric duct. The thruster was similar in design and function to that which will be used in the flight test series of Dryden's Ducted-Rocket Experiment. The determination of duct ignition limits was made based on the variation of secondary air flow rates and primary thruster equivalence ratios. Thrust augmentation and improvements in specific impulse were studied along with the pressure and temperature profiles of the duct to study mixing lengths and thermal choking. The occurrence of ignition was favored by lower rocket equivalence ratios. However, among ignition cases, better thrust and specific impulse performance were seen with higher equivalence ratios owing to the increased fuel available for combustion. Thrust and specific impulse improvements by factors of 1.2 to 1.7 were seen. The static pressure and temperature profiles allowed regions of mixing and heat addition to be identified. The mixing lengths were found to be shorter at lower rocket equivalence ratios. Total pressure measurements allowed plume-based calculation of thrust, which agreed with load-cell measured values to within 6.5-8.0%. The corresponding Mach Number profile indicated the flow was not thermally choked for the highest duct static pressure case.
    Keywords: Spacecraft Propulsion and Power
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  • 10
    Publication Date: 2018-06-12
    Description: Contents include the following: Oxygen Compatible Materials. Manufacturing Technology Demonstrations. Turbopump Inducer Waterflow Test. Turbine Damping "Whirligig" Test. Single Element Preburner and Main Injector Test. 40K Multi-Element Preburner and MI. Full-Scale Battleship Preburner. Prototype Preburner Test Article. Full-Scale Prototype TCA. Turbopump Hot-Fire Test Article. Prototype Engine. Validated Analytical Models.
    Keywords: Spacecraft Propulsion and Power
    Type: Fifth International Symposium on Liquid Space Propulsion; NASA/CP-2005-213607
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  • 11
    Publication Date: 2018-06-12
    Description: Development of Liquid Rocket Engines is expensive. Extensive testing at large scales usually required. In order to verify engine lifetime, large number of tests required. Limited Resources available for development. Sub-scale cold-flow and hot-fire testing is extremely cost effective. Could be a necessary (but not sufficient) condition for long engine lifetime. Reduces overall costs and risk of large scale testing. Goal: Determine knowledge that can be gained from sub-scale cold-flow and hot-fire evaluations of LRE injectors. Determine relationships between cold-flow and hot-fire data.
    Keywords: Spacecraft Propulsion and Power
    Type: Fifth International Symposium on Liquid Space Propulsion; NASA/CP-2005-213607
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  • 12
    Publication Date: 2018-06-12
    Description: Major Causes: Limited Initial Materials Properties. Limited Structural Models - especially fatigue. Limited Thermal Models. Limited Aerodynamic Models. Human Errors. Limited Component Test. High Pressure. Complicated Control.
    Keywords: Spacecraft Propulsion and Power
    Type: Fifth International Symposium on Liquid Space Propulsion; NASA/CP-2005-213607
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  • 13
    Publication Date: 2018-06-12
    Description: The subject of mathematical modeling of the transient operation of liquid rocket engines is presented in overview form from the perspective of engineers working at the NASA Marshall Space Flight Center. The necessity of creating and utilizing accurate mathematical models as part of liquid rocket engine development process has become well established and is likely to increase in importance in the future. The issues of design considerations for transient operation, development testing, and failure scenario simulation are discussed. An overview of the derivation of the basic governing equations is presented along with a discussion of computational and numerical issues associated with the implementation of these equations in computer codes. Also, work in the field of generating usable fluid property tables is presented along with an overview of efforts to be undertaken in the future to improve the tools use for the mathematical modeling process.
    Keywords: Spacecraft Propulsion and Power
    Type: Fifth International Symposium on Liquid Space Propulsion; NASA/CP-2005-213607
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  • 14
    Publication Date: 2018-06-12
    Description: Contents include the following: SLI initiated under NASA Research Announcement (NRA) 8-30. Strategic Objectives. Make spaceflight safer (1 in 10000 mission LOV). Make spaceflight cheaper ($1000/lb payload). Two prototype LOX/LH2 engine systems funded under Cycle-1 of NRA8-30. COBRA (Pratt & Whitney / Aerojet). RS-83 (Rocketdyne).
    Keywords: Spacecraft Propulsion and Power
    Type: Fifth International Symposium on Liquid Space Propulsion; NASA/CP-2005-213607
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  • 15
    Publication Date: 2018-06-12
    Description: A) MSFC funded an internal study on Altitude Compensating Nozzles: 1) Develop an ACN design and performance prediction tool. 2) Design, build and test cold flow ACN nozzles. 3) An annular aerospike nozzle was designed and tested. 4) Incorporated differential throttling to assess Thrust Vector Control. B) Objective of the test hardware: 1) Provide design tool verification. 2) Provide benchmark data for CFD calculations. 3) Experimentally measure side force, or TVC, for a differentially throttled annular aerospike.
    Keywords: Spacecraft Propulsion and Power
    Type: Fifth International Symposium on Liquid Space Propulsion; NASA/CP-2005-213607
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  • 16
    Publication Date: 2018-06-12
    Description: It is well known that under some operating conditions, rocket engines (using solid or liquid fuels) exhibit unstable modes of operation that can lead to engine malfunction and shutdown. The sources of these instabilities are diverse and are dependent on fuel, chamber geometry and various upstream sources such as pumps, valves and injection mechanism. It is believed that combustion-acoustic instabilities occur when the acoustic energy increase due to the unsteady heat release of the flame is greater than the losses of acoustic energy from the system [1, 2]. Giammar and Putnam [3] performed a comprehensive study of noise generated by gasfired industrial burners and made several key observations; flow noise was sometimes more intense than combustion roar, which tended to have a characteristic frequency spectrum. Turbulence was amplified by the flame. The noise power varied directly with combustion intensity and also with the product of pressure drop and heat release rate. Karchmer [4] correlated the noise emitted from a turbofan jet engine with that in the combustion chamber. This is important, since it quantified how much of the noise from an engine originates in the combustor. A physical interpretation of the interchange of energy between sound waves and unsteady heat release rates was given by Rayleigh [5] for inviscid, linear perturbations. Bloxidge et al [6] extended Rayleigh s criterion to describe the interaction of unsteady combustion with one-dimensional acoustic waves in a duct. Solutions to the mass, momentum and energy conservation equations in the pre- and post-flame zones were matched by making several assumptions about the combustion process. They concluded that changes in boundary conditions affect the energy balance of acoustic waves in the combustor. Abouseif et al [7] also solved the one-dimensional flow equations, but they used a onestep reaction to evaluate the unsteady heat release rate by relating it to temperature and velocity perturbations. Their analysis showed that oscillations arise from coupling between entropy waves produced at the flame and pressure waves originating from the nozzle. Yang and Culick [8] assumed a thin flame sheet, which is distorted by velocity and pressure oscillations. Conservation equations were expressed in integral form and solutions for the acoustic wave equations and complex frequencies were obtained. The imaginary part of the frequency indicated stability regions of the flame. Activation energy asymptotics together with a one-step reaction were used by McIntosh [9] to study the effects of acoustic forcing and feedback on unsteady, one-dimensional flames. He found that the flame stability was altered by the upstream acoustic feedback. Shyy et al [10] used a high-accuracy TVD scheme to simulate unsteady, one-dimensional longitudinal, combustion instabilities. However, numerical diffusion was not completely eliminated. Recently, Prasad [11] investigated numerically the interactions of pressure perturbations with premixed flames. He used complex chemistry to study responses of pressure perturbations in one-dimensional combustors. His results indicated that reflected and transmitted waves differed significantly from incident waves.
    Keywords: Spacecraft Propulsion and Power
    Type: The 2004 NASA Faculty Fellowship Program Research Reports; XV-1 - XV-24; NASA/CR-2005-213847
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  • 17
    Publication Date: 2018-06-12
    Description: Shuttle Redesigned Solid Rocket Motor (RSRM) nozzle interiors fabricated from carbon phenolic composite exhibit "ply lift" when hot fired. The composite surface is smooth when fabricated, but the individual plies separate and lift away from the surface when exposed to high temperature and high-pressure exhaust gas. It shows a cross section of a post-fired composite in which ply lift is evident as dark fissures. Surface charring is also visible as a darker band about 0.2 inches thick. Charring is normal, but ply lift is not desirable since the fissures could possibly initiate an abnormal exhaust path from the RSRM. The underlying mechanisms of ply lift are under investigation as part of the Shuttle Return-To-Flight Program.
    Keywords: Spacecraft Propulsion and Power
    Type: The 2004 NASA Faculty Fellowship Program Research Reports; XII-1 - XII-5; NASA/CR-2005-213847
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  • 18
    Publication Date: 2018-06-12
    Description: When the President offered his new vision for space exploration in January of 2004, he said, "Our third goal is to return to the moon by 2020, as the launching point for missions beyond," and, "With the experience and knowledge gained on the moon, we will then be ready to take the next steps of space exploration: human missions to Mars and to worlds beyond." A human mission to Mars implies the need to move large payloads as rapidly as possible, in an efficient and cost-effective manner. Furthermore, with the scientific advancements possible with Project Prometheus and its Jupiter Icy Moons Orbiter (JIMO), (these use electric propulsion), there is a renewed interest in deep space exploration propulsion systems. According to many mission analyses, nuclear thermal propulsion (NTP), with its relatively high thrust and high specific impulse, is a serious candidate for such missions. Nuclear rockets utilize fission energy to heat a reactor core to very high temperatures. Hydrogen gas flowing through the core then becomes superheated and exits the engine at very high exhaust velocities. The combination of temperature and low molecular weight results in an engine with specific impulses above 900 seconds. This is almost twice the performance of the LOX/LH2 space shuttle engines, and the impact of this performance would be to reduce the trip time of a manned Mars mission from the 2.5 years, possible with chemical engines, to about 12-14 months.
    Keywords: Spacecraft Propulsion and Power
    Type: The 2004 NASA Faculty Fellowship Program Research Reports; XXIV-1 - XXIV-7; NASA/CR-2005-213847
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  • 19
    Publication Date: 2018-06-11
    Description: In this paper, we will describe the electronic propulsion technologies of interest and our role in developing and interjecting these technologies into JPL missions.
    Keywords: Spacecraft Propulsion and Power
    Type: 2005 AIAA Joint Propulsion Conference; Tucson, AZ; United States
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  • 20
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    Publication Date: 2018-06-11
    Description: This Final Report serves as an executive summary of the Prometheus Project's activities and deliverables from November 2002 through September 2005. It focuses on the challenges from a technical and management perspective, what was different and innovative about this project, and identifies the major options, decisions, and accomplishments of the Project team as a whole. However, the details of the activities performed by DOE NR and its contractors will be documented separately in accordance with closeout requirements of the DOE NR and consistent with agreements between NASA and NR.
    Keywords: Spacecraft Propulsion and Power
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  • 21
    Publication Date: 2018-06-11
    Description: A bismuth feed system was developed for the VHITAL Program to deliver 8-12 mg/s of bismuth vapor at a few Torr to the VHITAL-160. A carbon vaporizer developed to control vapor flow rates to the thruster.
    Keywords: Spacecraft Propulsion and Power
    Type: International Electric Propulsion Conference 2005
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  • 22
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    Publication Date: 2018-06-11
    Description: This study has advanced state-of-the-art dishcarge modeling and revealed important aspects of discharge plasma processes.
    Keywords: Spacecraft Propulsion and Power
    Type: Joint Propulsion Conference
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  • 23
    Publication Date: 2018-06-11
    Description: The power, Isp and thrust of ion thrusters are constrained by ther fixed grid gap in the ion accellerator, which limits performance and life to a limited range in Isp and thrust.
    Keywords: Spacecraft Propulsion and Power
    Type: 2005 AIAA Joint Propulsion Conference; Tucson, AZ; United States
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  • 24
    facet.materialart.
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    In:  CASI
    Publication Date: 2018-06-05
    Description: The majority of new satellites generate electrical power using photovoltaic solar arrays and store energy in batteries for use during eclipse periods. Careful regulation of battery charging during insolation can greatly increase the expected lifetime of the satellite. The battery charge regulator is usually custom designed for each satellite and its specific mission. Economic competition in the small satellite market requires battery charge regulators that are lightweight, efficient, inexpensive, and modular enough to be used in a wide variety of satellites. A new battery charge regulator topology has been developed at the NASA Lewis Research Center to address these needs. The new regulator topology uses industry-standard dc-dc converters and a unique interconnection to provide size, weight, efficiency, fault tolerance, and modularity benefits over existing systems. A transformer-isolated buck converter is connected such that the high input line is connected in series with the output. This "bypass connection" biases the converter's output onto the solar array voltage. Because of this biasing, the converter only processes the fraction of power necessary to charge the battery above the solar array voltage. Likewise, the same converter hookup can be used to regulate the battery output to the spacecraft power bus with similar fractional power processing. The advantages of this scheme are: 1) Because only a fraction of the power is processed through the dc-dc converter, the single- stage conversion efficiency is 94 to 98 percent; 2) Costly, high-efficiency dc-dc converters are not necessary for high end-to-end system efficiency; 3) The system is highly fault tolerant because the bypass connection will still deliver power if the dc-dc converter fails; and 4) The converters can easily be connected in parallel, allowing higher power systems to be built from a common building block. This new technology will be spaceflight tested in the Photovoltaic Regulator Kit Experiment (PRKE) on TRW's Small Spacecraft Technology Initiative (SSTI) satellite scheduled for launch in 1996. This experiment uses commercial dc-dc converters (28 to 15 Vdc) and additional control circuitry to regulate current to a battery load. The 60-W, 87- percent efficiency converters can control 180 W of power at an efficiency of 94 percent in the new configuration. The power density of the Photovoltaic Regulator Kit Experiment is about 200 W/kg.
