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  • Spacecraft Propulsion and Power  (1,025)
  • 2000-2004  (1,025)
  • 1
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    Taming Liquid Hydrogen: The Centaur Upper Stage Rocket, 1958-2002 (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-28
    Description: During its maiden voyage in May 1962, a Centaur upper stage rocket, mated to an Atlas booster, exploded 54 seconds after launch, engulfing the rocket in a huge fireball. Investigation revealed that Centaur's light, stainless-steel tank had split open, spilling its liquid-hydrogen fuel down its sides, where the flame of the rocket exhaust immediately ignited it. Coming less than a year after President Kennedy had made landing human beings on the Moon a national priority, the loss of Centaur was regarded as a serious setback for the National Aeronautics and Space Administration (NASA). During the failure investigation, Homer Newell, Director of Space Sciences, ruefully declared: "Taming liquid hydrogen to the point where expensive operational space missions can be committed to it has turned out to be more difficult than anyone supposed at the outset." After this failure, Centaur critics, led by Wernher von Braun, mounted a campaign to cancel the program. In addition to the unknowns associated with liquid hydrogen, he objected to the unusual design of Centaur. Like the Atlas rocket, Centaur depended on pressure to keep its paper-thin, stainless-steel shell from collapsing. It was literally inflated with its propellants like a football or balloon and needed no internal structure to give it added strength and stability. The so-called "pressure-stabilized structure" of Centaur, coupled with the light weight of its high- energy cryogenic propellants, made Centaur lighter and more powerful than upper stages that used conventional fuel. But, the critics argued, it would never become the reliable rocket that the United States needed.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/SP-2004-4230 , LC-2004-042092
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  • 2
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    Fabrication process for combustion chamber/nozzle assembly (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-27
    Description: An integral, lightweight combustion chamber/nozzle assembly for a rocket engine has a refractory metal shell defining a chamber of generally frusto-conical contour. The shell communicates at its smaller end with a rocket body, and terminates at its larger end in a generally contact contour, which is open at its terminus and which serves as a nozzle for the rocket engine. The entire inner surface of the refractory metal shell has a thermal and oxidation barrier layer applied thereto. An ablative silica phenolic insert is bonded to the exposed surface of the thermal and oxidation barrier layer. The ablative phenolic insert provides a chosen inner contour for the combustion chamber and has a taper toward the open terminus of the nozzle. A process for fabricating the integral, lightweight combustion chamber/nozzle assembly is simple and efficient, and results in economy in respect of both resources and time.
    Keywords: Spacecraft Propulsion and Power
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  • 3
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    Injector for liquid fueled rocket engine (2000)
    In:  CASI
    Details
    Publication Date: 2019-08-27
    Description: An injector for liquid fueled rocket engines wherein a generally flat core having a frustoconical dome attached to one side of the core to serve as a manifold for a first liquid, with the core having a generally circular configuration having an axis. The other side of the core has a plurality of concentric annular first slots and a plurality of annular concentric second slots alternating with the first slots, the second slots having a greater depth than said first slots. A bore extends through the core for inletting a second liquid into said core, the bore intersecting the second slots to feed the second liquid into the second slots. The core also has a plurality of first passageways leading from the manifold to the first annular slots for feeding the first liquid into said first slots. A faceplate brazed to said other side of the core is provided with apertures extending from the first and second slots through said face plate, these apertures being positioned to direct fuel and liquid oxygen into contact with each other in the combustion chamber. The first liquid may be liquid oxygen and the second liquid may be kerosene or liquid hydrogen.
    Keywords: Spacecraft Propulsion and Power
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  • 4
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    In-Space Propulsion Program Overview and Status (2003)
    In:  Other Sources
    Details
    Publication Date: 2019-08-27
    Description: NASA's In-Space Propulsion (ISP) Program is designed to develop advanced propulsion technologies that can enable or greatly enhance near and mid-term NASA science missions by significantly reducing cost, mass, and/or travel times. These technologies include: Solar Electric Propulsion, Aerocapture, Solar Sails, Momentum Exchange Tethers, Plasma Sails and other technologies such as Advanced Chemical Propulsion. The ISP Program intends to develop cost-effective propulsion technologies that will provide a broad spectrum of mission possibilities, enabling NASA to send vehicles on longer, more useful voyages and in many cases to destinations that were previously unreachable using conventional means. The ISP approach to identifying and prioritizing these most promising technologies is to use mission and system analysis and subsequent peer review. The ISP program seeks to develop technologies under consideration to Technology Readiness Level (TRL) -6 for incorporation into mission planning within 3-5 years of initiation. The NASA TRL 6 represents a level where a technology is ready for system level demonstration in a relevant environment, usually a space environment. In addition, maximum use of open competition is encouraged to seek optimum solutions under ISP. Several NASA Research Announcements (NRA's) have been released asking industry, academia and other organizations to propose propulsion technologies designed to improve our ability to conduct scientific study of the outer planets and beyond. The ISP Program is managed by NASA Headquarters Office of Space Science and implemented by the Marshall Space Flight Center in Huntsville, Alabama.
    Keywords: Spacecraft Propulsion and Power
    Type: International Electric Propulsion Conference 2003; Mar 17, 2003 - Mar 21, 2003; Toulouse; France
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  • 5
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    NASA Development of Aerocapture Technologies (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Aeroassist technology development is a vital part of the NASA In-Space Propulsion Program (ISP), which is managed by the NASA Headquarters Office of Space Science, and implemented by the Marshall Space Flight Center in Huntsville, Alabama. Aeroassist is the general term given to various techniques to maneuver a space vehicle within an atmosphere, using aerodynamic forces in lieu of propulsive fuel. Within the ISP, the current aeroassist technology development focus is aerocapture. The objective of the ISP Aerocapture Technology Project (ATP) is to develop technologies that can enable and/or benefit NASA science missions by significantly reducing cost, mass, and/or travel times. To accomplish this objective, the ATP identifies and prioritizes the most promising technologies using systems analysis, technology advancement and peer review, coupled with NASA Headquarters Office of Space Science target requirements. Plans are focused on developing mid-Technology Readiness Level (TRL) technologies to TRL 6 (ready for technology demonstration in space).
    Keywords: Spacecraft Propulsion and Power
    Type: 36th Annual Division for Planetary Science; Nov 08, 2004 - Nov 10, 2004; Louisville, KY; United States
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  • 6
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    Modeling and Validation of a Propellant Mixer for Controller Design (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A mixing chamber used in rocket engine testing at the NASA Stennis Space Center is modelled by a system of two nonlinear ordinary differential equations. The mixer is used to condition the thermodynamic properties of cryogenic liquid propellant by controlled injection of the same substance in the gaseous phase. The three inputs of the mixer are the positions of the valves regulating the liquid and gas flows at the inlets, and the position of the exit valve regulating the flow of conditioned propellant. Mixer operation during a test requires the regulation of its internal pressure, exit mass flow, and exit temperature. A mathematical model is developed to facilitate subsequent controller designs. The model must be simple enough to lend itself to subsequent feedback controller design, yet its accuracy must be tested against real data. For this reason, the model includes function calls to thermodynamic property data. Some structural properties of the resulting model that pertain to controller design, such as uniqueness of the equilibrium point, feedback linearizability and local stability are shown to hold under conditions having direct physical interpretation. The existence of fixed valve positions that attain a desired operating condition is also shown. Validation of the model against real data is likewise provided.
    Keywords: Spacecraft Propulsion and Power
    Type: SE-2002-12-00083-SSC
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  • 7
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    In-Space Chemical Propulsion System Model (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Multiple, new technologies for chemical systems are becoming available and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic space missions. A mass model for pressure-fed, Earth and space-storable, advanced chemical propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystem. The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with characteristics of current interest.
    Keywords: Spacecraft Propulsion and Power
    Type: 36th Annual Division for Planetary Science; Nov 08, 2004 - Nov 10, 2004; Louisville, KY; United States
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  • 8
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    Performance Expectations of Closed-Brayton-Cycle Heat Exchangers in 100-kWe Nuclear Space Power Systems (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Performance expectations of closed-Brayton-cycle heat exchangers to be used in 100-kWe nuclear space power systems were forecast. Proposed cycle state points for a system supporting a mission to three of Jupiter s moons required effectiveness values for the heat-source exchanger, recuperator and rejection exchanger (gas cooler) of 0.98,0.95 and 0.97, respectively. Performance parameters such as number of thermal units (Nm), equivalent thermal conductance (UA), and entropy generation numbers (Ns) varied from 11 to 19,23 to 39 kWK, and 0.019 to 0.023 for some standard heat exchanger configurations. Pressure-loss contributions to entropy generation were significant; the largest frictional contribution was 114% of the heat-transfer irreversibility. Using conventional recuperator designs, the 0.95 effectiveness proved difficult to achieve without exceeding other performance targets; a metallic, plate-fin counterflow solution called for 15% more mass and 33% higher pressure-loss than the target values. Two types of gas-coolers showed promise. Single-pass counterflow and multipass cross-counterflow arrangements both met the 0.97 effectiveness requirement. Potential reliability-related advantages of the cross-countefflow design were noted. Cycle modifications, enhanced heat transfer techniques and incorporation of advanced materials were suggested options to reduce system development risk. Carbon-carbon sheeting or foam proved an attractive option to improve overall performance.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2003-212597 , AIAA Paper 2003-5956 , NAS 1.15:212597 , E-14139 , First International Energy Conversion Engineering Conference; Aug 17, 2003 - Aug 21, 2003; Portsmouth, VA; United States
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  • 9
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    NASA In-Space Propulsion Technology Program: Overview and Update (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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 they are insufficient to meet needs for many ambitious science missions currently being considered. The In-Space Propulsion Technology Program's technology portfolio includes many advanced propulsion systems. From the next-generation ion propulsion system operating in the 5- to 10-kW range to aerocapture and solar sails, 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, aeroassist and aerocapture. This paper will provide an overview of both propellantless and propellant-based advanced propulsion technologies, as well as NASA's plans for advancing them as part of the In-Space Propulsion Technology Program.
    Keywords: Spacecraft Propulsion and Power
    Type: 36th Annual Division for Planetary Science; Nov 08, 2004 - Nov 10, 2004; Louisville, KY; United States
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  • 10
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    Effect of Random Thermal Spikes on Stirling Convertor Heater Head Reliability (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Onboard radioisotope power systems being developed to support future NASA exploration missions require reliable design lifetimes of up to 14 yr and beyond. The structurally critical heater head of the high-efficiency developmental Stirling power converter has undergone extensive computational analysis of operating temperatures (up to 650 C), stresses, and creep resistance of the thin-walled Inconel 718 bill of material. Additionally assessment of the effect of uncertainties in the creep behavior of the thin-walled heater head, the variation in the manufactured thickness, variation in control temperature, and variation in pressure on the durability and reliability were performed. However, it is possible for the heater head to experience rare incidences of random temperature spikes (excursions) of short duration. These incidences could occur randomly with random magnitude and duration during the desired mission life. These rare incidences could affect the creep strain rate and therefore the life. The paper accounts for these uncertainties and includes the effect of such rare incidences, random in nature, on the reliability. The sensitivities of variables affecting the reliability are quantified and guidelines developed to improve the reliability are outlined. Furthermore, the quantified reliability is being verified with test data from the accelerated benchmark tests being conducted at the NASA Glenn Research Center.
