Tungsten and barium transport in the internal plasma of hollow cathodes

The effect of tungsten erosion, transport, and redeposition on the operation of dispenser hollow cathodes was investigated in detailed examinations of the discharge cathode inserts from 8200 h and 30 352 h ion engine wear tests. Erosion and subsequent redeposition of tungsten in the electron emission zone at the downstream end of the insert reduce the porosity of the tungsten matrix, preventing the flow of barium from the interior. This inhibits the interfacial reactions of the barium-calcium-aluminate impregnant with the tungsten in the pores. A numerical model of barium transport in the internal xenon discharge plasma shows that the barium required to reduce the work function in the emission zone can be supplied from upstream through the gas phase. Barium that flows out of the pores of the tungsten insert is rapidly ionized in the xenon discharge and pushed back to the emitter surface by the electric field and drag from the xenon ion flow. This barium ion flux is sufficient to maintain a barium surface coverage at the downstream end greater than 0.6, even if local barium production at that point is inhibited by tungsten deposits. The model also shows that the neutral barium pressure exceeds the equilibrium vapor pressure of the impregnant decomposition reaction over much of the insert length, so the reactions are suppressed. Only a small region upstream of the zone blocked by tungsten deposits is active and supplies the required barium. These results indicate that hollow cathode failure models based on barium depletion rates in vacuum dispenser cathodes are very conservative

NASA Propulsion Investments for Exploration and Science

The National Aeronautics and Space Administration (NASA) invests in chemical and electric propulsion systems to achieve future mission objectives for both human exploration and robotic science. Propulsion system requirements for human missions are derived from the exploration architecture being implemented in the Constellation Program. The Constellation Program first develops a system consisting of the Ares I launch vehicle and Orion spacecraft to access the Space Station, then builds on this initial system with the heavy-lift Ares V launch vehicle, Earth departure stage, and lunar module to enable missions to the lunar surface. A variety of chemical engines for all mission phases including primary propulsion, reaction control, abort, lunar ascent, and lunar descent are under development or are in early risk reduction to meet the specific requirements of the Ares I and V launch vehicles, Orion crew and service modules, and Altair lunar module. Exploration propulsion systems draw from Apollo, space shuttle, and commercial heritage and are applied across the Constellation architecture vehicles. Selection of these launch systems and engines is driven by numerous factors including development cost, existing infrastructure, operations cost, and reliability. Incorporation of green systems for sustained operations and extensibility into future systems is an additional consideration for system design. Science missions will directly benefit from the development of Constellation launch systems, and are making advancements in electric and chemical propulsion systems for challenging deep space, rendezvous, and sample return missions. Both Hall effect and ion electric propulsion systems are in development or qualification to address the range of NASA s Heliophysics, Planetary Science, and Astrophysics mission requirements. These address the spectrum of potential requirements from cost-capped missions to enabling challenging high delta-v, long-life missions. Additionally, a high specific impulse chemical engine is in development that will add additional capability to performance-demanding space science missions. In summary, the paper provides a survey of current NASA development and risk reduction propulsion investments for exploration and science

Overview of Al-based nanoenergetic ingredients for solid rocket propulsion

The introduction of nano-sized energetic ingredients first occurred in Russia about 60 years ago and arose great expectations in the rocket propulsion community, thanks to the higher energy densities and faster energy release rates exhibited with respect to conventional ingredients. But, despite intense worldwide research programs, still today mostly laboratory level applications are reported and often for scientific purposes only. A number of practical reasons prevent the applications at industrial level: inert native coating of the energetic particles, nonuniform dispersion, aging, excessive viscosity of the slurry propellant, possible limitations in mechanical properties, more demanding safety issues, cost, and so on. This paper describes the main features in terms of performance of solid rocket propellants loaded with nanometals and intends to emphasize the unique properties or operating conditions made possible by the addition of the nano-sized energetic ingredients. Steady and unsteady combustion regimes are examined. Keywords: Nanoaluminum, Solid rocket propellant, Burning rate, Combustion, Propulsion, Performanc

Green Comparable Alternatives of Hydrazines-Based Monopropellant and Bipropellant Rocket Systems

Concepts are presented for “green” (with reduced hazards) replacements for monopropellant hydrazine propulsion systems and for hypergolic bipropellant systems while maintaining similar performance. At the onset of the “green propulsion” age, “green” alternatives to hydrazine propulsion have been emerging. The introduction rate of these into space systems is very slow due to the conservatism of the space propulsion industry. The concept presented here for monopropellant hydrazine systems offers gradual conversion to “green propellants” by dual capability of conventional hydrazine systems and ammonium dinitramide (ADN)-based systems. An initial risk reduction program has been carried out for materializing the concept. It includes proof of concept of dual use of all propulsion system parts. Materials compatibility and actual operation have been demonstrated. For bipropellants, we present the emerging “green” hypergolic system based on kerosene and peroxide, similar in performance to MMH/N2O4. Results of the proof-of-concept and development model systems are presented. The experimental results of various engine types demonstrate the capability to operate in both pulse and steady-state modes and the ability to produce different thrust levels. The fuel and oxidizer show very robust hypergolicity and short ignition delay times, as well as characteristic velocity efficiency exceeding 98%