    Keywords: Spacecraft Propulsion and Power
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  • 25
    Publication Date: 2018-06-11
    Description: Planar laser-induced fluorescence visualisation is used to investigate nonuniformities in the flow of a hypersonic conical nozzle. Possible causes for the nonuniformity are outlined and investigated, and the problem is shown to be due to a small step at the nozzle throat. Entrainment of cold boundary layer gas is postulated as the cause of the signal nonuniformity.
    Keywords: Spacecraft Propulsion and Power
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  • 26
    Publication Date: 2018-06-05
    Description: The NASA Glenn Research Center initiated baseline testing of ultracapacitors to obtain empirical data in determining the feasibility of using ultracapacitors for the Next Generation Launch Transportation (NGLT) Project. There are large transient loads associated with NGLT that require a very large primary energy source or an energy storage system. The primary power source used for this test was a proton-exchange-membrane (PEM) fuel cell. The energy storage system can consist of batteries, flywheels, or ultracapacitors. Ultracapacitors were used for these tests. NASA Glenn has a wealth of experience in ultracapacitor technology through the Hybrid Power Management (HPM) Program, which the Avionics, Power and Communications Branch of Glenn s Engineering Development Division initiated for the Technology Transfer and Partnership Office. HPM is the innovative integration of diverse, state-ofthe- art power devices in optimal configurations for space and terrestrial applications. The appropriate application and control of the various advanced power devices (such as ultracapacitors and fuel cells) significantly improves overall system performance and efficiency. HPM has extremely wide potential. Applications include power generation, transportation systems, biotechnology systems, and space power systems. HPM has the potential to significantly alleviate global energy concerns, improve the environment, and stimulate the economy.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2004; NASA/TM-2005-213419
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  • 27
    Publication Date: 2018-06-05
    Description: Low-pressure turbine (LPT) airfoils are subject to increasingly stronger pressure gradients as designers impose higher loading in an effort to improve efficiency and lower cost by reducing the number of airfoils in an engine. When the adverse pressure gradient on the suction side of these airfoils becomes strong enough, the boundary layer will separate. Separation bubbles, particularly those that fail to reattach, can result in a significant loss of lift and a subsequent degradation of engine efficiency. The problem is particularly relevant in aircraft engines. Airfoils optimized to produce maximum power under takeoff conditions may still experience boundary layer separation at cruise conditions because of the thinner air and lower Reynolds numbers at altitude. Component efficiency can drop significantly between takeoff and cruise conditions. The decrease is about 2 percent in large commercial transport engines, and it could be as large as 7 percent in smaller engines operating at higher altitudes. Therefore, it is very beneficial to eliminate, or at least reduce, the separation bubble.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2004; NASA/TM-2005-213419
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  • 28
    Publication Date: 2018-06-05
    Description: A free-piston Stirling power convertor is being considered as an advanced power-conversion technology for future NASA deep-space missions requiring long-life radioisotope power systems. The NASA Glenn Research Center has identified key areas where advanced technologies can enhance the capability of Stirling energy-conversion systems. One of these is power electronic controls. Current power-conversion technology for Glenn-tested Stirling systems consists of an engine-driven linear alternator generating an alternating-current voltage controlled by a tuning-capacitor-based alternating-current peak voltage load controller. The tuning capacitor keeps the internal alternator electromotive force (EMF) in phase with its respective current (i.e., passive power factor correction). The alternator EMF is related to the piston velocity, which must be kept in phase with the alternator current in order to achieve stable operation. This tuning capacitor, which adds volume and mass to the overall Stirling convertor, can be eliminated if the controller can actively drive the magnitude and phase of the alternator current.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2004; NASA/TM-2005-213419
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  • 29
    Publication Date: 2018-06-05
    Description: NASA's Next Generation Launch Technology (NGLT) Program has successfully demonstrated cooled ceramic matrix composite (CMC) technology in a scramjet engine test. This demonstration represented the world s largest cooled nonmetallic matrix composite panel fabricated for a scramjet engine and the first cooled nonmetallic composite to be tested in a scramjet facility. Lightweight, high-temperature, actively cooled structures have been identified as a key technology for enabling reliable and low-cost space access. Tradeoff studies have shown this to be the case for a variety of launch platforms, including rockets and hypersonic cruise vehicles. Actively cooled carbon and CMC structures may meet high-performance goals at significantly lower weight, while improving safety by operating with a higher margin between the design temperature and material upper-use temperature. Studies have shown that using actively cooled CMCs can reduce the weight of the cooled flow-path component from 4.5 to 1.6 lb/sq ft and the weight of the propulsion system s cooled surface area by more than 50 percent. This weight savings enables advanced concepts, increased payload, and increased range. The ability of the cooled CMC flow-path components to operate over 1000 F hotter than the state-of-the-art metallic concept adds system design flexibility to space-access vehicle concepts. Other potential system-level benefits include smaller fuel pumps, lower part count, lower cost, and increased operating margin.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2004; NASA/TM-2005-213419
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  • 30
    Publication Date: 2018-06-05
    Description: The Forward Technology Solar Cell Experiment (FTSCE) is a space solar cell experiment built as part of the Fifth Materials on the International Space Station Experiment (MISSE-5): Data Acquisition and Control Hardware and Software. It represents a collaborative effort between the NASA Glenn Research Center, the Naval Research Laboratory, and the U.S. Naval Academy. The purpose of this experiment is to place current and future solar cell technologies on orbit where they will be characterized and validated. This is in response to recent on-orbit and ground test results that raised concerns about the in-space survivability of new solar cell technologies and about current ground test methodology. The various components of the FTSCE are assembled into a passive experiment container--a 2- by 2- by 4-in. folding metal container that will be attached by an astronaut to the outer structure of the International Space Station. Data collected by the FTSCE will be relayed to the ground through a transmitter assembled by the U.S. Naval Academy. Data-acquisition electronics and software were designed to be tolerant of the thermal and radiation effects expected on orbit. The experiment has been verified and readied for flight on STS-114.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2004; NASA/TM-2005-213419
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  • 31
    Publication Date: 2018-06-05
    Description: Electric power system performance predictions are critical to spacecraft, such as the International Space Station (ISS), to ensure that sufficient power is available to support all the spacecraft s power needs. In the case of the ISS power system, analyses to date have been deterministic, meaning that each analysis produces a single-valued result for power capability because of the complexity and large size of the model. As a result, the deterministic ISS analyses did not account for the sensitivity of the power capability to uncertainties in model input variables. Over the last 10 years, the NASA Glenn Research Center has developed advanced, computationally fast, probabilistic analysis techniques and successfully applied them to large (thousands of nodes) complex structural analysis models. These same techniques were recently applied to large, complex ISS power system models. This new application enables probabilistic power analyses that account for input uncertainties and produce results that include variations caused by these uncertainties. Specifically, N&R Engineering, under contract to NASA, integrated these advanced probabilistic techniques with Glenn s internationally recognized ISS power system model, System Power Analysis for Capability Evaluation (SPACE).
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2004; NASA/TM-2005-213419
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  • 32
    Publication Date: 2018-06-05
    Description: The Space Shuttle Main Engine (SSME), developed 30 years ago, remains a strong candidate for use in the new Exploration Initiative as part of a shuttle-derived heavy-lift expendable booster. This is because the Boeing-Rocket- dyne man-rated SSME remains the most highly efficient liquid rocket engine ever developed. There are only enough parts for 12-15 existing SSMEs, however, so one NASA option is to reinitiate SSME production to use it as a throw-away, as opposed to a reusable, powerplant for NASA s new heavy-lift booster.
    Keywords: Spacecraft Propulsion and Power
    Type: Aviation Week and Space Technology (ISSN 0005-2175); Volume 163; No. 2; 59
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  • 33
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    Publication Date: 2018-06-11
    Description: A briefing on the propulsion system modification of the STS-114 Discovery is presented. June Malone, NASA Public Affairs, introduces the panel who consists of: Sandy Coleman, External Tank Project Manager, Neil Otte, External Tank Chief Engineer, and Tom Williams, Solid Rocket Booster, Deputy Project Manager. Neil Otte presents charts on new requirements for foam debris reduction on the external tank. He also presents charts describing the Forward Bipod Redesign, LO2 Feedline Bellows Location, LH2 Intertank Flange Location, and In-Flight Imagery. Tom Williams presents charts describing Solid Rocket Booster Activities and Return to Flight efforts.
    Keywords: Spacecraft Propulsion and Power
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  • 34
    facet.materialart.
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    In:  CASI
    Publication Date: 2018-06-06
    Description: Mars has greatly intrigued scientists and the general public for many years because, of all the planets, its environment is most like Earth's. Many scientists believe that Mars once had running water, although surface water is gone today. The planet is very cold with a very thin atmosphere consisting mainly of CO2. Mariner 4, 6, and 7 explored the planet in flybys in the 1960s and by the orbiting Mariner 9 in 1971. NASA then mounted the ambitious Viking mission, which launched two orbiters and two landers to the planet in 1975. The landers found ambiguous evidence of life. Mars Pathfinder landed on the planet on July 4, 1997, delivering a mobile robot rover that demonstrated exploration of the local surface environment. Mars Global Surveyor is creating a highest-resolution map of the planet's surface. These prior and current missions to Mars have paved the way for a complex Mars Sample Return mission planned for 2003 and 2005. Returning surface samples from Mars will necessitate retrieval of material from Mars orbit. Sample mass and orbit are restricted to the launch capability of the Mars Ascent Vehicle. A small sample canister having a mass less than 4 kg and diameter of less than 16 cm will spend from three to seven years in a 600 km orbit waiting for retrieval by a second spacecraft consisting of an orbiter equipped with a sample canister retrieval system, and a Earth Entry Vehicle. To allow rapid detection of the on-orbit canister, rendezvous, and collection of the samples, the canister will have a tracking beacon powered by a surface mounted solar array. The canister must communicate using RF transmission with the recovery vehicle that will be coming in 2006 or 2009 to retrieve the canister. This paper considers the aspect and conclusion that went into the design of the power system that achieves the maximum power with the minimum risk. The power output for the spherical orbiting canister was modeled and plotted in various views of the orbit by the Satellite Orbit Analysis Program (SOAP).
    Keywords: Spacecraft Propulsion and Power
    Type: 16th Space Photovoltaic Research and Technology Conference; 238-241; NASA/CP-2001-210747/REV1
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  • 35
    Publication Date: 2019-07-27
    Description: This paper provides a summary of testing of Space Shuttle Main Engine (SSME) flowmeter bearings and cage material. These tests were con&cM over a several month period in 2004 at the Marshall Space Flight Center. The test program's primary objective was to compare the performance of bearings using the existing cage material and bearings using a proposed replacement cage material. In order to meet the test objectives for this program, a flowmeter test rig was designed and fabricated to measure both breakaway and running torque for a flowmeter assembly. Other test parameters,,such as motor current and shaft speed, were also recorded and provide a means of comparing bearing performance. The flowmeter and bearings were tested in liquid hydrogen to simulate the flowmeter's operating environment as closely as possible. Based on the results from this testing, the bearings with the existing cage material are equivalent to the bearings with the proposed replacement cage material. No major differences exist between the old and new cage materials. Therefore, the new cage material is a suitable replacement for the existing cage material.