    Keywords: Spacecraft Propulsion and Power
    Type: E-14918 , International Energy Conversion Engineering Conference; Aug 16, 2004 - Aug 19, 2004; Providence, RI; United States
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  • 11
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    Reliability Assessment Approach for Stirling Convertors and Generators (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Stirling power conversion is being considered for use in a Radioisotope Power System for deep-space science missions because it offers a multifold increase in the conversion efficiency of heat to electric power. Quantifying the reliability of a Radioisotope Power System that utilizes Stirling power conversion technology is important in developing and demonstrating the capability for long-term success. A description of the Stirling power convertor is provided, along with a discussion about some of the key components. Ongoing efforts to understand component life, design variables at the component and system levels, related sources, and the nature of uncertainties is discussed. The requirement for reliability also is discussed, and some of the critical areas of concern are identified. A section on the objectives of the performance model development and a computation of reliability is included to highlight the goals of this effort. Also, a viable physics-based reliability plan to model the design-level variable uncertainties at the component and system levels is outlined, and potential benefits are elucidated. The plan involves the interaction of different disciplines, maintaining the physical and probabilistic correlations at all the levels, and a verification process based on rational short-term tests. In addition, both top-down and bottom-up coherency were maintained to follow the physics-based design process and mission requirements. The outlined reliability assessment approach provides guidelines to improve the design and identifies governing variables to achieve high reliability in the Stirling Radioisotope Generator design.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2004-213078 , E-14554 , Space Technology and Applications International Forum (STAIF-2003); Feb 02, 2003 - Feb 05, 2003; Albuquerque, NM; United States
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  • 12
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    EPS (Electric Particulate Suspension) Microgravity Technology Provides NASA with New Tools (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The Electric Particulate Suspension is a fire safety ignition test system being developed at Iowa State University with NASA support for evaluating combustion properties of powders, powder-gas mixtures, and pure gases in microgravity and gravitational atmospheres (quenching distance, ignition energy, flammability limits). A separate application is the use of EPS technology to control heat transfer in vacuum and space environment enclosures. In combustion testing, ignitable powders (aluminum, magnesium) are introduced in the EPS test cell and ignited by spark, while the addition of inert particles act as quenching media. As a combustion research tool, the EPS method has potential as a benchmark design for quenching powder flames that would provide NASA with a new fire safety standard for powder ignition testing. The EPS method also supports combustion modeling by providing accurate measurement of flame-quenching distance as an important parameter in laminar flame theory since it is closely related to characteristic flame thickness and flame structure. In heat transfer applications, inert powder suspensions (copper, steel) driven by electric fields regulate heat flow between adjacent surfaces enclosures both in vacuum (or gas) and microgravity. This simple E-field control can be particularly useful in space environments where physical separation is a requirement between heat exchange surfaces.
    Keywords: Spacecraft Propulsion and Power
    Type: Strategic Research to Enable NASA's Exploration Missions Conference and Workshop: Poster Session, Volume 2; 52-53; NASA/CP-2004-213205/VOL2
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  • 13
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    Development of a Dynamic, End-to-End Free Piston Stirling Convertor Model (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A dynamic model for a free-piston Stirling convertor is being developed at the NASA Glenn Research Center. The model is an end-to-end system model that includes the cycle thermodynamics, the dynamics, and electrical aspects of the system. The subsystems of interest are the heat source, the springs, the moving masses, the linear alternator, the controller, and the end-user load. The envisioned use of the model will be in evaluating how changes in a subsystem could affect the operation of the convertor. The model under development will speed the evaluation of improvements to a subsystem and aid in determining areas in which most significant improvements may be found. One of the first uses of the end-toend model will be in the development of controller architectures. Another related area is in evaluating changes to details in the linear alternator.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2004-212941 , E-14381 , Space Technology and Applications International Forum; Feb 02, 2003 - Feb 05, 2003; Albuquerque, NM; United States
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  • 14
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    Stirling Technology Development at NASA GRC (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The Department of Energy, Stirling Technology Company (STC), and NASA Glenn Research Center (NASA Glenn) are developing a free-piston Stirling convertor for a high-efficiency Stirling Radioisotope Generator (SRG) for NASA Space Science missions. The SRG is being developed for multimission use, including providing electric power for unmanned Mars rovers and deep space missions. NASA Glenn is conducting an in-house technology project to assist in developing the convertor for space qualification and mission implementation. Recent testing, of 55-We Technology Demonstration Convertors (TDC's) built by STC includes mapping, of a second pair of TDC's, single TDC testing, and TDC electromagnetic interference and electromagnetic compatibility characterization on a nonmagnetic test stand. Launch environment tests of a single TDC without its pressure vessel to better understand the convertor internal structural dynamics and of dual-opposed TDC's with several engineering mounting structures with different natural frequencies have recently been completed. A preliminary life assessment has been completed for the TDC heater head, and creep testing of the IN718 material to be used for the flight convertors is underway. Long-term magnet aging tests are continuing to characterize any potential aging in the strength or demagnetization resistance of the magnets used in the linear alternator (LA). Evaluations are now beginning on key organic materials used in the LA and piston/rod surface coatings. NASA Glenn is also conducting finite element analyses for the LA, in part to look at the demagnetization margin on the permanent magnets. The world's first known integrated test of a dynamic power system with electric propulsion was achieved at NASA Glenn when a Hall-effect thruster was successfully operated with a free-piston Stirling power source. Cleveland State University is developing a multidimensional Stirling computational fluid dynamics code to significantly improve Stirling loss predictions and assist in identifying convertor areas for further improvements. This paper will update the status and results for these efforts.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211315/REV1 , E-13119/REV1 , NAS 1.15:211315/REV1 , Space Technology and Applications International Forum; Feb 03, 2002 - Feb 07, 2002; Albuquerque, NM; United States
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  • 15
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    Mars Mission Surface Operation Simulation Testing of Lithium-Ion Batteries (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The objectives of this program are to 1) Assess viability of using lithium-ion technology for future NASA applications, with emphasis upon Mars landers and rovers which will operate on the planetary surface; 2) Support the JPL 2003 Mars Exploration Rover program to assist in the delivery and testing of a 8 AHr Lithium-Ion battery (Lithion/Yardney) which will power the rover; 3) Demonstrate applicability of using lithium-ion technologyfor future Mars applications: Mars 09 Science Laboratory (Smart Lander) and Future Mars Surface Operations (General). Mission simulation testing was carried out for cells and batteries on the Mars Surveyor 2001 Lander and the 2003 Mars Exploration Rover.
    Keywords: Spacecraft Propulsion and Power
    Type: The 2002 NASA Aerospace Battery Workshop; NASA/CP-2003-212344
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  • 16
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    A Propulsion System for a Scientific Micro/Nanospacecraft (2000)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A preliminary propulsion system design for a 10-100 kg. micro/nanospacecraft for generic scientific missions is described. The design drivers resulting in this proposed system are discussed as well as current technology developments.
    Keywords: Spacecraft Propulsion and Power
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  • 17
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    Ram Burn Observations (RAMBO) (2002)
    In:  Other Sources
    Details
    Publication Date: 2019-08-15
    Description: Ram Burn Observations (RAMBO) is a Department of Defense experiment that observes shuttle Orbital Maneuvering System engine burns for the purpose of improving plume models. On STS-107 the appropriate sensors will observe selected rendezvous and orbit adjust burns.
    Keywords: Spacecraft Propulsion and Power
    Type: STS 107 Shuttle Press Kit: Providing 24/7 Space Science Research; 96
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  • 18
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    NASA GRC Stirling Technology Development Overview (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The Department of Energy, Lockheed Martin (LM), Stirling Technology Company, and NASA Glenn Research Center (GRC) are developing a high-efficiency Stirling Radioisotope Generator (SRG) for potential NASA Space Science missions. The SRG is being developed for multimission use, including providing spacecraft onboard electric power for NASA deep space missions and power for unmanned Mars rovers. NASA GRC is conducting an in- house supporting technology project to assist in developing the Stirling convertor for space qualification and mission implementation. Preparations are underway for a thermalhacuum system demonstration and unattended operation during endurance testing of the 55-We Technology Demonstration Convertors. Heater head life assessment efforts continue, including verification of the heater head brazing and heat treatment schedules and evaluation of any potential regenerator oxidation. Long-term magnet aging tests are continuing to characterize any possible aging in the strength or demagnetization resistance of the permanent magnets used in the linear alternator. Testing of the magnet/lamination epoxy bond for performance and lifetime characteristics is now underway. These efforts are expected to provide key inputs as the system integrator, LM, begins system development of the SRG. GRC is also developing advanced technology for Stirling convertors. Cleveland State University (CSU) is progressing toward a multi-dimensional Stirling computational fluid dynamics code, capable of modeling complete convertors. Validation efforts at both CSU and the University of Minnesota are complementing the code development. New efforts have been started this year on a lightweight convertor, advanced controllers, high-temperature materials, and an end-to-end system dynamics model. Performance and mass improvement goals have been established for second- and third-generation Stirling radioisotope power systems.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2003-212454 , E-13976 , NAs 1.15:212454 , Space Technology and Applications International Forum (STAIF-2003); Feb 02, 2003 - Feb 05, 2003; Albuquerque, NM; United States
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  • 19
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    A Power Conversion Concept for the Jupiter Icy Moons Orbiter (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The Jupiter Icy Moons Orbiter (JIMO) mission is currently under study by the Office of Space Science under the Project Prometheus Program. JIMO is examining the use of Nuclear Electric Propulsion (NEP) to carry scientific payloads to three Jovian moons. A potential power system concept includes dual 100 kWe Brayton converters, a deployable pumped loop heat rejection subsystem, and a 400 Vac Power Management and Distribution (PMAD) bus. Many trades were performed in aniving at this candidate power system concept. System-level studies examined design and off-design operating modes, determined startup requirements, evaluated subsystem redundancy options, and quantified the mass and radiator area of reactor power systems from 20 to 200 kWe. In the Brayton converter subsystem, studies were performed to investigate converter packaging options, and assess the induced torque effects on spacecraft dynamics due to rotating machinery. In the heat rejection subsystem, design trades were conducted on heat transport approaches, material and fluid options, and deployed radiator geometries. In the PMAD subsystem, the overall electrical architecture was defined and trade studies examined distribution approaches, voltage levels, and cabling options.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2003-212596 , E-14153 , NAS 1.15:212596 , AIAA Paper 2003-6007 , First International Energy Conversion Engineering Conference; Aug 17, 2003 - Aug 21, 2003; Portsmouth, VA; United States
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  • 20
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    Gyro-based Maximum-Likelihood Thruster Fault Detection and Identification (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: When building smaller, less expensive spacecraft, there is a need for intelligent fault tolerance vs. increased hardware redundancy. If fault tolerance can be achieved using existing navigation sensors, cost and vehicle complexity can be reduced. A maximum likelihood-based approach to thruster fault detection and identification (FDI) for spacecraft is developed here and applied in simulation to the X-38 space vehicle. The system uses only gyro signals to detect and identify hard, abrupt, single and multiple jet on- and off-failures. Faults are detected within one second and identified within one to five accords,
    Keywords: Spacecraft Propulsion and Power
    Type: 2002 American Control Conference; May 08, 2002 - May 10, 2002; Anchorage, AL; United States
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  • 21
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    Asteroids as Propulsion Systems of Space Ships (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Currently, rockets are used to change the trajectory of space ships and probes. This method is very expensive and requires a lot of fuel, which limits the feasibility of space stations, interplanetary space ships, and probes. Sometimes space probes use the gravity field of a planet However, there am only nine planets in the Solar System, all separated by great distances. There are tons of millions of asteroids in outer space. This paper offers a revolutionary method for changing the trajectory of space probes. The method uses the kinetic or rotary energy of asteroids, comet nuclei, meteorites or other space bodies (small planets, natural planetary satellites, space debris, etc.) to increase (to decrease) ship (probe) speed up to 1000 m/sec (or more) and to achieve any new direction in outer space. The flight possibilities of space ships and probes are increased by a factor of millions.