Summary of Propulsion System Needs in Support of Project Constellation

In January 2004, the President of the United States established the Vision for Space Exploration (VSE) to return man to the moon and ultimately to extend manned space travel to Mars. This paper will summarize the manned space flight liquid propulsion system needs in support of Project Constellation over the next 10 years. It will include all engine needs to return man to the moon. An overview of engines currently under contract, those baselined but not yet under contract, and those engine needs that hav.e yet to be initiated. Project Constellation includes the components as shown Figure 1. Liquid propulsion systems supporting the manned portion of these elements include the following: the Crew Exploration Vehicle named Orion (crew module reaction control system (CMRCS), service module Orion Main Engine (OME), service module auxiliary RCS, and service module reaction control system (SMRCS)), the Crew Launch Vehicle named Ares 1 (J2X upper stage, first stage roll control system, second stage reaction control system, and the Ares I-X roll control system), the Heavy Lift Launch Vehicle named Ares V (RS68B first stage booster, J-2X upper stage, roll control systems, and the Earth Departure Stage (EDS) (powered by the same Ares V Upper Stage J-2X), and the Lunar Lander named Altair with both descent and ascent stages (lunar orbit insertion and descent main engine, ascent main engine, and attitude control systems for both stages). In addition, there may be additional engine needs for early demonstrators, but those will not be speculated on as part of this paper. Also, other portions of the VSE architecture, including the planned Orion abort demonstrations and the Lunar Precursor Robotic Program, are not addressed here as they either use solid motors or are focused on unmanned precursor missions

A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing

Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review

Liquid-Propellant Rocket Engines (LREs) are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable LREs can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable LREs can also continuously follow the most economical thrust curve in a given situation, compared to discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an LRE as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of LRE throttling centered around engines from the United States. Several LRE throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared

Heat Transfer Processes for Hydrogen and Methane in Cooling Channels of Regeneratively Cooled Thrust Chambers of Cryogenic Rocket Engines

Die enormen Temperaturen und Wärmeströme in einer Raketenbrennkammer machen eine aktive Kühlung der Brennkammerstruktur unabdingbar. Die Regenerativkühlung, bei der der Treibstoff vor der Verbrennung durch die Struktur geleitet wird, ist eine sehr effiziente und weitverbreitete Methode der Kühlung. Kühlkanäle mit einem hohen Aspektverhältnis (Höhe zu Breite Verhältnis) können zu einer besseren Kühlung bei gleichzeitig geringerem Druckverlust führen. Die thermische Schichtung, die bei dieser Art von Kühlkanälen auftreten kann, wirkt dem positiven Effekt entgegen und limitiert das Aspektverhältnis. In der vorliegenden Arbeit werden experimentelle Untersuchungen zur regenerativen Kühlung mit Wasserstoff und Methan bei für Raketenbrennkammern repräsentativen Bedingungen vorgestellt und ausgewertet. Das verwendete Brennkammersegment ist in Umfangsrichtung in vier Quadranten unterteilt, wobei in jeden Quadrant Kühlkanäle mit einem anderen Aspektverhältnis eingebracht worden sind. Für die Auswertung der experimentellen Daten wurde eine inverse Methode verwendet, die es ermöglicht, anhand der gemessenen Strukturtemperaturen den lokalen Wärmestrom und Wärmeübergangskoeffizienten zu bestimmen. Die thermische Schichtung aufgrund von mangelnder Durchmischung in Kühlkanälen mit hohem Aspektverhältnis tritt sowohl bei Wasserstoff als auch bei Methan auf. Für Wasserstoff ist die Ausprägung allerdings deutlich größer. Der Einfluss auf die Heißgaswandtemperatur ist allerdings für beide Kühlmedien vergleichsweise gering. Bei Methan kann es aufgrund der Nähe zum kritischen Punkt zur Trennung einer heißen gasartigen Schicht an der Wand und der kalten flüssigartigen Kernströmung kommen. Dieser heat transfer deterioration (HTD) genannte Effekt führt zu einem lokalen Abfall des Wärmeübergangs und einem drastischen Anstieg der Heißgaswandtemperatur. Die wesentlichen Einflussfaktoren auf diesen Effekt sind der Druck bzw. die Nähe zum kritischen Punkt, das Verhältnis von Wärmestromdichte und Massenstrom pro Fläche qw/G sowie das Aspektverhältnis. Basierend auf den Experimenten wurden numerische Simulationen durchgeführt, die die Ergebnisse stützen und erweitern. Die Auswertung dieser Simulationen zeigt eine weitestgehend gute Übereinstimmung mit den experimentellen Ergebnissen. In der Nähe des kritischen Punktes und insbesondere wenn es zu HTD kommt, zeigen die Simulationen eine systematische Abweichung und ein deutliches Überschätzen der Heißgaswandtemperatur