    Keywords: Spacecraft Propulsion and Power
    Type: WTC2005-63299 , World Tribology Conference III; 12-16 Sept. 2005; Washington DC.; United States
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  • 36
    Publication Date: 2019-07-27
    Description: The complex interactions between internal motor generated pressure oscillations and motor structural vibration modes associated with the static test configuration of a Reusable Solid Rocket Motor have potential to generate significant dynamic thrust loads in the 5-segment configuration (Engineering Test Motor 3). Finite element model load predictions for worst-case conditions were generated based on extrapolation of a previously correlated 4-segment motor model. A modal survey was performed on the largest rocket motor to date, Engineering Test Motor #3 (ETM-3), to provide data for finite element model correlation and validation of model generated design loads. The modal survey preparation included pretest analyses to determine an efficient analysis set selection using the Effective Independence Method and test simulations to assure critical test stand component loads did not exceed design limits. Historical Reusable Solid Rocket Motor modal testing, ETM-3 test analysis model development and pre-test loads analyses, as well as test execution, and a comparison of results to pre-test predictions are discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: IMAC XX111; 31 Jan. 3 Feb. 2005; Orlando, FL; United States
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  • 37
    Publication Date: 2019-07-27
    Description: This paper describes potential heat rejection design concepts for closed Brayton cycle (CBC) power conversion systems. Brayton conversion systems are currently under study by NASA for Nuclear Electric Propulsion (NEP) applications. The Heat Rejection Subsystem (HRS) must dissipate waste heat generated by the power conversion system due to inefficiencies in the thermal-to-electric conversion process. Space Brayton conversion system designs tend to optimize at efficiencies of about 20 to 25 percent with radiator temperatures in the 400 to 600 K range. A notional HRS was developed for a 100 kWe-class Brayton power system that uses a pumped sodium-potassium (NaK) heat transport loop coupled to a water heat pipe radiator. The radiator panels employ a sandwich construction consisting of regularly-spaced circular heat pipes contained within two composite facesheets. Heat transfer from the NaK fluid to the heat pipes is accomplished by inserting the evaporator sections into the NaK duct channel. The paper evaluates various design parameters including heat pipe diameter, heat pipe spacing, and facesheet thickness. Parameters were varied to compare design options on the basis of NaK pump pressure rise and required power, heat pipe unit power and radial flux, radiator panel areal mass, and overall HRS mass.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2005-213337 , E-14807 , AIAA Paper 2004-5654 , Second International Energy Conversion Engineering Conference; 16-19 aAug. 2004; Providence, RI; United States
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  • 38
    Publication Date: 2019-07-18
    Description: A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines. These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine. The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster. Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows.
    Keywords: Spacecraft Propulsion and Power
    Type: 53rd JPM/2nd LPS/SP Joint Meeting; Dec 05, 2005 - Dec 08, 2005; Monterey, CA; United States
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  • 39
    Publication Date: 2019-07-18
    Description: NASA's mission to "reach the Moon and Mars" will be obtained only if research begins now to develop materials with expanded capabilities to reduce mass, cost and risk to the program. Current materials cannot function satisfactorily in the deep space environments and do not meet the requirements of long term space propulsion concepts for manned missions. Directed research is needed to better understand materials behavior for optimizing their processing. This research, generating a deeper understanding of material behavior, can lead to enhanced implementation of materials for future exploration vehicles. materials providing new approaches for manufacture and new options for In response to this need for more robust materials, NASA's Exploration Systems Mission Directorate (ESMD) has established a strategic research initiative dedicated to materials development supporting NASA's space propulsion needs. The Advanced Materials for Exploration (AME) element directs basic and applied research to understand material behavior and develop improved materials allowing propulsion systems to operate beyond their current limitations. This paper will discuss the approach used to direct the path of strategic research for advanced materials to ensure that the research is indeed supportive of NASA's future missions to the moon, Mars, and beyond.
    Keywords: Spacecraft Propulsion and Power
    Type: 43rd AIAA Aerospace Sciences Meeting and Exhibit; Jan 10, 2005 - Jan 13, 2005; Reno, NV; United States
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  • 40
    Publication Date: 2019-07-18
    Description: The Space Launch Initiative (SLI) procurement mechanism NRA8-30 initiated the Auxiliary Propulsion System/Main Propulsion System (APS/MPS) Project in 2001 to address technology gaps and development risks for non-toxic and cryogenic propellants for auxiliary propulsion applications. These applications include reaction control and orbital maneuvering engines, and storage, pressure control, and transfer technologies associated with on-orbit maintenance of cryogens. The project has successfully evolved over several years in response to changing requirements for re-usable launch vehicle technologies, general launch technology improvements, and, most recently, exploration technologies. Lessons learned based on actual hardware performance have also played a part in the project evolution to focus now on those technologies deemed specifically relevant to the Exploration Initiative. Formal relevance reviews held in the spring of 2004 resulted in authority for continuation of the Auxiliary Propulsion Project through Fiscal Year 2005 (FY05), and provided for a direct reporting path to the Exploration Systems Mission Directorate. The tasks determined to be relevant under the project were: continuation of the development, fabrication, and delivery of three 870 lbf thrust prototype LOX/ethanol reaction control engines; the fabrication, assembly, engine integration and testing of the Auxiliary Propulsion Test Bed at White Sands Test Facility; and the completion of FY04 cryogenic fluid management component and subsystem development tasks (mass gauging, pressure control, and liquid acquisition elements). This paper presents an overview of those tasks, their scope, expectations, and results to-date as carried forward into the Exploration Initiative.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA-2005-TBD , AIAA 1st Space Exploration Conference; Jan 30, 2005 - Feb 02, 2005; Orlando, FL; United States
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  • 41
    Publication Date: 2019-07-18
    Description: A significant problem in the use of electric thrusters in spacecraft is the formation of low-energy ions in the thruster plume. Low-energy ions are formed in the plume via random collisions between high-velocity ions ejected from the thruster and slow-moving neutral atoms of propellant effusing from the engine. The sputtering of spacecraft materials due to interactions with low-energy ions may result in erosion or contamination of the spacecraft. The trajectory of these ions is determined primarily by the plasma potential of the plume. Thus, accurate characterization of the plasma potential is essential to predicting low-energy ion contamination. Emissive probes were utilized to determine the plasma potential. When the ion and electron currents to the probe are balanced, the potential of such probes float to the plasma potential. Two emissive probes were fabricated; one utilizing a DC power supply, another utilizing a rectified AC power source. Labview programs were written to coordinate and automate probe motion in the thruster plume. Employing handshaking interaction, these motion programs were synchronized to various data acquisition programs to ensure precision and accuracy of the measurements. Comparing these experimental values to values from theoretical models will allow for a more accurate prediction of low-energy ion interaction.
    Keywords: Spacecraft Propulsion and Power
    Type: Summer Student Research Presentations; 31; JPL-Publ-05-07
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  • 42
    Publication Date: 2019-07-18
    Description: The required type and amount of numerical dissipation/filter to accurately resolve all relevant multiscales of complex MHD unsteady high-speed shock/shear/turbulence/combustion problems are not only physical problem dependent, but also vary from one flow region to another. In addition, proper and efficient control of the divergence of the magnetic field (Div(B)) numerical error for high order shock-capturing methods poses extra requirements for the considered type of CPU intensive computations. The goal is to extend our adaptive numerical dissipation control in high order filter schemes and our new divergence-free methods for ideal MHD to non-ideal MHD that include viscosity and resistivity. The key idea consists of automatic detection of different flow features as distinct sensors to signal the appropriate type and amount of numerical dissipation/filter where needed and leave the rest of the region free from numerical dissipation contamination. These scheme-independent detectors are capable of distinguishing shocks/shears, flame sheets, turbulent fluctuations and spurious high-frequency oscillations. The detection algorithm is based on an artificial compression method (ACM) (for shocks/shears), and redundant multiresolution wavelets (WAV) (for the above types of flow feature). These filters also provide a natural and efficient way for the minimization of Div(B) numerical error.
    Keywords: Spacecraft Propulsion and Power
    Type: Advanced Space Transportation Systems; Apr 25, 2005 - Apr 29, 2005; Rome; Italy
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  • 43
    Publication Date: 2019-07-18
    Description: The concept of the electrodynamic propulsion has a number of attractive features and has been widely discussed for different applications. A number of system designs have been proposed and compared during the last ten years. In spite of this, the choice of the proper design, for a specific mission, is far from evident. Such characteristics of tether performance as system acceleration, efficiency, etc. should be calculated and compared. The code that calculates the current for bare and partly insulated tethers with circular (wire) and rectangle (tape) cross-sections is presented. It takes into account the corrections to the OML current due to the tether cross-section geometry and the magnetic field produced by the tether current. There are two options in this code: for current calculation with the prescribed energy supply and with the prescribed end-point potential. This permits us to calculate the parameters characterizing tether performance. Results for the current calculated for tethers with different designs for the currently proposed Momentum exchange Electrodynamic Reboost (MXER) Tether System are presented.
    Keywords: Spacecraft Propulsion and Power
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  • 44
    Publication Date: 2019-07-18
    Description: Propulsion for aerospace applications is limited by two basic parameters: specific energy (MJ/kg) and specific power (KW/kg). Specific energy can perhaps be improved by increasing the energy content of propellants, increasing energy storage of other on-board devices, and by the use of intense off-board energy sources such as beamed energy. Several beamed energy concepts for space access have been investigated using Lasers and Microwave beams. Several preliminary concepts have been examined for high altitude platforms for commercial or military applications. Some of these results are described. Additionally, two concepts are briefly described for potentially improving on-board specific energy: Metallic Hydrogen and Magnetic Energy Storage.
    Keywords: Spacecraft Propulsion and Power
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  • 45
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2019-07-18
    Description: The gasdynamic mirror has been proposed as a concept which could form the basis of a highly efficient fusion rocket engine. Gasdynamic mirrors differ from most other mirror type plasma confinement schemes in that they have much larger aspect ratios and operate at somewhat higher plasma densities. There are several types of instabilities which are known to plague mirror type confinement schemes. These instabilities fall into two general classes. One class of instability is the Magnetohydrodynamic or MHD instability which induces gross distortions in the plasma geometry. The other class of instability is the "loss cone" microinstability which leads to general plasma turbulence. The "loss cone" microinstability is caused by velocity space asymmetries resulting from the loss of plasma having constituent particle velocities within the angle of the magnetic mirror "loss cone." These instabilities generally manifest themselves in high temperature, moderately dense plasmas. The present study indicates that a GDM configured as a rocket engine might operate in a plasma regime where microinstabilities could potentially be significant.
    Keywords: Spacecraft Propulsion and Power
    Type: Joint Propulsion Conference; Jul 11, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 46
    Publication Date: 2019-07-18
    Description: An advanced Cu-8(at.%)Cr-4%Nb alloy developed at NASA's Glenn Research Center, and designated as GRCop-84, is currently being considered for use as liners in combustor chambers and nozzle ramps in NASA s future generations of reusable launch vehicles (RLVs). However, past experience has shown that unprotected copper alloys undergo an environmental attack called "blanching" in rocket engines using liquid hydrogen as fuel and liquid oxygen as the oxidizer. Potential for sulfidation attack of the liners in hydrocarbon-fueled engines is also of concern. As a result, protective overlay coatings alloys are being developed for GRCop-84. The oxidation behavior of several new coating alloys has been evaluated. GRCop-84 specimens were coated with several copper and nickel-based coatings, where the coatings were deposited by either vacuum plasma spraying or cold spraying techniques. Coated and uncoated specimens were thermally cycled in a furnace at different temperatures in order to evaluate the performance of the coatings. Additional studies were conducted in a high pressure burner rig using a hydrocarbon fuel and subjected to a high heat flux hydrogen-oxygen combustion flame in NASA s Quick Access Rocket Exhaust (QARE) rig. The performance of these coatings are discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: 2005 TMS Annual Meeting; Feb 13, 2005 - Feb 17, 2005; San Francisco, CA; United States
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  • 47
    Publication Date: 2019-07-18
    Description: Results of an experimental effort on pulse detonation driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE)/ejector setup that was specifically designed for the study. The results of various experiments designed to probe different aspects of the PDE/ejector setup are reported. The baseline PDE was operated using ethylene (C2H4) as the fuel and an oxygen/nitrogen (O2 + N2) mixture at an equivalence ratio of one. The PDE only experiments included propellant mixture characterization using a laser absorption technique, high fidelity thrust measurements using an integrated spring-damper system, and shadowgraph imaging of the detonation/shock wave structure emanating from the tube. The baseline PDE thrust measurement results are in excellent agreement with experimental and modeling results reported in the literature. These PDE setup results were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups with constant diameter ejector tubes and various detonation tube/ejector tube overlap distances. The results show that for the geometries studied here, a maximum thrust augmentation of 24% is achieved. Further increases are possible by tailoring the ejector geometry based on CFD predictions conducted elsewhere. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2003-4972 , Appendix A. Publications and Presentation Abstracts; 44|39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 20, 2003 - Jul 23, 2003; Huntsville, AL; United States
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  • 48
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2019-07-18
    Description: The objective of this effort is t o develop an efficient and accurate thermo-fluid computational methodology to predict environments for hypothetical thrust chamber design and analysis. The current task scope is to perform multidimensional, multiphysics analysis of thrust performance and heat transfer analysis for a hypothetical solid-core, nuclear thermal engine including thrust chamber and nozzle. The multiphysics aspects of the model include: real fluid dynamics, chemical reactivity, turbulent flow, and conjugate heat transfer. The model will be designed to identify thermal, fluid, and hydrogen environments in all flow paths and materials. This model would then be used to perform non- nuclear reproduction of the flow element failures demonstrated in the Rover/NERVA testing, investigate performance of specific configurations and assess potential issues and enhancements. A two-pronged approach will be employed in this effort: a detailed analysis of a multi-channel, flow-element, and global modeling of the entire thrust chamber assembly with a porosity modeling technique. It is expected that the detailed analysis of a single flow element would provide detailed fluid, thermal, and hydrogen environments for stress analysis, while the global thrust chamber assembly analysis would promote understanding of the effects of hydrogen dissociation and heat transfer on thrust performance. These modeling activities will be validated as much as possible by testing performed by other related efforts.