    Keywords: Spacecraft Propulsion and Power
    Type: World Space Congress 2002; Oct 10, 2002 - Oct 19, 2002; Houston, TX; United States|JBIS; 56; 98-107
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  • 22
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    Laser Transmission Measurements of Soot Extinction Coefficients in the Exhaust Plume of the X-34 60k-lb Thrust Fastrac Rocket Engine (2000)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A four-channel laser transmissometer has been used to probe the soot content of the exhaust plume of the X-34 60k-lb thrust Fastrac rocket engine at NASA's Marshall Space Flight Center. The transmission measurements were made at an axial location about equal 1.65 nozzle diameters from the exit plane and are interpreted in terms of homogeneous radial zones to yield extinction coefficients from 0.5-8.4 per meter. The corresponding soot mass density, spatially averaged over the plume cross section, is, for Rayleigh particles, approximately equal to 0.7 micrograms/cubic cm and alternative particle distributions are briefly considered. Absolute plume radiance at the laser wavelength (515 nm) is estimated from the data at approximately equal to 2.200 K equivalent blackbody temperature, and temporal correlations in emission from several spatial locations are noted.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TP-2000-210075 , M-972 , NAS 1.60:210075
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    Damping Goes the Distance in Golf (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: In the late 1980s, Dr. Benjamin Dolgin of NASA s Jet Propulsion Laboratory developed a concept for a high-damping graphite/viscoelastic material for the Strategic Defense Initiative (popularly referred to as "Star Wars"), as part of a space-based laser anti-missile program called "Asterix." Dolgin drummed up this concept with the intention of stabilizing weapons launch platforms in space, where there is no solid ground to firmly support these structures. Without the inclusion of high-damping material, the orbital platforms were said to vibrate for 20 minutes after force was applied - a rate deemed "unacceptable" by leaders of the Strategic Defense Initiative.
    Keywords: Spacecraft Propulsion and Power
    Type: Spinoff; 45-46; NASA/NP-2004-10-374-HQ
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    Diesel Goes Digital to Save Energy, Lower Emissions (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Forty years ago, actuators requiring constant energy to help power the Apollo spacecraft in space were replaced by magnetically holding and releasing, electronically controlled valves. Today, these same magnetic, electronic valves are on the verge of replacing entire camshaft systems in cars and trucks on Earth, thus leading to a whole new generation of low-emission engines.
    Keywords: Spacecraft Propulsion and Power
    Type: Spinoff; 63-64; NASA/NP-2004-10-374-HQ
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    Overview 2004 of NASA-Stirling Convertor CFD Model Development and Regenerator R and D Efforts (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: This paper reports on accomplishments in 2004 in (1) development of Stirling-convertor CFD models at NASA Glenn and via a NASA grant, (2) a Stirling regenerator-research effort being conducted via a NASA grant (a follow-on effort to an earlier DOE contract), and (3) a regenerator-microfabrication contract for development of a "next-generation Stirling regenerator." Cleveland State University is the lead organization for all three grant/contractual efforts, with the University of Minnesota and Gedeon Associates as subcontractors. Also, the Stirling Technology Company and Sunpower, Inc. are both involved in all three efforts, either as funded or unfunded participants. International Mezzo Technologies of Baton Rouge, Louisiana is the regenerator fabricator for the regenerator-microfabrication contract. Results of the efforts in these three areas are summarized.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2004-213404 , E-14912 , Space Technology and Applications International Forum; Feb 13, 2005 - Feb 17, 2005; Albuquerque, NM; United States
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    Modeling of Spacecraft Advanced Chemical Propulsion Systems (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: This paper outlines the development of the Advanced Chemical Propulsion System (ACPS) model for Earth and Space Storable propellants. This model was developed by the System Technology Operation of SAIC-Huntsville for the NASA MSFC In-Space Propulsion Project Office. Each subsystem of the model is described. Selected model results will also be shown to demonstrate the model's ability to evaluate technology changes in chemical propulsion systems.
    Keywords: Spacecraft Propulsion and Power
    Type: 36th Annual Division for Planetary Science; Nov 06, 2004 - Nov 10, 2004; Louisville, KY; United States
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    In-Space Propulsion (ISP) Solar Sail Propulsion Technology Development (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An overview of the rationale and content for Solar Sail Propulsion (SSP), the on-going project to advance solar technology from technology readiness level 3 to 6 will be provided. A descriptive summary of the major and minor component efforts underway will include identification of the technology providers and a listing of anticipated products Recent important results from major system ground demonstrators will be provided. Finally, a current status of all activities will provided along with the most recent roadmap for the SSP technology development program.
    Keywords: Spacecraft Propulsion and Power
    Type: 36th Annual Division for Planetary Sciences; Nov 08, 2004 - Nov 10, 2004; Louisville, KY; United States
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    Review of Computational Stirling Analysis Methods (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Nuclear thermal to electric power conversion carries the promise of longer duration missions and higher scientific data transmission rates back to Earth for both Mars rovers and deep space missions. A free-piston Stirling convertor is a candidate technology that is considered an efficient and reliable power conversion device for such purposes. While already very efficient, it is believed that better Stirling engines can be developed if the losses inherent its current designs could be better understood. However, they are difficult to instrument and so efforts are underway to simulate a complete Stirling engine numerically. This has only recently been attempted and a review of the methods leading up to and including such computational analysis is presented. And finally it is proposed that the quality and depth of Stirling loss understanding may be improved by utilizing the higher fidelity and efficiency of recently developed numerical methods. One such method, the Ultra HI-Fl technique is presented in detail.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2004-213300 , AIAA Paper 2004-5582 , E-14747 , Second International Energy Conversion Engineering Conference; Aug 16, 2004 - Aug 19, 2004; Providence, RI; United States
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    Solar Sail Models and Test Measurements Correspondence for Validation Requirements Definition (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Solar sails are being developed as a mission-enabling technology in support of future NASA science missions. Current efforts have advanced solar sail technology sufficient to justify a flight validation program. A primary objective of this activity is to test and validate solar sail models that are currently under development so that they may be used with confidence in future science mission development (e.g., scalable to larger sails). Both system and model validation requirements must be defined early in the program to guide design cycles and to ensure that relevant and sufficient test data will be obtained to conduct model validation to the level required. A process of model identification, model input/output documentation, model sensitivity analyses, and test measurement correspondence is required so that decisions can be made to satisfy validation requirements within program constraints.
    Keywords: Spacecraft Propulsion and Power
    Type: Solar Sail Technology and Applications Conference; Sep 26, 2004 - Sep 29, 2004; Greenbelt, MD; United States
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    Advanced Technology Development for Stirling Convertors (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A high-efficiency Stirling Radioisotope Generator (SRG) for use on potential NASA Space Science missions is being developed by the Department of Energy, Lockheed Martin, Stirling Technology Company, and NASA Glenn Research Center (GRC). These missions may include providing spacecraft onboard electric power for deep space missions or power for unmanned Mars rovers. GRC is also developing advanced technology for Stirling convertors, aimed at substantially improving the specific power and efficiency of the convertor and the overall power system. Performance and mass improvement goals have been established for second- and thirdgeneration Stirling radioisotope power systems. Multiple efforts are underway to achieve these goals, both in-house at GRC and under various grants and contracts. The status and results to date for these efforts will be discussed in this paper. Cleveland State University (CSU) is developing a multi-dimensional Stirling computational fluid dynamics code, capable of modeling complete convertors. A 2-D version of the code is now operational, and validation efforts at both CSU and the University of Minnesota are complementing the code development. A screening of advanced superalloy, refractory metal alloy, and ceramic materials has been completed, and materials have been selected for creep and joining characterization as part of developing a high-temperature heater head. A breadboard characterization is underway for an advanced controller using power electronics for active power factor control with a goal of eliminating the heavy tuning capacitors that are typically needed to achieve near unity power factors. Key Stirling developments just initiated under recent NRA (NASA Research Announcement) awards will also be discussed. These include a lightweight convertor to be developed by Sunpower Inc. and an advanced microfabricated regenerator to be done by CSU.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2004-213186 , E-14685 , Space Technology and Applications International Forum; Feb 08, 2004 - Feb 12, 2004; Albuquerque, NM; United States
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    Space Solar Power Demonstrations: Challenges and Progress (2002)
    In:  Other Sources
    Details
    Publication Date: 2019-08-14
    Description: The prospects of using electrical power beamed from space are coming closer to reality with the continued pursuit and improvements in the supporting space solar research and technology. Space Solar Power (SSP) has been explored off and on for approximately three decades as a viable alternative and clean energy source. Results produced through the more recent Space Solar Power Exploratory Research and Technology (SERT) program involving extensive participation by industry, universities, and government has provided a sound technical basis for believing that technology can be improved to the extent that SSP systems can be built, economically feasible, and successfully deployed in space. Considerable advancements have been made in conceptual designs and supporting technologies including solar power generation, wireless power transmission, power management distribution, thermal management and materials, and the integrated systems engineering assessments. Basic technologies have progressed to the point were the next logical step is to formulate and conduct sophisticated demonstrations involving prototype hardware as final proof of concepts and identify high end technology readiness levels in preparation for full scale SSP systems designs. In addition to continued technical development issues, environmental and safety issues must be addressed and appropriate actions taken to reassure the public and prepare them for the future use of this alternative renewable energy resource. Accomplishing these objectives will allow informed future decisions regarding further SSP and related R&D investments by both NASA management and prospective external partners. In particular, accomplishing these objectives will also guide further definition of SSP and related technology roadmaps including performance objectives, resources and schedules; including 'multi-purpose' applications (terrestrial markets, science, commercial development of space, and other government missions).
    Keywords: Spacecraft Propulsion and Power
    Type: 53rd International Astro. Congress; Oct 10, 2002 - Oct 19, 2002; Houston, TX; United States
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    Space Solar Power Technology Demonstration for Lunar Polar Applications: Laser-Photovoltaic Wireless Power Transmission (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-14
    Description: Space Solar Power technology offers unique benefits for near-term NASA space science missions, which can mature this technology for other future applications. "Laser-Photo-Voltaic Wireless Power Transmission" (Laser-PV WPT) is a technology that uses a laser to beam power to a photovoltaic receiver, which converts the laser's light into electricity. Future Laser-PV WPT systems may beam power from Earth to satellites or large Space Solar Power satellites may beam power to Earth, perhaps supplementing terrestrial solar photo-voltaic receivers. In a near-term scientific mission to the moon, Laser-PV WPT can enable robotic operations in permanently shadowed lunar polar craters, which may contain ice. Ground-based technology demonstrations are proceeding, to mature the technology for this initial application, in the moon's polar regions.
    Keywords: Spacecraft Propulsion and Power
    Type: International Astronautical Congress; Oct 17, 2002; Houston, TX; United States
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    In-Space Propulsion Program Overview and Status (2002)
    In:  Other Sources
    Details
    Publication Date: 2019-08-14
    Description: NASA's In-Space Propulsion (ISP) Program is designed to develop advanced propulsion technologies that can enable or greatly enhance near and mid-term NASA science missions by significantly reducing cost, mass, and/or travel times. These technologies include: Electric Propulsion (Solar and Nuclear Electric) [note: The Nuclear Electric Propulsion work will be transferred to the NSI program in FY03]; Propellantless Propulsion (aerocapture, solar sails, plasma sails, and momentum exchange tethers); Advanced Chemical Propulsion. The ISP approach to identifying and prioritizing these most promising technologies is to use mission analysis and subsequent peer review. These technologies under consideration are mid-Technology Readiness Level (TRL) up to TRL-6 for incorporation into mission planning within three - five years of initiation. In addition, maximum use of open competition is encouraged to seek optimum solutions under ISP. Several NASA Research Announcements (NRAs) have been released asking industry, academia and other organizations to propose propulsion technologies designed to improve our ability to conduct scientific study of the outer planets and beyond. The ISP Program is managed by NASA HQ (Headquarters) and implemented by the Marshall Space Flight Center in Huntsville, Alabama.
    Keywords: Spacecraft Propulsion and Power
    Type: 28th International Electric Propulsion Conference (IEPC); Mar 17, 2003 - Mar 21, 2003; Toulouse; France
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    Prospects for Breakthrough Propulsion From Physics (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: "Space drives", "Warp drives", and "Wormholes:" these concepts may sound like science fiction, but they are being written about in reputable journals. To assess the implications of these emerging prospects for future spaceflight, NASA supported the Breakthrough Propulsion Physics Project from 1996 through 2002. This Project has three grand challenges: (1) Discover propulsion that eliminates the need for propellant; (2) Discover methods to achieve hyper-fast travel; and (3) Discover breakthrough methods to power spacecraft. Because these challenges are presumably far from fruition, and perhaps even impossible, a special emphasis is placed on selecting incremental and affordable research that addresses the critical issues behind these challenges. Of 16 incremental research tasks completed by the project and from other sponsors, about a third were found not to be viable, a quarter have clear opportunities for sequels, and the rest remain unresolved.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2004-213082 , E-14560 , 2004 Conference on Evolvable Hardware; Jun 24, 2004 - Jun 26, 2004; Seattle, WA; United States
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    Space Shuttle Main Engine Implications for the Abort-to-Orbit Off-the-Pad Study (2003)
    In:  Other Sources
    Details
    Publication Date: 2019-08-13
    Description: In 2001, the Space Shuttle Main Engine (SSME) project office was contacted by the Space Shuttle Ascent Guidance, Navigation, and Control group to provide the engine perspective for an Abort-to-Orbit (ATO) study. The purpose of the AT0 Off-the-Pad study was to determine the feasibility of eliminating the Return to Launch Site and Transatlantic abort modes by using a five-segment solid rocket booster and throttling the remaining SSMEs to a higher power level. This would enable all abort modes to be Abort-to-Orbit. The SSME project office at Marshall Space Flight Center collaborated with MSFC's Space Transportation Directorate and Rocketdyne Propulsion and Power to provide the AT0 Off-the-Pad study with the analysis required. Power levels at 109%, 1 11% and 1 13% of rated power level were studied as well as mixture ratio decreases down to 5.85. SSME was to evaluate and define the technical and programmatic impacts to certify the SSME to these abort power levels. The SSME systems analysis group performed a steady state analysis using the SSME power balance model to determine if there were any technical issues associated with higher power level, low mixture ratio operation. Based on each power level/mixture ratio combination, an engine certification plan was created and a preliminary probabilistic risk assessment was performed. The results showed favorable results for higher power 1evel/lower mixture ratio SSME operation. In nearly all performance and redline parameters, the traded engine operation was encompassed by nominal engine performance of a prior engine configuration.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF Conference; Dec 01, 2003; Colorado Springs, CO; United States
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    Support of Integrated Health Management (IHM) through Automated Analyses of Flowfield-Derived Spectrographic Data (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Flow-field analysis techniques under continuing development at NASA's Marshall Space Flight Center are the foundation for a new type of health monitoring instrumentation for propulsion systems and a vast range of other applications. Physics, spectroscopy, mechanics, optics, and cutting-edge computer sciences merge to make recent developments in such instrumentation possible. Issues encountered in adaptation of such a system to future space vehicles, or retrofit in existing hardware, are central to the work. This paper is an overview of the collaborative efforts results, current efforts, and future plans.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF Interagency Propulsion Committee; Dec 01, 2003 - Dec 05, 2003; Colorado Springs, CO; United States
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    Investigation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications (2003)
    In:  Other Sources
    Details
    Publication Date: 2019-08-13
    Description: This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Spac Flight Center (MSFC). Initial hot-fire tests in a small-scale rocket chamber at MSFC have demonstrated the dual pulse laser-induced spark (DPLIS) technique, which has an advantage over existing single-pulse laser ignition techniques in that it can be optimized in its laser pulse format to maximize the initial plasma volume, the plasm: lifetime, as well as the flame kernel growth rate. The distribution of the total laser energy into two separate pulse also lowers the peak power that would need to be sent through fiber optics to the combustion chamber, making the implementation of this technique more practical than other single-pulse techniques. A first generation prototype of an optic fiber-coupled laser ignition system will be tested a rocket chamber with RP-1/GOX and GH2/GOX propellants systems. Other relevant technology, such as optical windows, flight-qualified laser system etc. will be discussed in this paper.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF CS/APS/PSHS/MSS Meeting; Dec 03, 2003; Colorado Springs, CO; United States
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    Building a Propulsion Experiment Project Management Environment (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: What do you get when you cross rocket scientists with computer geeks? It is an interactive, distributed computing web of tools and services providing a more productive environment for propulsion research and development. The Rocket Engine Advancement Program 2 (REAP2) project involves researchers at several institutions collaborating on propulsion experiments and modeling. In an effort to facilitate these collaborations among researchers at different locations and with different specializations, researchers at the Information Technology and Systems Center,' University of Alabama in Huntsville, are creating a prototype web-based interactive information system in support of propulsion research. This system, to be based on experience gained in creating similar systems for NASA Earth science field experiment campaigns such as the Convection and Moisture Experiments (CAMEX), will assist in the planning and analysis of model and experiment results across REAP2 participants. The initial version of the Propulsion Experiment Project Management Environment (PExPM) consists of a controlled-access web portal facilitating the drafting and sharing of working documents and publications. Interactive tools for building and searching an annotated bibliography of publications related to REAP2 research topics have been created to help organize and maintain the results of literature searches. Also work is underway, with some initial prototypes in place, for interactive project management tools allowing project managers to schedule experiment activities, track status and report on results. This paper describes current successes, plans, and expected challenges for this project.
    Keywords: Spacecraft Propulsion and Power
    Type: 52nd JANNAF Joint Propulsion Meeting; May 10, 2004 - May 14, 2004; Las Vegas, NV; United States
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    Power Control and Monitoring Requirements for Thermal Vacuum/Thermal Balance Testing of the MAP Observatory (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The specific heater control requirements for the thermal vacuum and thermal balance testing of the Microwave Anisotropy Probe (MAP) Observatory at the Goddard Space Flight Center (GSFC) in Greenbelt, Maryland are described. The testing was conducted in the 10m wide x 18.3m high Space Environment Simulator (SES) Thermal Vacuum Facility. The MAP thermal testing required accurate quantification of spacecraft and fixture power levels while minimizing heater electrical emissions. The special requirements of the MAP test necessitated construction of five (5) new heater racks.
    Keywords: Spacecraft Propulsion and Power
    Type: 22nd IEST-NASA/ASTM/AIAA/CSA Space Simulation Conference; Oct 21, 2002 - Oct 24, 2002
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    Turbulent Mixing of Primary and Secondary Flow Streams in a Rocket-Based Combined Cycle Engine (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This viewgraph presentation gives an overview of the turbulent mixing of primary and secondary flow streams in a rocket-based combined cycle (RBCC) engine. A significant RBCC ejector mode database has been generated, detailing single and twin thruster configurations and global and local measurements. On-going analysis and correlation efforts include Marshall Space Flight Center computational fluid dynamics modeling and turbulent shear layer analysis. Potential follow-on activities include detailed measurements of air flow static pressure and velocity profiles, investigations into other thruster spacing configurations, performing a fundamental shear layer mixing study, and demonstrating single-shot Raman measurements.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF 38th Combustion Subcommittee Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
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    Design and Study of a LOX/GH2 Throttleable Swirl Injector for Rocket Applications (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A LOX/GH2 swirl injector was designed for a 10:1 propellant throttling range. To accomplish this, a dual LOX (liquid oxygen) manifold was used feeding a single common vortex chamber of the swirl element. Hot-fire experiments were conducting for rocket chamber pressures from 80 to 800 psia at a mixture ratio of nominally 6.0 using steady flow, single-point-per-firing cases as well as dynamic throttling conditions. Low frequency (mean) and high frequency (fluctuating) pressure transducer data, flow meter measurements, and Raman spectroscopy images for mixing information were obtained. The injector design, experimental setup, low frequency pressure data, and injector performance analysis will be presented. C efficiency was very high (approximately 100%) at the middle of the throttle-able range with somewhat lower performance at the high and low ends. From the analysis of discreet steady state operating conditions, injector pressure drop was slightly higher than predicted with an inviscid analysis, but otherwise agreed well across the design throttling range. Analysis of the dynamic throttling data indicates that the injector may experience transient conditions that effect pressure drop and performance when compared to steady state results.
    Keywords: Spacecraft Propulsion and Power
    Type: 2002 JANNAF 38th Combustion Subcommittee Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
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    Design and Performance of 40 cm Ion Optics (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A 40 cm ion thruster is being developed at the NASA Glenn Research Center to obtain input power and propellant throughput capabilities of 10 kW and 550 kg. respectively. The technical approach here is a continuation of the "derating" technique used for the NSTAR ion thruster. The 40 cm ion thruster presently utilizes the NSTAR ion optics aperture geometry to take advantage of the large database of lifetime and performance data already available. Dome-shaped grids were chosen for the design of the 40 cm ion optics because this design is naturally suited for large-area ion optics. Ion extraction capabilities and electron backstreaming limits for the 40 cm ion optics were estimated by utilizing NSTAR 30 cm ion optics data. A preliminary service life assessment showed that the propellant throughput goal of 550 kg of xenon may be possible with molybdenum 40 cm ion optics. One 40 cm ion optics' set has been successfully fabricated to date. Additional ion optics' sets are presently being fabricated. Preliminary performance tests were conducted on a laboratory model 40 cm ion thruster.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211275 , E-13074 , IEPC-01-090 , NAS 1.15:211275 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    X-33 XRS-2200 Linear Aerospike Engine Sea Level Plume Radiation (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Wide band plume radiation data were collected during ten sea level tests of a single XRS-2200 engine at the NASA Stennis Space Center in 1999 and 2000. The XRS-2200 is a liquid hydrogen/liquid oxygen fueled, gas generator cycle linear aerospike engine which develops 204,420 lbf thrust at sea level. Instrumentation consisted of six hemispherical radiometers and one narrow view radiometer. Test conditions varied from 100% to 57% power level (PL) and 6.0 to 4.5 oxidizer to fuel (O/F) ratio. Measured radiation rates generally increased with engine chamber pressure and mixture ratio. One hundred percent power level radiation data were compared to predictions made with the FDNS and GASRAD codes. Predicted levels ranged from 42% over to 7% under average test values.