    Keywords: Spacecraft Propulsion and Power
    Type: Space Technology and Applications International Forum (STAIF-2006); Feb 12, 2006 - Feb 16, 2006; Albuquerque, NM; United States
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  • 49
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-08-17
    Description: Before any rocket is allowed to fly and be used for a manned mission, it is first test-fired on a static test stand to verify its flight readiness. NASA s Stennis Space Center provides testing of Space Shuttle Main Engines, rocket propulsion systems, and related components with several test facilities. It has been NASA s test-launch site since 1961. The testing stations age with time and repeated use; and with aging comes maintenance; and with maintenance comes expense. NASA has been seeking ways to lower the cost of maintaining the stations, and has aided in the development of an improved reliable linear actuator that arrives onsite quickly and costs less money than other actuators. In general terms, a linear actuator is a servomechanism that supplies a measured amount of energy for the operation of another mechanical system. Accuracy, reliability, and speed of the actuator are critical to performance of the entire system, and these actuators are critical components of the engine test stands. Partnership An actuator was developed as part of a Dual-Use Cooperative Agreement between BAFCO, Inc., of Warminister, Pennsylvania, and Stennis. BAFCO identified four suppliers that manufactured actuator components that met the rigorous testing standards imposed by the Space Agency and then modified these components for application on the rocket test stands. In partnership with BAFCO, the existing commercial products size and weight were reworked, reducing cost and delivery time. Previously, these parts would cost between $20,000 and $22,000, but with the new process, they now run between $11,000 and $13,000, a substantial savings, considering NASA has already purchased over 120 of the units. Delivery time of the cost-saving actuators has also been cut from over 20 to 22 weeks to within 8 to 10 weeks. The redesigned actuator is commercially available, and the company is successfully supplying them to customers other than NASA.
    Keywords: Spacecraft Propulsion and Power
    Type: Spinoff 2005; 98-99; NASA/NP-2005-12-419-HQ
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  • 50
    Publication Date: 2019-08-17
    Description: The NASA In-Space Propulsion Technology (ISPT) Program is managed by the NASA Headquarters Science Mission Directorate and is implemented by the Marshall Space Flight Center in Huntsville, Alabama. The ISPT objective is to fund development of promising in- space propulsion technologies that can decrease flight times, decrease cost, or increase delivered payload mass for future science missions. Before ISPT will invest in a technology, the Technology Readiness Level (TRL) of the concept must be estimated to be at TRL 3. A TRL 3 signifies that the technical community agrees that the feasibility of the concept has been proven through experiment or analysis. One of the highest priority technology investments for ISPT is Aerocapture. The aerocapture maneuver uses a planetary atmosphere to reduce or alter the speed of a vehicle allowing for quick, propellantless (or using very little propellant) orbit capture. The atmosphere is used as a brake, transferring the energy associated with the vehicle s high speed into thermal energy. The ISPT Aerocapture Technology Area (ATA) is currently investing in the development of advanced lightweight ablative thermal protection systems, high temperature composite structures, and heat-flux sensors for rigid aeroshells. The heritage of rigid aeroshells extends back to the Apollo era and this technology will most likely be used by the first generation aerocapture vehicle. As a second generation aerocapture technology, ISPT is investing in three inflatable aerodynamic decelerator concepts for planetary aerocapture. They are: trailing ballute (balloon-parachute), attached afterbody ballute, and an inflatable aeroshell. ISPT also leverages the NASA Small Business Innovative Research Program for additional inflatable decelerator technology development. In mid-2004 ISPT requested an independent review of the three inflatable decelerator technologies funded directly by ISPT to validate the TRL and to identify technology maturation concerns. An independent panel with expertise in advanced thin film materials, aerothermodynamics, trajectory design, and inflatable structures was convened to assess the ISPT investments. The panel considered all major technical subsystems including materials, aerothermodynamics, structural dynamics, packaging, and inflation systems. The panel assessed the overall technology readiness of inflatable decelerators to be a 3 and identified fluid-structure interaction, aeroheating, and structural adhesives to be of highest technical concern.
    Keywords: Spacecraft Propulsion and Power
    Type: 18th AIAA Aerodynamic Decelerator Technology Conference and Seminar; May 23, 2005 - May 26, 2005; Munich; Germany
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  • 51
    Publication Date: 2019-08-16
    Description: The constraints of future Exploration Missions will require unique Integrated System Health Management (ISHM) capabilities throughout the mission. An ambitious launch schedule, human-rating requirements, long quiescent periods, limited human access for repair or replacement, and long communication delays all require an ISHM system that can span distinct yet interdependent vehicle subsystems, anticipate failure states, provide autonomous remediation, and support the Exploration Mission from beginning to end. NASA Glenn Research Center has developed and applied health management system technologies to aerospace propulsion systems for almost two decades. Lessons learned from past activities help define the approach to proper ISHM development: sensor selection- identifies sensor sets required for accurate health assessment; data qualification and validation-ensures the integrity of measurement data from sensor to data system; fault detection and isolation-uses measurements in a component/subsystem context to detect faults and identify their point of origin; information fusion and diagnostic decision criteria-aligns data from similar and disparate sources in time and use that data to perform higher-level system diagnosis; and verification and validation-uses data, real or simulated, to provide variable exposure to the diagnostic system for faults that may only manifest themselves in actual implementation, as well as faults that are detectable via hardware testing. This presentation describes a framework for developing health management systems and highlights the health management research activities performed by the Controls and Dynamics Branch at the NASA Glenn Research Center. It illustrates how those activities contribute to the development of solutions for Integrated System Health Management.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2005-214026 , E-15380 , First International Forum on Integrated System Health Engineering and Management in Aerospace; Nov 07, 2005 - Nov 10, 2005; Napa, CA; United States
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  • 52
    Publication Date: 2019-08-16
    Description: A high-efficiency, 110-W(sub e) (watts electric) Stirling Radioisotope Generator (SRG110) for possible use on future NASA Space Science missions is being developed by the Department of Energy, Lockheed Martin, Stirling Technology Company (STC), and NASA Glenn Research Center (GRC). Potential mission use includes providing spacecraft onboard electric power for deep space missions and power for unmanned Mars rovers. GRC is conducting an in-house supporting technology project to assist in SRG110 development. One-, three-, and six-month heater head structural benchmark tests have been completed in support of a heater head life assessment. Testing is underway to evaluate the key epoxy bond of the permanent magnets to the linear alternator stator lamination stack. GRC has completed over 10,000 hours of extended duration testing of the Stirling convertors for the SRG110, and a three-year test of two Stirling convertors in a thermal vacuum environment will be starting shortly. GRC is also developing advanced technology for Stirling convertors, aimed at substantially improving the specific power and efficiency of the convertor and the overall generator. Sunpower, Inc. has begun the development of a lightweight Stirling convertor, under a NASA Research Announcement (NRA) award, that has the potential to double the system specific power to about 8 W(sub e) per kilogram. GRC has performed random vibration testing of a lowerpower version of this convertor to evaluate robustness for surviving launch vibrations. STC has also completed the initial design of a lightweight convertor. Status of the development of a multi-dimensional computational fluid dynamics code and high-temperature materials work on advanced superalloys, refractory metal alloys, and ceramics are also discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2005-213409 , E-14924 , Space Technology and Applications International Forum; Feb 13, 2005 - Feb 17, 2005; Albuquerque, NM; United States
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  • 53
    Publication Date: 2019-08-16
    Description: Performance of two solar sail attitude control implementations is evaluated. One implementation employs four articulated reflective vanes located at the periphery of the sail assembly to generate control torque about all three axes. A second attitude control configuration uses mass on a gimbaled boom to alter the center-of-mass location relative to the center-of-pressure producing roll and pitch torque along with a pair of articulated control vanes for yaw control. Command generation algorithms employ linearized dynamics with a feedback inversion loop to map desired vehicle attitude control torque into vane and/or gimbal articulation angle commands. We investigate the impact on actuator deflection angle behavior due to variations in how the Jacobian matrix is incorporated into the feedback inversion loop. Additionally, we compare how well each implementation tracks a commanded thrust profile, which has been generated to follow an orbit trajectory from the sun-earth L1 point to a sub-L1 station.
    Keywords: Spacecraft Propulsion and Power
    Type: AAS-05-003 , AAS Guidance and Control Conference; Feb 05, 2005 - Feb 09, 2005; Breckenridge, CO; United States
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  • 54
    Publication Date: 2019-08-15
    Description: Lunar habitation modules need electricity and potentially heat to operate. Because of the low amounts of radiation emitted by General Purpose Heat Source (GPHS) modules, power plants incorporating these as heat sources could be placed in close proximity to habitation modules. A design concept is discussed for a high efficiency power plant based on a GPHS assembly integrated with a Stirling convertor. This system could provide both electrical power and heat, if required, for a lunar habitation module. The conceptual GPHS/Stirling system is modular in nature and made up of a basic 5.5 KWe Stirling convertor/GPHS module assembly, convertor controller/PMAD electronics, waste heat radiators, and associated thermal insulation. For the specific lunar application under investigation eight modules are employed to deliver 40 KWe to the habitation module. This design looks at three levels of Stirling convertor technology and addresses the issues of integrating the Stirling convertors with the GPHS heat sources assembly using proven technology whenever possible. In addition, issues related to the high-temperature heat transport system, power management, convertor control, vibration isolation, and potential system packaging configurations to ensure safe operation during all phases of deployment will be discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2005-213991 , AIAA Paper 2005-5716 , E-15315 , 3rd International Energy Conversion Enigeering Conference; Aug 15, 2005 - Aug 18, 2005; San Francisco, CA; United States
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  • 55
    Publication Date: 2019-07-12
    Description: A high-efficiency solid state power amplifier (SSPA) for specific use in a spacecraft is provided. The SSPA has a mass of less than 850 g and includes two different X-band power amplifier sections, i.e., a lumped power amplifier with a single 11-W output and a distributed power amplifier with eight 2.75-W outputs. These two amplifier sections provide output power that is scalable from 11 to 15 watts without major design changes. Five different hybrid microcircuits, including high-efficiency Heterostructure Field Effect Transistor (HFET) amplifiers and Monolithic Microwave Integrated Circuit (MMIC) phase shifters have been developed for use within the SSPA. A highly efficient packaging approach enables the integration of a large number of hybrid circuits into the SSPA.
    Keywords: Spacecraft Propulsion and Power
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  • 56
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2019-08-13
    Description: The Minimum Impulse Thruster (MIT) was developed to improve the state-of-the-art minimum impulse capability of hydrazine monopropellant thrusters. Specifically, a new fast response solenoid valve was developed, capable of responding to a much shorter electrical pulse width, thereby reducing the propellant flow time and the minimum impulse bit. The new valve was combined with the Aerojet MR-103, 0.2 lbf (0.9 N) thruster and put through an extensive Delta-qualification test program, resulting in a factor of 5 reduction in the minimum impulse bit, from roughly 1.1 milli-lbf-seconds (5 milliNewton seconds) to - 0.22 milli-lbf-seconds (1 mN-s). To maintain it's extensive heritage, the thruster itself was left unchanged. The Minimum Impulse Thruster provides mission and spacecraft designers new design options for precision pointing and precision translation of spacecraft.
    Keywords: Spacecraft Propulsion and Power
    Type: 53rd JANNAF Propulsion Meeting; Dec 05, 2005 - Dec 08, 2005; Monterey, CA; United States
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  • 57
    Publication Date: 2019-08-13
    Description: Two transportation architecture changes are presented at either end of a conventional two-stage rocket flight: 1) Air launch using a large, conventional, pod hauler design (i.e., Crossbow)ans 2) Momentum exchange tether (i.e., an in-space asset like MXER). Air launch has ana analytically justified cost reduction of approx. 10%, but its intangible benefits suggest real-world operations cost reductions much higher: 1) Inherent launch safety; 2) Mission Risk Reduction; 3) Favorable payload/rocket limitations; and 4) Leveraging the aircraft for other uses (military transport, commercial cargo, public outreach activities, etc.)