    Keywords: Spacecraft Propulsion and Power
    Type: 26th JANNAF Exhaust Plume Technology Subcommittee Meeting; Nov 05, 2001 - Nov 09, 2001; San Antonio, TX; United States
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    Status of Real-Time Laser Based Ion Engine Diagnostics at NASA Glenn Research Center (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The development status of laser based erosion diagnostics for ion engines at the NASA Glenn Research Center is discussed. The diagnostics are being developed to enhance component life-prediction capabilities. A direct measurement of the erosion product density using laser induced fluorescence (LIF) is described. Erosion diagnostics based upon evaluation of the ion dynamics are also under development, and the basic approach is presented. The planned implementation of the diagnostics is discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211281 , E-13081 , IEPC-01-304 , NAS 1.15:211281 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Correlation of Hollow Cathode Assembly and Plasma Contactor Data from Ground Testing and In-Space Operation on the International Space Station (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Charge control on the International Space Station (ISS) is currently being provided by two plasma contactor units (PCUs). The plasma contactor includes a hollow cathode assembly (HCA), power processing unit and Xe gas feed system. The hollow cathode assemblies in use in the ISS plasma contactors were designed and fabricated at the NASA Glenn Research Center. Prequalification testing of development HCAs as well as acceptance testing of the flight HCAs is presented. Integration of the HCAs into the Boeing North America built PCU and acceptance testing of the PCU are summarized in this paper. Finally, data from the two on-orbit PCUs is presented.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211293 , NAS 1.15:211293 , E-13097 , IEPC-01-252 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Compact Plasma Accelerator for Micropropulsion Applications (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: There is a need for a low power, light-weight (compact), high specific impulse electric propulsion device to satisfy mission requirements for microsatellite (1 to 20 kg) class missions. Satisfying these requirements entails addressing the general problem of generating a sufficiently dense plasma within a relatively small volume and then accelerating it. In the work presented here, the feasibility of utilizing a magnetic cusp to generate a dense plasma over small length scales of order 1 mm is investigated. This approach could potentially mitigate scaling issues associated with conventional ion thruster plasma containment schemes. Plume and discharge characteristics were documented using a Faraday probe and a retarding potential analyzer.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211282 , E-13082 , NAS 1.15:211282 , IEPC-01-221 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Hall Thruster Plume Measurements On-Board the Russian Express Satellites (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The operation of North-South and East-West station-keeping Hall thruster propulsion systems on-board two Russian Express-A geosynchronous communication satellites were investigated through a collaborative effort with the manufacturer of the spacecraft. Over 435 firings of 16 different thrusters with a cumulative run time of over 550 hr were reported with no thruster failures. Momentum transfer due to plume impingement was evaluated based on reductions in the effective thrust of the SPT-100 thrusters and induced disturbance torques determined based on attitude control system data and range data. Hall thruster plasma plume effects on the transmission of C-band and Ku-band communication signals were shown to be negligible. On-orbit ion current density measurements were made and subsequently compared to predictions and ground test data. Ion energy, total pressure, and electric field strength measurements were also measured on-orbit. The effect of Hall thruster operation on solar array performance over several months was investigated. A subset of these data is presented.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211217 , NAS 1.15:211217 , E-13069 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Absolute Density Calibration Cell for Laser Induced Fluorescence Erosion Rate Measurements (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Flight qualification of ion thrusters typically requires testing on the order of 10,000 hours. Extensive knowledge of wear mechanisms and rates is necessary to establish design confidence prior to long duration tests. Consequently, real-time erosion rate measurements offer the potential both to reduce development costs and to enhance knowledge of the dependency of component wear on operating conditions. Several previous studies have used laser-induced fluorescence (LIF) to measure real-time, in situ erosion rates of ion thruster accelerator grids. Those studies provided only relative measurements of the erosion rate. In the present investigation, a molybdenum tube was resistively heated such that the evaporation rate yielded densities within the tube on the order of those expected from accelerator grid erosion. This work examines the suitability of the density cell as an absolute calibration source for LIF measurements, and the intrinsic error was evaluated.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211279 , E-13078 , NAS 1.15:211279 , IEPC-01-300 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Performance Evaluation of Titanium Ion Optics for the NASA 30 cm Ion Thruster (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The results of performance tests with titanium ion optics were presented and compared to those of molybdenum ion optics. Both titanium and molybdenum ion optics were initially operated until ion optics performance parameters achieved steady state values. Afterwards, performance characterizations were conducted. This permitted proper performance comparisons of titanium and molybdenum ion optics. Ion optics' performance A,as characterized over a broad thruster input power range of 0.5 to 3.0 kW. All performance parameters for titanium ion optics of achieved steady state values after processing 1200 gm of propellant. Molybdenum ion optics exhibited no burn-in. Impingement-limited total voltages for titanium ion optics where up to 55 V greater than those for molybdenum ion optics. Comparisons of electron backstreaming limits as a function of peak beam current density for molybdenum and titanium ion optics demonstrated that titanium ion optics operated with a higher electron backstreaming limit than molybdenum ion optics for a given peak beam current density. Screen grid ion transparencies for titanium ion optics were as much as 3.8 percent lower than those for molybdenum ion optics. Beam divergence half-angles that enclosed 95 percent of the total beam current for titanium ion optics were within 1 to 3 deg. of those for molybdenum ion optics. All beam divergence thrust correction factors for titanium ion optics were within 1 percent of those with molybdenum ion optics.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211220 , E-13073 , NAS 1.15:211220 , IEPC-01-092 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Combustion Instability Analysis and the Effects of Drop Size on Acoustic Driving Rocket Flow (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: High frequency combustion instability, the most destructive kind, is generally solved on a per engine basis. The instability often is the result of compounding acoustic oscillations, usually from the propellant combustion itself. To counteract the instability the chamber geometry can be changed and/or the method of propellant injection can be altered. This experiment will alter the chamber dimensions slightly; using a cylindrical shape of constant diameter and the length will be varied from six to twelve inches in three-inch increments. The main flowfield will be the products of a high OF hydrogen/oxygen flow. The liquid fuel will be injected into this flowfield using a modulated injector. It will allow for varied droplet size, feed rate, spray pattern, and location for the mixture within the chamber. The response will be deduced from the chamber pressure oscillations.
    Keywords: Spacecraft Propulsion and Power
    Type: 52nd JANNAF Joint Propulsion Meeting; May 10, 2004 - May 14, 2004; Las Vegas, NV; United States
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    Thermal Imaging of Aerospace Battery Cells (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The objective of the research covered by this viewgraph presentation is to understand the thermal characteristics of a cylindrical aerospace battery cell by studying the surface thermal properties. The batteries tested include a nickel hydrogen cell (NiH2), and a lithium-ion cell (Li-ion).
    Keywords: Spacecraft Propulsion and Power
    Type: 2004 NASA Aerospace Battery Workshop; Nov 16, 2004 - Nov 18, 2004; Huntsville, AL; United States
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    Hubble Space Telescope 2004 Battery Update (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Battery cell wear out mechanisms and signatures are examined and compared to orbital data from the six on-orbit Hubble Space Telescope (HST) batteries, and the Flight Spare Battery (FSB) Test Bed at Marshall Space Fiight Center (MSFC), which is instrumented with individual cell voltage monitoring.
    Keywords: Spacecraft Propulsion and Power
    Type: 2004 NASA Aerospace Battery Workshop; Nov 16, 2004 - Nov 18, 2004; Huntsville, AL; United States
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    Performance of AEA 80 Ah Battery under GEO Profiles (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This viewgraph presentation is divided into the following sections: 1) AEA Geosynchronous Earth Orbit (GEO) Life Testing; 2) AEA/Goddard Space Flight Center (GSFC) 20 Ah Battery; 3) AEA/GSFC 80 Ah Battery; 4) Solar Dynamic Observatory (SDO) Life Test; 5) Test Results; 6) Correlation; 7) Conclusions.
    Keywords: Spacecraft Propulsion and Power
    Type: 2004 NASA Aerospace Battery Workshop; Nov 16, 2004 - Nov 18, 2004; Huntsville, AL; United States
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    Tests of Mach's Principle With a Mechanical Oscillator (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: James F. Woodward has made a prediction, based on Sciama's formulation of Mach's Principle in the framework of general relativity, that in the presence of an energy flow the inertial mass of an object may undergo sizable variations, changing as the second time derivative of the energy. We describe an attempt to test for the predicted effect with a charging capacitor, using a technique that does not require an unbalanced force or any local violation of Newton s 3rd law of motion. We attempt to observe: (1) the gravitational effect of the varying mass and (2) the effect of the mass variation on a driven harmonic oscillator with the charging capacitor as the oscillating mass. We report on the predicted effect, the design and implementation of the measurement apparatus, and initial experience with the apparatus. At this time, however, we will not report on observations of the presence or absence of the Woodward effect.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2004-213310 , E-14770
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    Low Temperature Life-cycle Testing of a Lithium-ion Battery for Low-earth-orbiting Spacecraft (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A flight-qualified, lithium-ion (Li-ion) battery developed for the Mars Surveyor Program 2001 lander is undergoing life-testing at low temperature under a low-Earth-orbit (LEO) profile to assess its capability to provide long term energy storage for aerospace missions. NASA has embarked upon an ambitious course to return humans to the moon by 2015-2020 in preparation for robotic and human exploration of Mars and robotic exploration of the moons of outer planets. Li-ion batteries are excellent candidates to provide power and energy storage for multiple aspects of these missions due to their low specific energy, low energy density, and excellent low temperature performance. Laboratory testing of Li-ion technology is necessary in order to assess lifetime, characterize multi-cell battery-level performance under aerospace conditions, and to gauge safety aspects of the technology. Life-cycle testing provides an opportunity to examine battery-level performance and the dynamics of individual cells in the stack over the entire life of the battery. Data generated through this testing will be critical to establish confidence in the technology for its widespread use in manned and unmanned mission. This paper discusses the performance of the 28 volt, 25 ampere-hour battery through 6000 LEO cycles, which corresponds to one year on LEO orbit. Testing is being performed at 0 C and 40% depth-of-discharge. Individual cell behaviors and their effect on the performance of the battery are described. Capacity, impedance, energy efficiency and end-of-discharge voltage at 1000 cycle intervals are reported. Results from this life-testing will help contribute to the database on battery-level performance of aerospace Li-ion batteries and low temperature cycling under LEO conditions.
    Keywords: Spacecraft Propulsion and Power
    Type: E-14936 , NASA Aerospace Battery Workshop; Nov 16, 2004 - Nov 18, 2004; Huntsville, AL; United States
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    Improving the Flow (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: In early 1995, NASA s Glenn Research Center (then Lewis Research Center) formed an industry-government team with several jet engine companies to develop the National Combustion Code (NCC), which would help aerospace engineers solve complex aerodynamics and combustion problems in gas turbine, rocket, and hypersonic engines. The original development team consisted of Allison Engine Company (now Rolls-Royce Allison), CFD Research Corporation, GE Aircraft Engines, Pratt and Whitney, and NASA. After the baseline beta version was established in July 1998, the team focused its efforts on consolidation, streamlining, and integration, as well as enhancement, evaluation, validation, and application. These activities, mainly conducted at NASA Glenn, led to the completion of NCC version 1.0 in October 2000. NCC version 1.0 features high-fidelity representation of complex geometry, advanced models for two-phase turbulent combustion, and massively parallel computing. Researchers and engineers at Glenn have been using NCC to provide analysis and design support for various aerospace propulsion technology development projects. NASA transfers NCC technology to external customers using non- exclusive Space Act Agreements. Glenn researchers also communicate research and development results derived from NCC's further development through publications and special sessions at technical conferences.