    Keywords: Spacecraft Propulsion and Power
    Type: 2005 JANNAF Conference; Dec 04, 2005 - Dec 08, 2005; Monterey, CA; United States
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  • 58
    Publication Date: 2019-08-13
    Description: This paper describes the structural dynamic tests conducted in-vacuum on the Scalable Square Solar Sail (S(sup 4)) System 20-meter test article developed by ATK Space Systems as part of a ground demonstrator system development program funded by NASA's In-Space Propulsion program. These tests were conducted for the purpose of validating analytical models that would be required by a flight test program to predict in space performance. Specific tests included modal vibration tests on the solar sail system in a 1 Torr vacuum environment using various excitation locations and techniques including magnetic excitation at the sail quadrant corners, piezoelectric stack actuation at the mast roots, spreader bar excitation at the mast tips, and bi-morph piezoelectric patch actuation on the sail cords. The excitation methods are evaluated for their suitability to in-vacuum ground testing and their traceability to the development of on-orbit flight test techniques. The solar sail masts were also tested in ambient atmospheric conditions and these results are also discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: TTA-M-ISP-03-23 , 2005 JANNAF JPM-LPS-SPS Joint Meeting; Dec 05, 2005 - Dec 08, 2005; Monterey, CA; United States
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  • 59
    Publication Date: 2019-08-13
    Description: The NASA In-Space Propulsion Technology (ISPT) Program is managed by the NASA Headquarters Science Mission Directorate and is implemented by the Marshall Space Flight Center in Huntsville, Alabama. The ISPT objective is to fund development of promising in-space propulsion technologies that can decrease flight times, decrease cost, or increase delivered payload mass for future science missions. Before ISPT will invest in a technology, the Technology Readiness Level (TRL) of the concept must be estimated to be at TRL 3. A TRL 3 signifies that the technical community agrees that the feasibility of the concept has been proven through experiment or analysis. One of the highest priority technology investments for ISPT is Aerocapture. The aerocapture maneuver uses a planetary atmosphere to reduce or alter the speed of a vehicle allowing for quick, propellantless (or using very little propellant) orbit capture. The atmosphere is used as a brake, transferring the energy associated with the vehicle's high speed into thermal energy. The ISPT Aerocapture Technology Area (ATA) is currently investing in the development of advanced lightweight ablative thermal protection systems, high temperature composite structures, and heat-flux sensors for rigid aeroshells. The heritage of rigid aeroshells extends back to the Apollo era and this technology will most likely be used by the first generation aerocapture vehicle. As a second generation aerocapture technology, ISPT is investing in three inflatable aerodynamic decelerator concepts for planetary aerocapture. They are: trailing ballute (balloon-parachute), attached afterbody ballute, and an inflatable aeroshell. ISPT also leverages the NASA Small Business Innovative Research Program for additional inflatable decelerator technology development. In mid-2004 ISPT requested an independent review of the three inflatable decelerator technologies funded directly by ISPT to validate the TRL and to identify technology maturation concerns. An independent panel with expertise in advanced thin film materials, aerothermodynamics, trajectory design, and inflatable structures was convened to assess the ISPT investments. The panel considered all major technical subsystems including materials, aerothermodynamics, structural dynamics, packaging, and inflation systems. The panel assessed the overall technology readiness of inflatable decelerators to be a 3 and identified fluid- structure interaction, aeroheating, and structural adhesives to be of highest technical concern.
    Keywords: Spacecraft Propulsion and Power
    Type: SPS Joint Meeting; Dec 05, 2005 - Dec 08, 2005; Monterey, CA; United States|JANNAF 53rd JPM Meeting; Dec 05, 2005 - Dec 08, 2005; Monterey, CA; United States|2nd LPS Meeting; Dec 05, 2005 - Dec 08, 2005; Monterey, CA; United States
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  • 60
    Publication Date: 2019-08-13
    Description: NASA and the U.S. Air Force are working on a joint project to develop a new hydrogen-fueled, full-flow, staged combustion rocket engine. The initial testing and modeling work for the Integrated Powerhead Demonstrator (IPD) project is being performed by NASA Marshall and Stennis Space Centers. A key factor in the testing of this engine is the ability to predict and measure the transient fluid flow during engine start and shutdown phases of operation. A model built by NASA Marshall in the ROCket Engine Transient Simulation (ROCETS) program is used to predict transient engine fluid flows. The model is initially calibrated to data from previous tests on the Stennis E1 test stand. The model is then used to predict the next run. Data from this run can then be used to recalibrate the model providing a tool to guide the test program in incremental steps to reduce the risk to the prototype engine. In this paper, they define this type of model as a calibrated model. This paper proposes a method to estimate the uncertainty of a model calibrated to a set of experimental test data. The method is similar to that used in the calibration of experiment instrumentation. For the IPD example used in this paper, the model uncertainty is determined for both LOX and LH flow rates using previous data. The successful use of this model is then demonstrated to predict another similar test run within the uncertainty bounds. The paper summarizes the uncertainty methodology when a model is continually recalibrated with new test data. The methodology is general and can be applied to other calibrated models.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF Meeting; Dec 05, 2005 - Dec 08, 2005; Monterey, CA; United States
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  • 61
    Publication Date: 2019-08-13
    Description: NASA s Marshall Space Flight Center (MSFC) is well known for its contributions to large ascent propulsion systems such as the Saturn V rocket and the Space Shuttle external tank, solid rocket boosters, and main engines. This paper highlights a lesser known but very rich side of MSFC-its heritage in the development of in-space chemical propulsion systems and its current capabilities for spacecraft propulsion system development and chemical propulsion research. The historical narrative describes the flight development activities associated with upper stage main propulsion systems such as the Saturn S-IVB as well as orbital maneuvering and reaction control systems such as the S-IVB auxiliary propulsion system, the Skylab thruster attitude control system, and many more recent activities such as Chandra, the Demonstration of Automated Rendezvous Technology (DART), X-37, the X-38 de-orbit propulsion system, the Interim Control Module, the US Propulsion Module, and multiple technology development activities. This paper also highlights MSFC s advanced chemical propulsion research capabilities, including an overview of the center s Propulsion Systems Department and ongoing activities. The authors highlight near-term and long-term technology challenges to which MSFC research and system development competencies are relevant. This paper concludes by assessing the value of the full range of aforementioned activities, strengths, and capabilities in light of NASA s exploration missions.
    Keywords: Spacecraft Propulsion and Power
    Type: 53rd JANNAF Joint Propulsion Meeting; Dec 05, 2005 - Dec 08, 2005; Monterey, CA; United States
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  • 62
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-08-13
    Description: The term, propulsion breakthrough, refers to concepts like propellantless space drives and faster-than-light travel, the kind of breakthroughs that would make interstellar exploration practical. Although no such breakthroughs appear imminent, a variety of investigations into these goals have begun. From 1996 to 2002, NASA supported the Breakthrough Propulsion Physics Project to examine physics in the context of breakthrough spaceflight. Three facets of these assessments are now reported: (1) predicting benefits, (2) selecting research, and (3) recent technical progress. Predicting benefits is challenging since the breakthroughs are still only notional concepts, but kinetic energy can serve as a basis for comparison. In terms of kinetic energy, a hypothetical space drive could require many orders of magnitude less energy than a rocket for journeys to our nearest neighboring star. Assessing research options is challenging when the goals are beyond known physics and when the implications of success are profound. To mitigate the challenges, a selection process is described where: (a) research tasks are constrained to only address the immediate unknowns, curious effects or critical issues, (b) reliability of assertions is more important than their implications, and (c) reviewers judge credibility rather than feasibility. The recent findings of a number of tasks, some selected using this process, are discussed. Of the 14 tasks included, six reached null conclusions, four remain unresolved, and four have opportunities for sequels. A dominant theme with the sequels is research about the properties of space, inertial frames, and the quantum vacuum.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2005-213998 , E-15322 , New Trends in Astrodynamics and Applications 2: An International Conference; Jun 03, 2005 - Jun 05, 2005; Princeton, NJ; United States
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  • 63
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2019-08-13
    Description: On NASA's ISTAR RBCC program packaging and performance requirements exceeded traditional H2O2 catalyst bed capabilities. Aerojet refined a high performance, monolithic 90% H202 catalyst bed previously developed and demonstrated. This approach to catalyst bed design and fabrication was an enabling technology to the ISTAR tri-fluid engine. The catalyst bed demonstrated 55 starts at throughputs greater than 0.60 lbm/s/sq in for a duration of over 900 seconds in a physical envelope approximately 114 of traditional designs. The catalyst bed uses photoetched plates of metal bonded into a single piece monolithic structure. The precise control of the geometry and complete mixing results in repeatable, quick starting, high performing catalyst bed. Three different beds were designed and tested, with the best performing bed used for tri-fluid engine testing.
    Keywords: Spacecraft Propulsion and Power
    Type: 40th JANNAF Combustion Meeting; Jun 13, 2005 - Jun 17, 2005; Charleston, NC; United States
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  • 64
    Publication Date: 2019-08-13
    Description: Thermo-nuclear fusion may be the key to a high Isp, high specific power (low alpha) propulsion system. In a fusion system energy is liberated within, and imparted directly to, the propellant. In principle, this can overcome the performance limitations inherent in systems that require thermal power transfer across a material boundary, and/or multiple power conversion stages (NTR, NEP). A thermo-nuclear propulsion system, which attempts to overcome some of the problems inherent in the ORION concept, is described. A passive tapered liner is launched behind a vehicle, through a hole in a pusher-plate, that is connected to the vehicle by a shock-absorbing mechanism. A dense FRC plasmoid is then accelerated to high velocity (in excess of 1,000 km/s) and shot through the hole into the liner, when it has reached a given point down-range. The kinetic energy of the FRC is converted into thermal and magnetic-field energy, igniting a fusion bum in the magnetically confined plasma. The fusion reaction serves as an ignition source for the liner, which is made out of detonable materials. The energy liberated in this process is converted to thrust by the pusher-plate, as in the classic ORION concept. However with this concept, the vehicle does not carry a magazine of pre-fabricated pulse-units. A magnetic nozzle may also be used, in place of the pusher-plate. Estimates of the conditions needed to achieve a sufficient gain will be presented, along with a description of the driver characteristics. The incorporation of this concept into the propulsion system of a spacecraft will also be discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/JPL/MSFC 16th Annual Event Propulsion Workshop; Apr 07, 2005 - Apr 08, 2005; Huntsville, AL; United States
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  • 65
    Publication Date: 2019-08-13
    Description: The Propulsion IVHM Technology Experiment (PITEX) has been an on-going research effort conducted over several years. PITEX has developed and applied a model-based diagnostic system for the main propulsion system of the X-34 reusable launch vehicle, a space-launch technology demonstrator. The application was simulation-based using detailed models of the propulsion subsystem to generate nominal and failure scenarios during captive carry, which is the most safety-critical portion of the X-34 flight. Since no system-level testing of the X-34 Main Propulsion System (MPS) was performed, these simulated data were used to verify and validate the software system. Advanced diagnostic and signal processing algorithms were developed and tested in real-time on flight-like hardware. In an attempt to expose potential performance problems, these PITEX algorithms were subject to numerous real-world effects in the simulated data including noise, sensor resolution, command/valve talkback information, and nominal build variations. The current research has demonstrated the potential benefits of model-based diagnostics, defined the performance metrics required to evaluate the diagnostic system, and studied the impact of real-world challenges encountered when monitoring propulsion subsystems.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2005-213422 , AIAA Paper 2004-6361 , E-14948 , AIAA 1st Intelligent Systems Technical Conference; Sep 20, 2004 - Sep 22, 2004; Chicago, IL; United States
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  • 66
    Publication Date: 2019-07-11
    Description: Spotless days are examined as a predictor for the size and timing of a sunspot cycle. For cycles 16-23 the first spotless day for a new cycle, which occurs during the decline of the old cycle, is found to precede minimum amplitude for the new cycle by about approximately equal to 34 mo, having a range of 25-40 mo. Reports indicate that the first spotless day for cycle 24 occurred in January 2004, suggesting that minimum amplitude for cycle 24 should be expected before April 2007, probably sometime during the latter half of 2006. If true, then cycle 23 will be classified as a cycle of shorter period, inferring further that cycle 24 likely will be a cycle of larger than average minimum and maximum amplitudes and faster than average rise, peaking sometime in 2010.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TP-2005-213608 , M1130
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  • 67
    Publication Date: 2019-07-13
    Description: In 1996 Stennis Space Center was given management authority for all Propulsion Testing for NASA. Over the next few years several research and development (R&D) test facilities were completed and brought up to full operation in what is known as the E-Complex Test Facility at Stennis Space Center. To construct, activate and operate these test facilities, a manual paper-based work control system was created. After utilizing this paper-based work control system for approximately three years, it became apparent that the research and development test area needed a better method to execute, monitor, and report on tasks required to further propulsion testing. The paper based system did not provide the engineers adequate visibility into work tasks or the tracking of testing or hardware discrepancies. This system also restricted the engineer s ability to utilize and access past knowledge and experiences given the severe schedule limitations for most R&D propulsion testing projects. Therefore a system was developed to meet the growing need of Test Operations called the Propulsion Test Directorate (PTD) Work Control System. This system is used to plan, perform, and track tasks that support testing and also to capture lessons learned while doing so.
    Keywords: Spacecraft Propulsion and Power
    Type: SSTI-8080-0003-EPLEX , AIAA Paper 2005-1130 , AIAA Aerodynamic Measurement Technology and Ground Testing Conference; Jan 01, 2005 - Jan 31, 2005; Portland, OR; United States
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  • 68
    Publication Date: 2019-07-13
    Description: The performance of a prototype Hall thruster designed for Discovery-class NASA science mission applications was evaluated at input powers ranging from 0.2 to 2.9 kilowatts. These data were used to construct a throttle profile for a projected Hall thruster system based on this prototype thruster. The suitability of such a Hall thruster system to perform robotic exploration missions was evaluated through the analysis of a near Earth asteroid sample return mission. This analysis demonstrated that a propulsion system based on the prototype Hall thruster offers mission benefits compared to a propulsion system based on an existing ion thruster.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2005-214020 , AIAA Paper 2005-3675 , E-15335 , 41st Joint Propulsion Conference and Exhibit; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 69
    Publication Date: 2019-07-13
    Description: One of the advantages of using a Radioisotope Power System (RPS) for deep space or planetary surface missions is the readily available waste heat, which can be used to maintain electronic components within a controlled temperature range, to warm propulsion tanks and mobility actuators, and to gasify liquid propellants. Previous missions using Radioisotope Thermoelectric Generators (RTGs) dissipated a very large quantity of waste heat due to the relatively low efficiency of the thermoelectric conversion technology. The next generation RPSs, such as the 110-watt Stirling Radioisotope Generator (SRG110) will have much higher conversion efficiencies than their predecessors and therefore may require alternate approaches to transferring waste heat to the spacecraft. RTGs, with efficiencies of approx. 6 to 7% and 200 C housing surface temperatures, would need to use large and heavy radiator heat exchangers to transfer the waste heat to the internal spacecraft components. At the same time, sensitive spacecraft instruments must be shielded from the thermal radiation by using the heat exchangers or additional shields. The SRG110, with an efficiency around 22% and 50 C nominal housing surface temperature, can use the available waste heat more efficiently by more direct heat transfer methods such as heat pipes, thermal straps, or fluid loops. The lower temperatures allow the SRG110 much more flexibility to the spacecraft designers in configuring the generator without concern of overheating nearby scientific instruments, thereby eliminating the need for thermal shields. This paper will investigate using a high efficiency SRG110 for spacecraft thermal management and outline potential methods in several conceptual missions (Lunar Rover, Mars Rover, and Titan Lander) to illustrate the advantages with regard to ease of assembly, less complex interfaces, and overall mass savings.
    Keywords: Spacecraft Propulsion and Power
    Type: 3rd International Energy Conversion Engineering Conference; Aug 15, 2005 - Aug 18, 2005; San Francisco, CA; United States
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  • 70
    Publication Date: 2019-07-13
    Description: Contents include the following: 1. Closed-Brayton-cycle (CBC) thermal energy conversion is one available option for future spacecraft and surface systems. 2. Brayton system conceptual designs for milliwatt to megawatt power converters have been developed 3. Numerous features affect overall optimized power conversion system performance: Turbomachinery efficiency. Heat exchanger effectiveness. Working-fluid composition. Cycle temperatures and pressures.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2005-5700 , 3rd International Energy Conversion Engineering Conference; Aug 15, 2005; San Francisco, CA; United States
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  • 71
    Publication Date: 2019-07-13
    Description: The primary obstacle to any space-based mission is, and has always been, the cost of access to space. Even with impressive efforts toward reusability, no system has come close to lowering the cost a significant amount. It is postulated here, that architectural innovation is necessary to make reusability feasible, not incremental subsystem changes. This paper shows two architectural approaches of reusability that merit further study investments. Both #inherently# have performance increases and cost advantages to make affordable access to space a near term reality. A rocket launched from a subsonic aircraft (specifically the Crossbow methodology) and a momentum exchange tether, reboosted by electrodynamics, offer possibilities of substantial reductions in the total transportation architecture mass - making access-to-space cost-effective. They also offer intangible benefits that reduce risk or offer large growth potential. The cost analysis indicates that approximately a 50% savings is obtained using today#s aerospace materials and practices.
    Keywords: Spacecraft Propulsion and Power
    Type: 2005 Joint Army Navy Nasa Air Force (JANNAF) Conference; Dec 05, 2005 - Dec 08, 2005; Monterey, CA; United States
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  • 72
    Publication Date: 2019-07-13
    Description: Development is underway on a unique high-power solar concentrator array called Stretched Lens Array (SLA) for direct drive electric propulsion. These SLA performance attributes closely match the critical needs of solar electric propulsion (SEP) systems, which may be used for "space tugs" to fuel-efficiently transport cargo from low earth orbit (LEO) to low lunar orbit (LLO), in support of NASA s robotic and human exploration missions. Later SEP systems may similarly transport cargo from the earth-moon neighborhood to the Mars neighborhood. This paper will describe the SLA SEP technology, discuss ground tests already completed, and present plans for future ground tests and future flight tests of SLA SEP systems.
    Keywords: Spacecraft Propulsion and Power
    Type: IAC-05-C3.4-D2.8.04 , 56th International Astronautical Congress; Oct 17, 2005 - Oct 21, 2005; Fukuoka; Japan
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  • 73
    Publication Date: 2019-07-13
    Description: This document is a viewgraph presentation that reviews the Lithium Ion Battery for the Space Technology-5 (ST-5) mission. Included in the document is a review of the ST-5 Mission, a review of the battery requirements, a description of the battery and the battery materials. The testing and the integration and qualification data is reviewed.
    Keywords: Spacecraft Propulsion and Power
    Type: 3rd International Energy Conversion Engineering Conference; Aug 15, 2005 - Aug 18, 2005; San Francisco, CA; United States
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  • 74
    Publication Date: 2019-07-13
    Description: UltraSail is a next-generation high-risk, high-payoff sail system for the launch, deployment, stabilization and control of very large (sq km class) solar sails enabling high payload mass fractions for high (Delta)V. Ultrasail is an innovative, non-traditional approach to propulsion technology achieved by combining propulsion and control systems developed for formation-flying micro-satellites with an innovative solar sail architecture to achieve controllable sail areas approaching 1 sq km, sail subsystem area densities approaching 1 g/sq m, and thrust levels many times those of ion thrusters used for comparable deep space missions. Ultrasail can achieve outer planetary rendezvous, a deep space capability now reserved for high-mass nuclear and chemical systems. One of the primary innovations is the near-elimination of sail supporting structures by attaching each blade tip to a formation-flying micro-satellite which deploys the sail, and then articulates the sail to provide attitude control, including spin stabilization and precession of the spin axis. These tip micro-satellites are controlled by 3-axis micro-thruster propulsion and an on-board metrology system. It is shown that an optimum spin rate exists which maximizes payload mass.
    Keywords: Spacecraft Propulsion and Power
    Type: 41st Joint Propulsion Conference; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 75
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-07-13
    Description: Useful design rules and simple scaling models have been developed for solar sails. Chief among the conclusions are: 1. Sail distortions contribute to the thrust and moments primarily though the mean squared value of their derivatives (slopes), and the sail behaves like a flat sheet if the value is small. The RMS slope is therefore an important figure of merit, and sail distortion effects on the spacecraft can generally be disregarded if the RMS slope is less than about 10% or so. 2. The characteristic slope of the sail distortion varies inversely with the tension in the sail, and it is the tension that produces the principle loading on the support booms. The tension is not arbitrary, but rather is the value needed to maintain the allowable RMS slope. That corresponds to a halyard force about equal to three times the normal force on the supported sail area. 3. Both the AEC/SRS and L Garde concepts appear to be structurally capable of supporting sail sizes up to a kilometer or more with 1AU solar flux, but select transverse dimensions must be changed to do so. Operational issues such as fabrication, handling, storage and deployment will be the limiting factors.
    Keywords: Spacecraft Propulsion and Power
    Type: 41st AIAA Joint Propulsion Conference; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 76
    Publication Date: 2019-07-13
    Description: A 20-meter Scalable Square Solar Sail (S(sup 4)) System was produced and successfully completed functional vacuum testing in NASA Glenn's Space Power Facility at Plum Brook Station Ohio in May 2005. The S(sup 4) system was designed and developed by ATK Space Systems, and the design and production of the Solar Sails for this system was carried out by SRS Technologies. The S(sup 4) system consists of a central structure with four deployable carbon fiber masts that support four triangular sails. SRS has developed an effective and efficient design for triangular sail quadrants that are supported at three points and provide a flat reflective surface with a high fill factor. This sail design is robust enough for deployments in a one atmosphere, one gravity environment and incorporates several advanced features including adhesiveless seaming of membrane strips, compliant edge borders to allow for film membrane cord strain mismatch without causing wrinkling and low mass (3% of total sail mass) ripstop. This paper will outline some of the sail design and fabrication processes and the mature production, packaging and deployment processes that have been developed. This paper will also detail the successful ambient and vacuum testing of the sails and the ATK spacecraft structure. Based on recent experience and testing, SRS is confidant that high Technology Readiness Level (TRL) 5-6 solar sails in the 40-120-meter size range with areal density in the 4-5 grams per square meters (sail minus structure) range can be produced with existing technology. Additional film production research will lead to further reductions in film thickness to less than 1 micron enabling production of sails with areal densities as low as 2.0 grams per square meters using the current design, resulting in a system areal densities as low as 5.3 grams per square meters (sail and structure). These areal densities are low enough to allow nearly all of the Solar Sail missions that have been proposed by the scientific community. The fundamental technologies required to produce these systems has been demonstrated on the 20-meter S(sup 4) sails that have recently completed ground testing demonstrating a mature and technology suitable for incorporation into future flight validation and future mission. Solar Sails can support NASA's Vision for Space Exploration by allowing communication satellite orbits that can maintain continuous communication with the polar regions of the Moon and Mars and to support solar weather monitoring to provide early warning of solar flares and storms that could threaten the safety of astronauts and other spacecraft.
    Keywords: Spacecraft Propulsion and Power
    Type: 41st AIAA/ASME Joint Propulsion Conference and Exhibit; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 77
    Publication Date: 2019-07-13
    Description: The purpose of the STD 6001 test 17 is to determine the flammability of materials in GOX at ambient temperature and at use pressure. The purpose of the new Heated Promoted combustion test is to determine the flammability of material in GOX at use temperature and pressure. The objective is to present the new heated promoted combustion method and show initial data and trends for three representative metals.
    Keywords: Spacecraft Propulsion and Power
    Type: National Space and Missile Materials Symposium; Jun 27, 2005 - Jul 01, 2005; Summerlin, NV; United States
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  • 78
    Publication Date: 2019-07-13
    Description: NASA Marshall Space Flight Center (MSFC) is well known for its contributions to large ascent propulsion systems such as the Saturn V and the Space Shuttle. This paper highlights a lesser known but equally rich side of MSFC - its heritage in spacecraft chemical propulsion systems and its current capabilities for in-space propulsion system development and chemical propulsion research. The historical narrative describes the efforts associated with developing upper-stage main propulsion systems such as the Saturn S-IVB as well as orbital maneuvering and reaction control systems such as the S-IVB auxiliary propulsion system, the Skylab thruster attitude control system, and many more recent activities such as Chandra, the Demonstration of Automated Rendezvous Technology, X-37, the X-38 de-orbit propulsion system, the Interim Control Module, the US Propulsion Module, and several technology development activities. Also discussed are MSFC chemical propulsion research capabilities, along with near- and long-term technology challenges to which MSFC research and system development competencies are relevant.
    Keywords: Spacecraft Propulsion and Power
    Type: 41st AIAA/ASME/ASEE/SAE Joint Propulsion Conference; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 79
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-07-13
    Description: Medium lift EELVs may still play a role in manned space flight. To be considered for manned flight, medium lift EELVs must address the short comings in their current boost assist motors. Two options exist: redesign and requalify the solid rocket motors. Replace solid rocket motors (SRMs) with hybrid rocket motors. Hybrid rocket motors are an attractive alternative. They are safer than SRMs. The TRL's Lockheed Martin Small Launch Vehicle booster development substantially lowers the development risk, cost risk, and the schedule risk for developing hybrid boost assist for EELVs. Hybrid boosters testability offsets SRMs higher inherent reliability.Hybrid booster development and recurring costs are lower than SRMs. Performance gains are readily achieved.