    Keywords: Spacecraft Propulsion and Power
    Type: Spinoff; 75-76; NASA/NP-2004-10-374-HQ
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    Subscale Testing of Nozzle Ablative Materials in a Supersonic Particle Impingement Environment (2004)
    In:  Other Sources
    Details
    Publication Date: 2019-08-13
    Description: Recent efforts to evaluate materials to replace the current NARC rayon used in the nozzle ablative of the NASA's Reusable Solid Rocket Motor (SRM), several tests were developed to look at the performance of supersonic particle impact region of the aft exit cone. It was seen in early testing that some potential candidates did not perform as well as the current NARC based material and so the 24 inch Solid Rocket Motor (SRTM) exit cone design was extended and contoured to induce particle impingement in the aft end. The SRTM testing provided a larger testbed to evaluate the results seen in the Solid Fuel Supersonic Blast Tube. Testing was performed in each test bed on two variants of the final two candidate materials. The materials were a standard prewoven Lyocell, a postwoven Lyocell, an Enka rayon, and Enka rayon processed at a higher carbonization temperature. This paper presents the results of the four materials in the SSBT and the SRTM tests as compared against the NARC baseline. Erosion, char, and plylift results are discussed in detail.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF 14th NDES/23rd RNTS/36th S/MBS Joint Subcommittee Meeting; Mar 30, 2004 - Apr 01, 2004; New Orleans, LA; United States
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    Comparison of On-Orbit and Ground Based Hollow Cathode Operation (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The Plasma Contactor Unit (PCU) was developed by the Rocketdyne division of Boeing to control charging of the International Space Station (ISS). Each PCU contains a Hollow Cathode Assembly (HCA), which emits the charge control electrons. The HCAs were designed and fabricated at NASA's Glenn Research Center (GRC). GRC's HCA development program included manufacture of engineering, qualification, and flight model HCAs as well as wear tests and qualification tests. GRC is currently tracking the on-orbit data for the flight HCAs. This data will be discussed with comparison to operating parameters verified by ground based HCA tests. The flight HCAs continue to operate flawlessly. The first unit has accumulated more than 3650 hours of on-orbit operation and the second unit has accumulated over 5550 hours.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2003-212585 , E-14140 , NAS 1.26:212585 , AIAA Paper 2003-5145 , 39th Joint Propulsion Conference and Exhibit; Jul 20, 2003 - Jul 23, 2003; Huntsville, AL; United States
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    Hall Effect Thruster Interactions Data From the Russian Express-A2 and Express-A3 Satellites (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This 12-part report documents the data obtained from various sensor measurements taken aboard the Russian Express-A2 and Express-A3 spacecraft in Geosynchronous Earth Orbit (GEO). These GEO communications satellites, which were designed and built by NPO Prikladnoy Mekhaniki (NPO PM) of Zheleznogorsk, Russia, utilize Hall thruster propulsion systems for north-south and east-west stationkeeping and as of June 2002, were still operating at 80 E. and 11 W., respectively. Express-A2 was launched on March 12, 2000, while Express-A3 was launched on June 24, 2000. The diagnostic equipment from which these data were taken includes electric field strength sensors, ion current and energy sensors, and pressure sensors. The diagnostics and the Hall thruster propulsion systems are described in detail along with lists of tabular data from those diagnostics and propulsion system and other satellite systems. Space Power, Inc., now part of Pratt & Whitney's Chemical Systems Division, under contract NAS3 99151 to the NASA Glenn Research Center, obtained these data over several periods from March 12, 2000, through September 30, 2001. Each of the 12 individual reports describe, in detail, the propulsion systems as well as the diagnostic sensors utilized. Finally, parts 11 and 12 include the requirements to which NPO PM prepared and delivered these data.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2003-212005/PT10 , E-13691-10/PT10 , NAS 1.26:212005/PT10
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    Hall Effect Thruster Interactions Data from the Russian Express-A2 and Express-A3 Satellites (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This 12-part report documents the data obtained from various sensor measurements taken aboard the Russian Express-A2 and Express-A3 spacecraft in Geosynchronous Earth Orbit (GEO). These GEO communications satellites, which were designed and built by NPO Prikladnoy Mekhaniki (NPO PM) of Zheleznogorsk, Russia, utilize Hall thruster propulsion systems for north-south and east-west stationkeeping and as of June 2002, were still operating at 80deg E. and 11deg W., respectively. Express-A2 was launched on March 12, 2000, while Express-A3 was launched on June 24, 2000. The diagnostic equipment from which these data were taken includes electric field strength sensors, ion current and energy sensors, and pressure sensors. The diagnostics and the Hall thruster propulsion systems are described in detail along with lists of tabular data from those diagnostics and propulsion system and other satellite systems. Space Power, Inc., now part of Pratt & Whitney's Chemical Systems Division, under contract NAS3 99151 to the NASA Glenn Research Center, obtained these data over several periods from March 12, 2000, through September 30, 2001. Each of the 12 individual reports describe, in detail, the propulsion systems as well as the diagnostic sensors utilized. Finally, parts 11 and 12 include the requirements to which NPO PM prepared and delivered these data.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2003-212005/PT8 , E-13691-8/PT8 , NAS 1.26:212005/PT8
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    Performance and Stability Characteristics of a Uni-Element Swirl Injector for Oxygen-Rich Stage Combustion Cycles (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A uni-element liquid propellant combustion performance and instability study for liquid RP-1 and hot oxygen-rich pre-burner products was conducted, at a chamber pressure of about 1000 psi. using flush and recessed swirl injectors. High-frequency pressure transducer measurements were analyzed to yield the characteristic frequencies which were compared to expected frequencies of the chamber. Modes, which were discovered to be present within the main chamber included, the first longitudinal, detected at approximately 1950 Hz, and the second longitudinal mode at approximately 3800 Hz. An additional first longitudinal quarter wave mode was measured at a frequency of approximately 23000 Hz for the recessed swirl injector configuration. The characteristic instabilities resulting from these experiments were relatively weak averaging 0.2% to 0.3% of the chamber pressure.
    Keywords: Spacecraft Propulsion and Power
    Type: 52nd JANNAF Joint Propulsion Meeting; May 10, 2004 - May 14, 2004; Las Vegas, NV; United States
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    Outbrief - Long Life Rocket Engine Panel (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This white paper is an overview of the JANNAF Long Life Rocket Engine (LLRE) Panel results from the last several years of activity. The LLRE Panel has met over the last several years in order to develop an approach for the development of long life rocket engines. Membership for this panel was drawn from a diverse set of the groups currently working on rocket engines (Le. government labs, both large and small companies and university members). The LLRE Panel was formed in order to determine the best way to enable the design of rocket engine systems that have life capability greater than 500 cycles while meeting or exceeding current performance levels (Specific Impulse and Thrust/Weight) with a 1/1,OOO,OOO likelihood of vehicle loss due to rocket system failure. After several meetings and much independent work the panel reached a consensus opinion that the primary issues preventing LLRE are a lack of: physics based life prediction, combined loads prediction, understanding of material microphysics, cost effective system level testing. and the inclusion of fabrication process effects into physics based models. With the expected level of funding devoted to LLRE development, the panel recommended that fundamental research efforts focused on these five areas be emphasized.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF Meeting; May 19, 2004; Las Vegas, NV; United States
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    NASA's In-Space Propulsion Technology Program: Overview and Status (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    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 they are insufficient to meet needs for many ambitious science missions currently being considered. The In-Space Propulsion Technology Program s technology portfolio includes many advanced propulsion systems. From the next generation ion propulsion system operating in the 5 - 10 kW range, to advanced cryogenic 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, aeroassist, and aerocapture. This paper will provide an overview of both propellantless and propellant-based advanced propulsion technologies, and NASA s plans for advancing them as part of the $60M per year In-Space Propulsion Technology Program.
    Keywords: Spacecraft Propulsion and Power
    Type: 52nd Joint Army-Navy-NASA-Air Force Propulsion Meeting; May 10, 2004 - May 13, 2004; Las Vegas, NV; United States
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    Retention of Sputtered Molybdenum on Ion Engine Discharge Chamber Surfaces (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Grit-blasted anode surfaces are commonly used in ion engines to ensure adherence of sputtered coatings. Next generation ion engines will require higher power levels, longer operating times, and thus there will likely be thicker sputtered coatings on their anode surfaces than observed to date on 2.3 kW-class xenon ion engines. The thickness of coatings on the anode of a 10 kW, 40-centimeter diameter thruster, for example, may be 22 micrometers or more after extended operation. Grit-blasted wire mesh, titanium, and aluminum coupons were coated with molybdenum at accelerated rates to establish coating stability after the deposition process and after thermal cycling tests. These accelerated deposition rates are roughly three orders of magnitude more rapid than the rates at which the screen grid is sputtered in a 2.3 kW-class, 30-centimeter diameter ion engine. Using both RF and DC sputtering processes, the molybdenum coating thicknesses ranged from 8 to 130 micrometers, and deposition rates from 1.8 micrometers per hour to 5.1 micrometers per hour. In all cases, the molybdenum coatings were stable after the deposition process, and there was no evidence of spalling of the coatings after 20 cycles from about -60 to +320 C. The stable, 130 micrometer molybdenum coating on wire mesh is 26 times thicker than the thickest coating found on the anode of a 2.3 kW, xenon ion engine that was tested for 8200 hr. Additionally, this coating on wire mesh coupon is estimated to be a factor of greater than 4 thicker than one would expect to obtain on the anode of the next generation ion engine which may have xenon throughputs as high as 550 kg.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211319 , E-13130 , NAS 1.15:211319 , IEPC-01-086 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Measurement of Doubly Charged Ions in Ion Thruster Plumes (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The ratio of doubly to singly charged ions was measured in the plumes of a 30 cm and of a 40 cm ion thruster. The measured ratio was correlated with observed erosion rates and thruster operating conditions. The measured and calculated erosion rates paralleled variation in the j(sup ++)/j(sup +) ratio and indicated that the erosion was dominated by Xe III. Simple models of cathode potential surfaces which were developed in support of this work were in agreement with this conclusion and provided a predictive capability of the erosion given the ratio of doubly to singly charged ion currents.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211295 , E-13098 , NAS 1.15:211295 , IEPC-01-310 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Thermal Analysis of a Carbon Fiber Rope Barrier for Use in the Reusable Solid Rocket Motor Nozzle Joint-2 (2001)
    In:  Other Sources
    Details
    Publication Date: 2019-08-13
    Description: This study provides for development and verification of analysis methods used to assess performance of a carbon fiber rope (CFR) thermal barrier system that is currently being qualified for use in Reusable Solid Rocket Motor (RSRM) nozzle joint-2. Modeled geometry for flow calculations considers the joint to be vented with the porous CFR barriers placed in the "open' assembly gap. Model development is based on a 1-D volume filling approach where flow resistances (assembly gap and CFRs) are defined by serially connected internal flow and the porous media "Darcy" relationships. Combustion gas flow rates are computed using the volume filling code by assuming a lumped distribution total joint fill volume on a per linear circumferential inch basis. Gas compressibility, friction and heat transfer are included in the modeling. Gas-to-wall heat transfer is simulated by concurrent solution of the compressible flow equations and a large thermal 2-D finite element (FE) conduction grid. The derived numerical technique loosely couples the FE conduction matrix with the compressible gas flow equations, Free constants that appear in the governing equations are calibrated by parametric model comparison to hot fire subscale test results. The calibrated model is then used to make full-scale motor predictions using RSRM aft dome environments. Model results indicate that CFR thermal barrier systems will provide a thermally benign and controlled pressurization environment for the RSRM nozzle joint-2 primary seal activation.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF Interagency Propulsion Committee Meeting; Mar 27, 2001 - Mar 29, 2001; Cocoa Beach, FL; United States
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    JANNAF 25Th Airbreathing Propulsion Subcommittee, 37Th Combustion Subcommittee and 1St Modeling and Simulation Subcommittee Joint Meeting (2000)
    In:  Other Sources
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    Publication Date: 2019-08-13
    Description: Contents include the following: 1. Hyper-X program: Propulsion development and verification. 2. GTX program: Airbreathing launch vehicles. 3. Hypersonic technology development: Technology program overviews. Ramjet/scramjet research. 4. Hypersonic test methods: Test medium effects. 5. Advanced propulsion: RBCC engine design and performance assessments. Advanced and combined cycle engine technology.