    Keywords: Spacecraft Propulsion and Power
    Type: 41st AIAA/ASME/ASEE/SAE Joint Propulsion Conference; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 80
    Publication Date: 2019-07-13
    Description: Fusion propulsion is inevitable if the human race remains dedicated to exploration of the solar system. There are fundamental reasons why fusion surpasses more traditional approaches to routine crewed missions to Mars, crewed missions to the outer planets, and deep space high speed robotic missions, assuming that reduced trip times, increased payloads, and higher available power are desired. A recent series of informal discussions were held among members from government, academia, and industry concerning fusion propulsion. We compiled a sufficient set of arguments for utilizing fusion in space. .If the U.S. is to lead the effort and produce a working system in a reasonable amount of time, NASA must take the initiative, relying on, but not waiting for, DOE guidance. Arguments for fusion propulsion are presented, along with fusion enabled mission examples, fusion technology trade space, and a proposed outline for future efforts.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2005-4140 , 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 81
    Publication Date: 2019-07-13
    Description: Solar sails reflect photons streaming from the sun and transfer momentum to the sail. The thrust, though small, is continuous and acts for the life of the mission without the need for propellant. Recent advances in materials and ultra-low mass gossamer structures have enabled a host of useful missions utilizing solar sail propulsion. The team of L'Garde, Jet Propulsion Laboratories, Ball Aerospace, and Langley Research Center, under the direction of the NASA In-Space Propulsion office, has been developing a scalable solar sail configuration to address NASA s future space propulsion needs. The baseline design currently in development and testing was optimized around the 1 AU solar sentinel mission. Featuring inflatably deployed sub-T(sub g), rigidized beam components, the 10,000 sq m sail and support structure weighs only 47.5 kg, including margin, yielding an areal density of 4.8 g/sq m. Striped sail architecture, net/membrane sail design, and L'Garde's conical boom deployment technique allows scalability without high mass penalties. This same structural concept can be scaled to meet and exceed the requirements of a number of other useful NASA missions. This paper discusses the interim accomplishments of phase 3 of a 3-phase NASA program to advance the technology readiness level (TRL) of the solar sail system from 3 toward a technology readiness level of 6 in 2005. Under earlier phases of the program many test articles have been fabricated and tested successfully. Most notably an unprecedented 4-quadrant 10 m solar sail ground test article was fabricated, subjected to launch environment tests, and was successfully deployed under simulated space conditions at NASA Plum Brook s 30m vacuum facility. Phase 2 of the program has seen much development and testing of this design validating assumptions, mass estimates, and predicted mission scalability. Under Phase 3 a much larger 20 m square test article including subscale vane has been fabricated and tested. A 20 m system ambient deployment has been successfully conducted after enduring Delta-2 launch environment testing. The program will culminate in a vacuum deployment of a 20 m subscale test article at the NASA Glenn s Plum Brook 30 m vacuum test facility to bring the TRL level as close to 6 as possible in 1 g. This focused program will pave the way for a flight experiment of this highly efficient space propulsion technology.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2005-3927 , 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 11, 2005 - Jul 15, 2005; Tucson, AZ; United States
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  • 82
    Publication Date: 2019-07-13
    Description: To meet the requirements for the 2nd Generation Reusable Launch Vehicle (RLV), a unique propulsion feed system concept was identified using crossfeed between the booster and orbiter stages that could reduce the Two-Stage-to-Orbit (TSTO) vehicle weight and development cost by approximately 25%. A Main Propulsion System (MPS) crossfeed water demonstration test program was configured to address all the activities required to reduce the risks for the MPS crossfeed system. A transient, one-dimensional system simulation was developed for the subscale crossfeed water flow tests. To ensure accurate representation of the crossfeed valve's dynamics in the system model, a high-fidelity, three-dimensional, computational fluid-dynamics (CFD) model was employed. The results from the CFD model were used to specify the valve's flow characteristics in the system simulation. This yielded a crossfeed system model that was anchored to the specific valve hardware and achieved good agreement with the measured test data. These results allowed the transient models to be correlated and validated and used for full scale mission predictions. The full scale model simulations indicate crossfeed is ' viable with the system pressure disturbances at the crossfeed transition being less than experienced by the propulsion system during engine start and shutdown transients.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2005-4370 , 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 83
    Publication Date: 2019-07-13
    Description: NASA's In-Space Propulsion Technology Program is investing in technologies that have the potential to revolutionize the robotic exploration of deep space. For robotic exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs and, in some cases, enable missions previously considered impossible. Continued reliance on conventional chemical propulsion alone will not enable the robust exploration of deep space. The maximum theoretical efficiencies have almost been reached and are insufficient to meet needs for many ambitious science missions currently being considered. By developing the capability to support mid-term robotic mission needs, the program is laying the technological foundation for travel to nearby interstellar space. The In-Space Propulsion Technology Program s technology portfolio includes many advanced propulsion systems. From the next-generation ion propulsion systems operating in the 5-10 kW range, to solar sail propulsion, substantial advances in spacecraft propulsion performance are anticipated. Some of the most promising technologies for achieving these goals use the environment of space itself for energy and propulsion and are generically called "propellantless" because they do not require onboard fuel to achieve thrust. Propellantless propulsion technologies include scientific innovations, such as solar sails, electrodynamic and momentum transfer tethers, and aerocapture. This paper will provide an overview of those propellantless and propellant-based advanced propulsion technologies that will most significantly advance our exploration of deep space.
    Keywords: Spacecraft Propulsion and Power
    Type: 41st Symposiumon Realistic Near-Term Advanced Scientific Space Missions; Jul 04, 2005 - Jul 06, 2005; Aosta; Italy
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  • 84
    Publication Date: 2019-07-13
    Description: Contents include the following: Introduction. Assumptions and scope. Reference mission (National). Propulsion elements amd opportunities for technology insertion. Issues. Summary and conclusion
    Keywords: Spacecraft Propulsion and Power
    Type: Topics in Engineering (TE20), NASA Training Workshop Technologies for Space Exploration; May 17, 2005 - May 18, 2005; Hampton, VA; United States
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  • 85
    Publication Date: 2019-07-13
    Description: At the sea level, a phenomenon common with all rocket engines, especially for a highly over-expanded nozzle, during ignition and shutdown is that of flow separation as the plume fills and empties the nozzle, Since the flow will be separated randomly. it will generate side loads, i.e. non-axial forces. Since rocket engines are designed to produce axial thrust to power the vehicles, it is not desirable to be excited by non-axial input forcing functions, In the past, several engine failures were attributed to side loads. During the development stage, in order to design/size the rocket engine components and to reduce the risks, the local dynamic environments as well as dynamic interface loads have to be defined. The methodology developed here is the way to determine the peak loads and shock environments for new engine components. In the past it is not feasible to predict the shock environments, e.g. shock response spectra, from one engine to the other, because it is not scaleable. Therefore, the problem has been resolved and the shock environments can be defined in the early stage of new engine development. Additional information is included in the original extended abstract.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Conference; Apr 18, 2005; Austin, TX; United States
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  • 86
    Publication Date: 2019-07-13
    Description: Solar electric propulsion (SEP) is being used for a variety of planetary missions sponsored by ESA, JAXA, and NASA and nuclear electric propulsion (NEP) is being considered for future, flagship-class interplanetary missions. Radioisotope electric propulsion (REP) has recently been shown to effectively complement SEP and NEP for missions to high-AU targets with modest payload requirements. This paper investigates the application of an advanced REP for a sample return from the comet Tempel 1. A set of mission and system parameters are varied with the goal of quantifying their impact on total mission payload. Mission parameters considered include trip-time and Earth return entry interface speed of the sample return system. System parameters considered include launch vehicle, power level of spacecraft at beginning of mission, and thruster specific impulse. For the baseline case of Atlas 401 and REP power level of 750 W, the mission time was 12 years, the payload was 144 kg, and the missions optimized to a single specific impulse generally within Hall ion thruster range. Other cases were investigated in support of graduate studies, and include the larger Atlas 551 launch vehicle and extended power level to 1 kW. The Atlas 551 cases tended to optimize dual specific impulses generally in the Hall ion thruster range for both legs of the mission. A power level of at least 1-kW and trip-time of approximately 11 years was required to obtain a total science payload close to 320 kg for the Atlas 401 launch vehicle. An Atlas 551 launch vehicle yielded a science payload of approximately 540 kg for the case of 1-kW of power and an 11-year trip time, and nearly 250 kg of science payload for the case of 1-kW of power and a 6-year trip time. Results are also reported indicating the performance ramifications of meeting a reduced Earth entry interface velocity constraint.
    Keywords: Spacecraft Propulsion and Power
    Type: 41st AIAA Joint Propulsion Conference; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 87
    Publication Date: 2019-07-13
    Description: The Certification of Propulsion Systems is costly and complex which involves development and qualification testing. The desire of the certification process is to assure all requirements can be demonstrated to be compliant. The purpose of this paper is to address the technical design concerns of certifying a system for flight. The authors of this paper have experience the lessons learned from supporting the Shuttle Program for Main Propulsion and On Orbit Propulsions Systems. They have collaborated design concerns for certifying propulsion systems. Presented are Pressurization, Tankage, Feed System and Combustion Instability concerns. Propulsion System Engineers are challenged with the dilemma for testing new systems to specific levels to reduce risk yet maintain budgetary targets. A methodical approach is presented to define the types of test suitable to address the technical issues for qualifying systems for retiring the risk levels.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2005-4314 , 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 10, 2005 - Jul 13, 2005; Tucson, AZ; United States
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  • 88
    Publication Date: 2019-07-13
    Description: In July 2004, a 10-meter solar sail structure developed by L Garde, Inc. was tested in vacuum at the NASA Glenn 30-meter Plum Brook Space Power Facility in Sandusky, Ohio. The three main objections of the test were to demonstrate unattended deployment from a stowed configuration, to measure the deployed shape of the sail at both ambient and cryogenic room temperatures, and to measure the deployed structural dynamic characteristics (vibration modes). This paper summarizes the work conducted to fulfill the second test objective. The deployed shape was measured photogrammetrically in vacuum conditions with four 2-megapixel digital video cameras contained in custom made pressurized canisters. The canisters included high-intensity LED ring lights to illuminate a grid of retroreflective targets distributed on the solar sail. The test results closely matched pre-test photogrammetry numerical simulations and compare well with ABAQUS finite-element model predictions.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2005-1889 , 46th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference; Apr 18, 2005 - Apr 21, 2005; Austin, TX; United States
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  • 89
    Publication Date: 2019-07-13
    Description: A simulation and modeling effort is conducted on gas foil thrust bearings. A foil bearing is a self acting hydrodynamic device capable of separating stationary and rotating components of rotating machinery by a film of air or other gaseous lubricant. Although simple in appearance these bearings have proven to be complicated devices in analysis. They are sensitive to fluid structure interaction, use a compressible gas as a lubricant, may not be in the fully continuum range of fluid mechanics, and operate in the range where viscous heat generation is significant. These factors provide a challenge to the simulation and modeling task. The Reynolds equation with the addition of Knudsen number effects due to thin film thicknesses is used to simulate the hydrodynamics. The energy equation is manipulated to simulate the temperature field of the lubricant film and combined with the ideal gas relationship, provides density field input to the Reynolds equation. Heat transfer between the lubricant and the surroundings is also modeled. The structural deformations of the bearing are modeled with a single partial differential equation. The equation models the top foil as a thin, bending dominated membrane whose deflections are governed by the biharmonic equation. A linear superposition of hydrodynamic load and compliant foundation reaction is included. The stiffness of the compliant foundation is modeled as a distributed stiffness that supports the top foil. The system of governing equations is solved numerically by a computer program written in the Mathematica computing environment. Representative calculations and comparisons with experimental results are included for a generation I gas foil thrust bearing.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2005-213811 , E-15168 , 2005 Annual Meeting and Exhibition, 60th Society of Tribologists and Lubrication Engineers; May 15, 2005 - May 19, 2005; Las Vegas, NV; United States
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  • 90
    Publication Date: 2019-07-13
    Description: High voltage, high power electron bombardment ion thrusters needed for deep space missions will be required to be operated for long durations in space as well as during ground laboratory life testing. Carbon based ion optics are being considered for such thrusters. The sputter deposition of carbon and arc vaporized carbon flakes from long duration operation of ion thrusters can result in deposition on insulating surfaces, causing them to become conducting. Because the sticking coefficient is less than one, secondary deposition needs to be considered to assure that shorting of critical components does not occur. The sticking coefficient for sputtered carbon and arc vaporized carbon is measured as well as directional ejection distribution data for carbon that does not stick upon first impact.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2005-213798 , AIAA Paper 2005-4413 , E-15156 , Joint Propulsion Conference; Jul 10, 2005 - Jul 13, 2005; Tuscon, AZ; United States
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  • 91
    Publication Date: 2019-07-13
    Description: The NASA Glenn Research Center in-house computer model Closed Cycle Engine Program (CCEP) was used to explore the design trade space and off-design performance characteristics of 100 kWe-class recuperated Closed Brayton Cycle (CBC) power conversion systems. Input variables for a potential design point included the number of operating units (1, 2, 4), cycle peak pressure (0.5, 1, 2 MPa), and turbo-alternator shaft speed (30, 45, 60 kRPM). The design point analysis assumed a fixed turbine inlet temperature (1150 K), compressor inlet temperature (400 K), working-fluid molecular weight (40 g/mol), compressor pressure ratio (2.0), recuperator effectiveness (0.95), and a Sodium-Potassium (NaK) pumped-loop radiator. The design point options were compared on the basis of thermal input power, radiator area, and mass. For a nominal design point with defined Brayton components and radiator area, off-design cases were examined by reducing turbine inlet temperature (as low as 900 K), reducing shaft speed (as low as 50% of nominal), and circulating a percentage (up to 20%) of the compressor exit flow back to the gas cooler. The off-design examination sought approaches to reduce thermal input power without freezing the radiator.