    Keywords: Spacecraft Propulsion and Power
    Type: CPIA-Publ-703-Vol-1 , JANNAF 37th Combustion Subcommittee Meeting; Nov 13, 2000 - Nov 17, 2000; Monterey, CA; United States|JANNAF 25th Airbreathing Propulsion Subcommittee Meeting; Nov 13, 2000 - Nov 17, 2000; Monterey, CA; United States|JANNAF 1st Modeling and Simulation Subcommittee Meeting; Nov 13, 2000 - Nov 17, 2000; Monterey, CA; United States
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    THRSTER: A Three-Stream Ejector Ramjet Analysis and Design Tool (2000)
    In:  Other Sources
    Details
    Publication Date: 2019-08-13
    Description: An engineering tool for analyzing ejectors in rocket based combined cycle (RBCC) engines has been developed. A key technology for multi-cycle RBCC propulsion systems is the ejector which functions as the compression stage of the ejector ramjet cycle. The THRee STream Ejector Ramjet analysis tool was developed to analyze the complex aerothermodynamic and combustion processes that occur in this device. The formulated model consists of three quasi-one-dimensional streams, one each for the ejector primary flow, the secondary flow, and the mixed region. The model space marches through the mixer, combustor, and nozzle to evaluate the solution along the engine. In its present form, the model is intended for an analysis mode in which the diffusion rates of the primary and secondary into the mixed stream are stipulated. The model offers the ability to analyze the highly two-dimensional ejector flowfield while still benefits from the simplicity and speed of an engineering tool. To validate the developed code, wall static pressure measurements from the Penn-State and NASA-ART RBCC experiments were used to compare with the results generated by the code. The calculated solutions were generally found to have satisfactory agreement with the pressure measurements along the engines, although further modeling effort may be required when a strong shock train is formed at the rocket exhaust. The range of parameters in which the code would generate valid results are presented and discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF 1st Modeling and Simulation Subcommittee Meeting; 151-166; CPIA-Publ-702|Modeling and Simulation Subcommittee; Nov 13, 2000 - Nov 17, 2000; Monterey, CA; United States
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    Long Term Performance Retention Test Using High Power COTS NiCd and NiMH Cells (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This slide presentation reviews the tests and results for performance retention of high powered commercial off the shelf (COTS) NiCd, and NiMH cells. Electromechanical actuators for space flight requires short duration high power batteries. The concern is that NiCd battery designs demonstrate an unfavorable power degradation after long periods of inactivity. Cycling can recover some of the decay, but this reduces the readiness that these batteries must have. Two 5-cell SubC stick test batteries ere chosen using NiCd and NiMH were tested and then the differences for charge maintenance were compared.
    Keywords: Spacecraft Propulsion and Power
    Type: JSC-CN-8284 , 2003 NASA Aerospace Battery Workshop; Nov 16, 2003 - Nov 18, 2003; Huntsville, AL; United States
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    Application of the GRC Stirling Convertor System Dynamic Model (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The GRC Stirling Convertor System Dynamic Model (SDM) has been developed to simulate dynamic performance of power systems incorporating free-piston Stirling convertors. This paper discusses its use in evaluating system dynamics and other systems concerns. Detailed examples are provided showing the use of the model in evaluation of off-nominal operating conditions. The many degrees of freedom in both the mechanical and electrical domains inherent in the Stirling convertor and the nonlinear dynamics make simulation an attractive analysis tool in conjunction with classical analysis. Application of SDM in studying the relationship of the size of the resonant circuit quality factor (commonly referred to as Q) in the various resonant mechanical and electrical sub-systems is discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2004-213364 , AIAA Paper 2004-5672 , E-14832 , Second International Energy Conversion Engineering Conference; Aug 16, 2004 - Aug 19, 2004; Providence, RI; United States
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    Combination Solar Sail and Electrodynamic Tether Propulsion System (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A propulsion system for a spacecraft includes a solar sail system and an electrodynamic tether system is presented. The solar sail system is used to generate propulsion to propel the spacecraft through space using solar photons and the electrodynamic tether system is used to generate propulsion to steer the spacecraft into orbit and to perform orbital maneuvers around a planet using the planet's magnetic field. The electrodynamic tether system can also be used to generate power for the spacecraft using the planet's magnetic field.
    Keywords: Spacecraft Propulsion and Power
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    Comparison of On-orbit and Ground Based Hollow Cathode Operation (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The Plasma Contactor Units (PCUs) were developed at NASA Glenn Research Center (GRC) and Boeing for charge control on board the International Space Station (ISS). Since the first ignition of a PCU on 10/16/2000 over 3,900 hours of operation have been demonstrated on a single unit. In order to guarantee that the PCUs hollow cathode assemblies (HCAs), which emit the electrons used for charge control, would satisfy the life requirement of 18,000 hours, a ground based hollow cathode life test program was initiated at GRC. The life test program aimed at 27,000 hours of operation on a single unit to demonstrate the industry standard 1.5 times operational life requirement. As of this printing, over 18,000 hours of operation have been accumulated on a single hollow cathode. By comparing the data received from the on-orbit HCAs to the data obtained for the life test cathodes, a comparison may be drawn to determine if the on-orbit HCAs are operating normally, with a final goal of predicting lifetime. Based on the data taken thus far, it can be concluded that the on-orbit HCAs are operating within their design specifications.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2003-211817 , E-13510 , NAS 1.26:211817 , AIAA Paper 2002-4098 , 38th Joint Propulsion Conference and Exhibit; Jul 07, 2002 - Jul 10, 2002; Indianapolis, IN; United States
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    Hall Effect Thruster Interactions Data From the Russian Express-A2 and Express-A3 Satellites (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This 12-part report documents the data obtained from various sensor measurements taken aboard the Russian Express-A2 and Express-A3 spacecraft in Geosynchronous Earth Orbit (GEO). These GEO communications satellites, which were designed and built by NPO Prikladnoy Mekhaniki (NPO PM) of Zheleznogorsk, Russia, utilize Hall thruster propulsion systems for north-south and east-west stationkeeping and as of June 2002, were still operating at 80 E. and 11 W., respectively. Express-A2 was launched on March 12, 2000, while Express-A3 was launched on June 24, 2000. The diagnostic equipment from which these data were taken includes electric field strength sensors, ion current and energy sensors, and pressure sensors. The diagnostics and the Hall thruster propulsion systems are described in detail along with lists of tabular data from those diagnostics and propulsion system and other satellite systems. Space Power, Inc., now part of Pratt & Whitney's Chemical Systems Division, under contract NAS3 99151 to the NASA Glenn Research Center, obtained these data over several periods from March 12, 2000, through September 30, 2001. Each of the 12 individual reports describe, in detail, the propulsion systems as well as the diagnostic sensors utilized. Finally, parts 11 and 12 include the requirements to which NPO PM prepared and delivered these data.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2003-212005/PT3 , E-13691-3/PT3 , NAS 1.26:212005/PT3
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    Hall Effect Thruster Interactions Data From the Russian Express-A2 and Express-A3 Satellites (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This 12-part report documents the data obtained from various sensor measurements taken aboard the Russian Express-A2 and Express-A3 spacecraft in Geosynchronous Earth Orbit (GEO). These GEO communications satellites, which were designed and built by NPO Prikladnoy Mekhaniki (NPO PM) of Zheleznogorsk, Russia, utilize Hall thruster propulsion systems for north-south and east-west stationkeeping and as of June 2002, were still operating at 80deg E. and 11deg W., respectively. Express-A2 was launched on March 12, 2000, while Express-A3 was launched on June 24, 2000. The diagnostic equipment from which these data were taken includes electric field strength sensors, ion current and energy sensors, and pressure sensors. The diagnostics and the Hall thruster propulsion systems are described in detail along with lists of tabular data from those diagnostics and propulsion system and other satellite systems. Space Power, Inc., now part of Pratt & Whitney's Chemical Systems Division, under contract NAS3-99151 to the NASA Glenn Research Center, obtained these data over several periods from March 12, 2000, through September 30, 2001. Each of the 12 individual reports describe, in detail, the propulsion systems as well as the diagnostic sensors utilized. Finally, parts 11 and 12 include the requirements to which NPO PM prepared and delivered these data.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2003-212005/PT11 , E-13691--11/PT11 , NAS 1.26:212005/PT11
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    Wavelength-Agile Optical Sensor for Exhaust Plume and Cryogenic Fluid Interrogation (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Two optical sensors developed in UW-Madison labs were evaluated for their potential to characterize rocket engine exhaust plumes and liquid oxygen (LOX) fluid properties. The plume sensor is based on wavelength-agile absorption spectroscopy A device called a chirped white pulse emitter (CWPE) is used to generate the wavelength agile light, scanning, for example, 1340 - 1560 nm every microsecond. Properties of the gases in the rocket plume (for example temperature and water mole fraction) can be monitored using these wavelength scans. We have performed preliminary tests in static gas cells, a laboratory GOX/GH2 thrust chamber, and a solid-fuel hybrid thrust chamber, and these initial tests demonstrate the potential of the CWPE for monitoring rocket plumes. The LOX sensor uses an alternative to wavelength agile sensing: two independent, fixed-wavelength lasers are combined into a single fiber. One laser is absorbed by LOX and the other not: by monitoring the differential transmission the LOX concentration in cryogenic feed lines can be inferred. The sensor was successful in interrogating static LOX pools in laboratory tests. Even in ice- and bubble-laden cryogenic fluids, LOX concentrations were measured to better than 1% with a 3 microsec time constant.
    Keywords: Spacecraft Propulsion and Power
    Type: SSTI-2200-0002-FLUIDS , 52nd JANNAF Propulsion Meeting; May 10, 2004 - May 13, 2004; Las Vegas, NV; United States
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    MEMS-Based Spinning Nozzle (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A nozzle body and assembly for delivering atomized fuel to a combustion chamber. The nozzle body is rotatably mounted onto a substrate. One or more curvilinear fuel delivery channels are in flow communication with an internal fuel distribution cavity formed in the nozzle body. Passage of pressurized fuel through the nozzle body causes the nozzle body to rotate. Components of the nozzle assembly are formed of silicon carbide having surfaces etched by deep reactive ion etching utilizing MEMS (micro-electro-mechanical systems) technology. A fuel premix chamber is carried on the substrate in flow communication with a supply passage in the nozzle body.