    Keywords: Spacecraft Propulsion and Power
    Type: E-15105 , Space Technology and Applications International Forum (STAIF-2005); Feb 13, 2005 - Feb 15, 2005; Albuquerque, NM; United States|University of New Mexico''s Institute for Space and Nuclear Power Studies (ISNPS) Conference Proceedings; Feb 13, 2005 - Feb 15, 2005; Albuquerque, NM; United States
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  • 92
    Publication Date: 2019-07-13
    Description: NASA is undertaking the design of a new spacecraft to explore the planet Jupiter and its three moons Calisto, Ganymede and Europa. This proposed mission, known as Jupiter Icy Moons Orbiter (JIMO) would use a nuclear reactor and an associated electrical generation system (Reactor Power Plant-RPP) to provide power to the spacecraft. The JIMO spacecraft is envisioned to use this power for science and communications as well as Electric Propulsion (EP). Among other potential power-generating concepts, previous studies have considered Thermoelectric and Brayton Power conversion systems, coupled to a liquid metal reactor for the JIMO mission. This paper will explore trades in system mass and radiator area for a nuclear reactor power conversion system, however this study will focus on Stirling power conversion. The Stirling convertor modeled in this study is based upon the Component Test Power Convertor design that was designed and operated successfully under the Civil Space Technology Initiative for use with the SP-100 nuclear reactor i the 1980's and early 1990's. The study design is such that two of the four convertors would operate at any time to generate the 100 kWe while the others are held in reserve. For this study the Stirling convertors hot-side temperature is 1050 K, would operate at a temperature ratio of 2.4 for a minimum mass system and would have a system efficiency of 29%. The Stirling convertor would generate high voltage (400 volt), 100 Hz single phase AC that is supplied to the Power Management and Distribution system. The waste hear is removed from the Stirling convertors by a flowing liquid sodium-potassium eutectic and then rejected by a shared radiator. The radiator consists of two coplanar wings, which would be deployed after the reactor is in space. System trades were performed to vary cycle state point temperatures and convertor design as well as power output. Other redundancy combinations were considered to understand the affects of convertor size and number of spares to the system mass.
    Keywords: Spacecraft Propulsion and Power
    Type: University of New Mexico''s Institute for Space and Nuclear Power Studies (UNM-ISNPS); Feb 13, 2005 - Feb 17, 2005; Albuquerque, NM; United States|Space Technology and Applications International Forum (STAIF-2005); Feb 13, 2005 - Feb 17, 2005; Albuquerque, NM; United States
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  • 93
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    Unknown
    In:  CASI
    Publication Date: 2019-07-13
    Description: Fusion propulsion is inevitable if the human race remains dedicated to exploration of the solar system. There are fundamental reasons why fusion surpasses more traditional approaches to routine crewed missions to Mars, crewed missions to the outer planets, and deep space high speed robotic missions, assuming that reduced trip times, increased payloads, and higher available power are desired. A recent series of informal discussions were held among members from government, academia, and industry concerning fusion propulsion. We compiled a sufficient set of arguments for utilizing fusion in space. If the U.S. is to lead the effort and produce a working system in a reasonable amount of time, NASA must take the initiative, relying on, but not waiting for, DOE guidance. In this talk those arguments for fusion propulsion are presented, along with fusion enabled mission examples, fusion technology trade space, and a proposed outline for future efforts.
    Keywords: Spacecraft Propulsion and Power
    Type: MSFC/JPL Advanced Propulsion Conference; Apr 07, 2005 - Apr 08, 2005; Huntsville, AL; United States
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  • 94
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-07-13
    Description: The need for cryogenic fuel tanks continues to expand, and research at NASA Marshall Space Flight Center (MSFC) is addressing these needs. This viewgraph presentation provides an overview of composite tank development, including tank testing, cryogenic materials research, tank liners, and dual-walled tanks, at MSFC.
    Keywords: Spacecraft Propulsion and Power
    Type: SAMPE Conference; May 02, 2005 - May 06, 2005; Long Beach, CA; United States
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  • 95
    Publication Date: 2019-07-13
    Description: The characterization of the electromagnetic interaction for a solar sail in the solar wind environment, and identification of viable charging mitigation strategies, is a critical solar sail mission design task, as spacecraft charging has important implications both for science applications and for sail lifetime. To that end, we have performed surface charging calculations of a candidate 150-meter-class solar sail spacecraft for the 0.5 solar polar orbit and a 1.0 AU L1 orbit. We construct a model of the spacecraft with candidate materials having appropriate electrical properties using Object Toolkit and perform the spacecraft charging analysis using NASCAP-2k, the NASA/AFRL sponsored spacecraft charging analysis tool. We use nominal and atypical solar wind environments appropriate for the 0.5 AU and 1.0 AU missions to establish current collection of solar wind ions and electrons. In addition, we include a geostationary orbit case to demonstrate a bounding example of extreme (negative) charging of a solar sail spacecraft in the geostationary orbit environment. Results form the charging analysis demonstrate that minimal differential potentials (and resulting threat of electrostatic discharge) occur when the spacecraft is constructed entirely of conducting materials, as expected. Examples with dielectric materials exposed to the space environment exhibit differential potentials ranging from a few volts to extreme potentials in the kilovolt range.
    Keywords: Spacecraft Propulsion and Power
    Type: 9th Spacecraft Charging Technology Conference; Apr 04, 2005 - Apr 08, 2005; Tsukuba; Japan
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  • 96
    Publication Date: 2019-07-13
    Description: During a normal inspection of the main propulsion system at Kennedy Space Center, small cracks were noticed near a slotted region of a gimbal joint flowliner located just upstream from one of the Space Shuttle Main Engines (SSME). These small cracks sparked an investigation of the entire Space Shuttle fleet main propulsion feedlines. The investigation was initiated to determine the cause of the small cracks and a repair method that would be needed to return the Shuttle fleet back to operation safely. The cracks were found to be initiated by structural resonance caused by flow fluctuations from the SSME low pressure fuel turbopump interacting with the flowliner. The pump induced backward traveling wakes that excited the liner and duct acoustics which also caused the liner to vibrate in complex mode shapes. The investigation involved an extensive effort by a team of engineers from the NASA civil servant and contractor workforce with the goal to characterize the root cause of the cracking behavior of the fuel side gimbal joint flowliners. In addition to working to identify the root cause, a parallel path was taken to characterize the material properties and fatigue capabilities of the liner material such that the life of the liners could be ascertained. As the characterization of the material and the most probable cause matured, the combination of the two with pump speed restrictions provided a means to return the Shuttle to flight in a safe manner. This paper traces the flowliner investigation results with respect to the structural dynamics analysis, component level testing and hot-fire flow testing on a static testbed. The paper will address the unique aspects of a very complex problem involving backflow from a high performance pump that has never been characterized nor understood to such detail. In addition, the paper will briefly address the flow phenomena that excited the liners, the unique structural dynamic modal characteristics and the variability of SSME operation which has ultimately ensured the safe and reliable operation of the shuttle main engines for each flight.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2005-1862 , AIAA Structures, Structural Dynamics and Materials Conference; Apr 18, 2005 - Apr 21, 2005; Austin, TX; United States
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  • 97
    Publication Date: 2019-07-13
    Description: The characterization of the electromagnetic interaction for a solar sail in the solar wind environment and identification of viable charging mitigation strategies are critical solar sail mission design task. Spacecraft charging has important implications both for science applications and for lifetime and reliability issues of sail propulsion systems. To that end, surface charging calculations of a candidate 150-meter-class solar sail spacecraft for the 0.5 AU solar polar and 1.0 AU L1 solar wind environments are performed. A model of the spacecraft with candidate materials having appropriate electrical properties is constructed using Object Toolkit. The spacecraft charging analysis is performed using Nascap-2k, the NASA/AFRL sponsored spacecraft charging analysis tool. Nominal and atypical solar wind environments appropriate for the 0.5 AU and 1.0 AU missions are used to establish current collection of solar wind ions and electrons. Finally, a geostationary orbit environment case is included to demonstrate a bounding example of extreme (negative) charging of a solar sail spacecraft. Results from the charging analyses demonstrate that minimal differential potentials (and resulting threat of electrostatic discharge) occur when the spacecraft is constructed entirely of conducting materials, as anticipated from standard guidelines for mitigation of spacecraft charging issues. Examples with dielectric materials exposed to the space environment exhibit differential potentials ranging from a few volts to extreme potentials in the kilovolt range.
    Keywords: Spacecraft Propulsion and Power
    Type: Solar Sail Technology and Applications Conference; Apr 04, 2005 - Apr 08, 2005; Greenbelt, MD; United States
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  • 98
    Publication Date: 2019-07-13
    Description: Nuclear electric propulsion (NEP) vehicles will be needed for future manned missions to Mars and beyond. Candidate vehicles must be identified through trade studies for further detailed design from a large array of possibilities. Genetic algorithms have proven their utility in conceptual design studies by effectively searching a large design space to pinpoint unique optimal designs. This research combines analysis codes for NEP subsystems with genetic algorithm-based optimization. Trade studies for a NEP reference mission to the asteroids were conducted to identify important trends, and to determine the effects of various technologies and subsystems on vehicle performance. It was found that the electric thruster type and thruster performance have a major impact on the achievable system performance, and that significant effort in thruster research and development is merited.
    Keywords: Spacecraft Propulsion and Power
    Type: 1st Space Exploration Conference; Jan 30, 2005 - Feb 01, 2005; Orlando, FL; United States
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  • 99
    Publication Date: 2019-07-13
    Description: Advanced power is one of the key capabilities that will be needed to achieve NASA's missions of exploration and scientific advancement. Significant gaps exist in advanced power capabilities that are on the critical path to enabling human exploration beyond Earth orbit and advanced robotic exploration of the solar system. Focused studies and investment are needed to answer key development issues for all candidate technologies before down-selection. The viability of candidate power technology alternatives will be a major factor in determining what exploration mission architectures are possible. Achieving the capabilities needed to enable the CEV, Moon, and Mars missions is dependent on adequate funding. Focused investment in advanced power technologies for human and robotic exploration missions is imperative now to reduce risk and to make informed decisions on potential exploration mission decisions beginning in 2008. This investment would begin the long lead-time needed to develop capabilities for human exploration missions in the 2015 to 2030 timeframe. This paper identifies some of the key technologies that will be needed to fill these power capability gaps. Recommendations are offered to address capability gaps in advanced power for Crew Exploration Vehicle (CEV) power, surface nuclear power systems, surface mobile power systems, high efficiency power systems, and space transportation power systems. These capabilities fill gaps that are on the critical path to enabling robotic and human exploration missions. The recommendations address the following critical technology areas: Energy Conversion, Energy Storage, and Power Management and Distribution.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2005-213600 , AIAA Paper 2005-2786 , E-15067 , First Space Exploration Conference: Continuing the Voyage of Discovery; Jan 30, 2005 - Feb 01, 2005; Orlando, FL; United States
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  • 100
    Publication Date: 2019-07-13
    Description: In recent years there has been increased interest in the development of a new generation of high performance boost rocket engines. These efforts, which will represent a substantial advancement in boost engine technology over that developed for the Space Shuttle Main Engines in the early 1970s, are being pursued both at NASA and the United States Air Force. NASA, under its Space Launch Initiative s Next Generation Launch Technology Program, is investigating the feasibility of developing a highly reliable, long-life, liquid oxygen/kerosene (RP-1) rocket engine for launch vehicles. One of the top technical risks to any engine program employing hydrocarbon fuels is the potential for fuel thermal stability and material compatibility problems to occur under the high-pressure, high-temperature conditions required for regenerative fuel cooling of the engine combustion chamber and nozzle. Decreased heat transfer due to carbon deposits forming on wetted fuel components, corrosion of materials common in engine construction (copper based alloys), and corrosion induced pressure drop increases have all been observed in laboratory tests simulating rocket engine cooling channels. To mitigate these risks, the knowledge of how these fuels behave in high temperature environments must be obtained. Currently, due to the complexity of the physical and chemical process occurring, the only way to accomplish this is empirically. Heated tube testing is a well-established method of experimentally determining the thermal stability and heat transfer characteristics of hydrocarbon fuels. The popularity of this method stems from the low cost incurred in testing when compared to hot fire engine tests, the ability to have greater control over experimental conditions, and the accessibility of the test section, facilitating easy instrumentation. These benefits make heated tube testing the best alternative to hot fire engine testing for thermal stability and heat transfer research. This investigation used the Heated Tube Facility at the NASA Glenn Research Center to perform a thermal stability and heat transfer characterization of RP-1 in an environment simulating that of a high chamber pressure, regenerative cooled rocket engine. The first step in the research was to investigate the carbon deposition process of previous heated tube experiments by performing scanning electron microscopic analysis in conjunction with energy dispersive spectroscopy on the tube sections. This analysis gave insight into the carbon deposition process and the effect that test conditions played in the formation of deleterious coke. Furthermore, several different formations were observed and noted. One other crucial finding of this investigation was that in sulfur containing hydrocarbon fuels, the interaction of the sulfur components with copper based wall materials presented a significant corrosion problem. This problem in many cases was more life limiting than those posed by the carbon deposition process. The results of this microscopic analysis was detailed and presented at the December 2003 JANNAF Air-Breathing Propulsion Meeting as a Materials Compatibility and Thermal Stability Analysis of common Hydrocarbon Fuels (reference 1).
    Keywords: Spacecraft Propulsion and Power
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