    Keywords: Spacecraft Propulsion and Power
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    Plasma Propulsion Research at NASA Marshall Space Flight Center (2004)
    In:  Other Sources
    Details
    Publication Date: 2019-08-13
    Description: The Propulsion Research Center at NASA Marshall Space Flight Center is pursuing a range of research efforts aimed at identifying and developing new technologies for primary spacecraft propulsion. Efficient high-power electric propulsion (Ep) thrusters are a particular area of emphasis; these would enable the relatively rapid transit of large payloads about the solar system for unmanned or manned science and exploration. Such a mission would make heavy demands on the propulsion system, which may be required to run reliably for several years at a specific impulse approaching 10,OOO s with an efficiency of turning electrical power into jet power of at least 70%. The transit time to a destination scales approximately inversely with the cube root of the specific power, which is the ratio of jet power to power-plant mass. Consequently, reducing a trip time by half requires roughly an eight-fold increase in specific power. Given a renewed NASA commitment to space nuclear power, developing efficient EP thrusters with high jet power (〉 100 kW) would seem to provide the most direct means of significantly increasing the specific power and hence reducing trip times. In particular, electromagnetic devices, with their high inherent thrust densities, should be better suited to high power applications than thrusters which depend exclusively on electrostatic forces for propellant acceleration.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF Propulsion Meeting; May 10, 2004 - May 14, 2004; Las Vegas, NV; United States
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    Evaluation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications (2001)
    In:  Other Sources
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    Publication Date: 2019-08-13
    Description: To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer the system simplicity, but they also would enhance the combustion performance. The test results showed that the chamber performance was markedly high even at a low chamber length-to-diameter ratio (L/D). This incentive can be translated to a convenience in the thrust chamber packaging. Variations of the vortex chamber concepts have been introduced in the past few decades. These investigations include an ongoing work at Orbital Technologies Corporation (ORBITEC). By injecting the oxidizer tangentially at the chamber convergence and fuel axially at the chamber head end, Knuth et al. were able to keep the wall relatively cold. A recent investigation of the low L/D vortex chamber concept for gel propellants was conducted by Michaels. He used both triplet (two oxidizer and one fuel orifices) and unlike impinging schemes to inject propellants tangentially along the chamber wall. Michaels called the subject injection scheme as Impinging Stream Vortex Chamber (ISVC). His preliminary tests showed that high performance, with an Isp efficiency of 92%, can be obtained. MSFC and the U.S. Army are jointly investigating an application of the ISVC concept for the cryogenic oxygen/hydrocarbon propellant system. This vortex chamber concept is currently tested with gel propellants at AMCOM at Redstone Arsenal, Alabama. A version of this concept for the liquid oxygen (LOX)/hydrocarbon fuel (RPM) system has been derived from the one for the gel propellant.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF CS/APS/PSHS/MSS Joint Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
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    Laser Ignition Technology for Bi-Propellant Rocket Engine Applications (2001)
    In:  Other Sources
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    Publication Date: 2019-08-13
    Description: The fiber optically coupled laser ignition approach summarized is under consideration for use in igniting bi-propellant rocket thrust chambers. This laser ignition approach is based on a novel dual pulse format capable of effectively increasing laser generated plasma life times up to 1000 % over conventional laser ignition methods. In the dual-pulse format tinder consideration here an initial laser pulse is used to generate a small plasma kernel. A second laser pulse that effectively irradiates the plasma kernel follows this pulse. Energy transfer into the kernel is much more efficient because of its absorption characteristics thereby allowing the kernel to develop into a much more effective ignition source for subsequent combustion processes. In this research effort both single and dual-pulse formats were evaluated in a small testbed rocket thrust chamber. The rocket chamber was designed to evaluate several bipropellant combinations. Optical access to the chamber was provided through small sapphire windows. Test results from gaseous oxygen (GOx) and RP-1 propellants are presented here. Several variables were evaluated during the test program, including spark location, pulse timing, and relative pulse energy. These variables were evaluated in an effort to identify the conditions in which laser ignition of bi-propellants is feasible. Preliminary results and analysis indicate that this laser ignition approach may provide superior ignition performance relative to squib and torch igniters, while simultaneously eliminating some of the logistical issues associated with these systems. Further research focused on enhancing the system robustness, multiplexing, and window durability/cleaning and fiber optic enhancements is in progress.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF CS/APS/PSHS/MSS Joint Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
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    Design of a Modular 5-kW Power Processing Unit for the Next-Generation 40-cm Ion Engine (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: NASA Glenn Research Center is developing a 5/10-kW ion engine for a broad range of mission applications. Simultaneously, a 5-kW breadboard poster processing unit is being designed and fabricated. The design includes a beam supply consisting of four 1.1 kW power modules connected in parallel, equally sharing the output current. A novel phase-shifted/pulse-width-modulated dual full-bridge topology was chosen for its soft-switching characteristics. The proposed modular approach allows scalability to higher powers as well as the possibility of implementing an N+1 redundant beam supply. Efficiencies in excess of 96% were measured during testing of a breadboard beam power module. A specific mass of 3.0 kg/kW is expected for a flight PRO. This represents a 50% reduction from the state of the art NSTAR power processor.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211359 , E-13179 , NAS 1.15:211359 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    The Floating Potential Probe (FPP) for ISS: Operations and Initial Results (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: In this paper we report early results from the Floating Potential Probe (FPP) recently installed on the International Space Station (ISS). The data show that FPP properly measures the electrical potential of ISS structure with respect to the plasma it is flying through. FPP Langmuir probe data seem to give accurate measurements of the ambient plasma density, and are generally consistent with the IRI-90 model. FPP data are used to judge the performance of the ISS Plasma Contacting Units (PCUs), and to evaluate the extent of ISS charging in the absence of the PCUs.
    Keywords: Spacecraft Propulsion and Power
    Type: 7th Spacecraft Charging Technology Conference; Apr 23, 2001 - Apr 27, 2001; Noordwijk; Netherlands
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    The Rocket Engine Advancement Program 2 (REAP2) (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The Rocket Engine Advancement Program (REAP) 2 program is being conducted by a university propulsion consortium consisting of the University of Alabama in Huntsville, Penn State University, Purdue University, Tuskegee University and Auburn University. It has been created to bring their combined skills to bear on liquid rocket combustion stability and thrust chamber cooling. The research team involves well established and known researchers in the propulsion community. The cure team provides the knowledge base, research skills, and commitment to achieve an immediate and continuing impact on present and future propulsion issues. through integrated research teams composed of analysts, diagnosticians, and experimentalists working together in an integrated multi-disciplinary program. This paper provides an overview of the program, its objectives and technical approaches. Research on combustion instability and thrust chamber cooling are being accomplished
    Keywords: Spacecraft Propulsion and Power
    Type: 52nd JANNAF Joint Propulsion Meeting; May 10, 2004 - May 14, 2004; Las Vegas, NV; United States
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    Computational Simulation of Acoustic Modes in Rocket Combustors (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A combination of computational fluid dynamic analysis and analytical solutions is being used to characterize the dominant modes in liquid rocket engines in conjunction with laboratory experiments. The analytical solutions are based on simplified geometries and flow conditions and are used for careful validation of the numerical formulation. The validated computational model is then extended to realistic geometries and flow conditions to test the effects of various parameters on chamber modes, to guide and interpret companion laboratory experiments in simplified combustors, and to scale the measurements to engine operating conditions. In turn, the experiments are used to validate and improve the model. The present paper gives an overview of the numerical and analytical techniques along with comparisons illustrating the accuracy of the computations as a function of grid resolution. A representative parametric study of the effect of combustor mean flow Mach number and combustor aspect ratio on the chamber modes is then presented for both transverse and longitudinal modes. The results show that higher mean flow Mach numbers drive the modes to lower frequencies. Estimates of transverse wave mechanics in a high aspect ratio combustor are then contrasted with longitudinal modes in a long and narrow combustor to provide understanding of potential experimental simulations.
    Keywords: Spacecraft Propulsion and Power
    Type: 52nd JANNAF Joint Propulsion Meeting; May 10, 2004 - May 14, 2004; Las Vegas, NV; United States
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    Transient Simulation of the Integrated Powerhead Demonstrator (IPD) Rocket Engine (2003)
    In:  Other Sources
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    Publication Date: 2019-08-13
    Description: The IPD rocket engine is the product of a joint Air Force/NASA program to demonstrate the concept of a full-flow staged combustion power cycle. In this type of rocket engine cycle, both the entire fuel flow and the entire oxidizer flow are combusted in one of two pre-bumers and used to drive two turbopumps, thus utilizing the flow work of the total propellant flow. The basis of the IPD engine program relies on integrating newly developed and pre-existing hardware to demonstrate the component and material technologies to make this concept feasible, while simultaneously saving development time and costs. To provide insight to the project team and contractors during engine development and test phases, the engine system was modeled at MSFC using the Rocket Engine Transient Simulation (ROCETS) software to analyze system performance and determine component integration issues. The ROCETS software is used extensively a MSFC to perform steady-state power-balances and transient simulations of thermodynamic power and general fluid systems. The software is favored for its capability to solve large systems of non-linear equations, its librarie of fluid properties and flow devices, its flexibility to modify existing code to improve the physics-derived approximations of real fluid thermodynamic behavior, and the ability to add unique system constraints. The purpose of this paper is to present the methodology used to model the IPD engine system, detail the pitfalls encountered with the software, and explain the approximations made to more accurately represent engine component and fluid behavior. Engine system performance output from the model will be presented and explained in comparison with real fluid and component behavior.
    Keywords: Spacecraft Propulsion and Power
    Type: 1st Liquid Propulsion Meeting; May 10, 2004 - May 14, 2004; Las Vegas, NV; United States|JANNAF 52nd Propulsion Meeting; May 10, 2004 - May 14, 2004; Las Vegas, NV; United States
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    50th JANNAF Propulsion Meeting (2001)
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    Publication Date: 2019-08-13
    Description: This volume, the first of two volumes, is a collection of 29 unclassified/unlimited-distribution papers which were presented at the 50th Joint Army-Navy-NASA-Air Force (JANNAF) Propulsion Meeting, held 11-13 July 2001 at the Salt Lake City Marriott Hotel in Salt Lake City, Utah.
    Keywords: Spacecraft Propulsion and Power
    Type: CPIA-Publ-705-Vol-1 , 50th JANNAF Propulsion Meeting; Jul 11, 2001 - Jul 13, 2001; Salt Lake City, UT; United States
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    Translation Optics for 30 cm Ion Engine Thrust Vector Control (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Data were obtained from a 30 cm xenon ion thruster in which the accelerator grid was translated in the radial plane. The thruster was operated at three different throttle power levels, and the accelerator grid was incrementally translated in the X, Y, and azimuthal directions. Plume data was obtained downstream from the thruster using a Faraday probe mounted to a positioning system. Successive probe sweeps revealed variations in the plume direction. Thruster perveance, electron backstreaming limit, accelerator current, and plume deflection angle were taken at each power level, and for each accelerator grid position. Results showed that the thruster plume could easily be deflected up to six degrees without a prohibitive increase in accelerator impingement current. Results were similar in both X and Y direction.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211502 , NAS 1.15:211502 , E-13268 , IEPC-01-116 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Carbon-Based Ion Optics Development at NASA GRC (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: With recent success of the NSTAR ion thruster on Deep Space 1, there is continued interest in long term, high propellant throughput thrusters to perform energetic missions. This requires flight qualified thrusters that can operate for long periods at high beam density, without degradation in performance resulting from sputter induced grid erosion. Carbon-based materials have shown nearly an order of magnitude improvement in sputter erosion resistance over molybdenum. NASA Glenn Research Center (GRC) has been active over the past several years pursuing carbon-based grid development. In 1995, NASA GRC sponsored work performed by the Jet Propulsion Laboratory to fabricate carbon/carbon composite grids using a machined panel approach. In 1999, a contract was initiated with a commercial vendor to produce carbon/carbon composite grids using a chemical vapor infiltration process. In 2001, NASA GRC purchased pyrolytic carbon grids from a commercial vendor. More recently, a multi-year contract was initiated with North Carolina A&T to develop carbon/carbon composite grids using a resin injection process. The following paper gives a brief overview of these four programs.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211501 , NAS 1.15:211501 , IEPC-01-94 , E-13241 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    Characterization of Low-Frequency Combustion Stability of the Fastrac Engine (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. During mainstage, the thrust chamber exhibited no large-amplitude chamber pressure oscillations that could be identified as low-frequency combustion instability or 'chug'. However, during start-up and shutdown, the thrust chamber very briefly exhibited large-amplitude chamber pressure oscillations that were identified as chug. These instabilities during start-up and shutdown were regarded as benign due to their brevity. Linear models of the thrust chamber and the propellant feed systems were formulated for both the thrust chamber component tests and the flight engine tests. These linear models determined the frequency and decay rate of chamber pressure oscillations given the design and operating conditions of the thrust chamber and feed system. The frequency of chamber pressure oscillations determined from the model closely matched the frequency of low-amplitude, low-frequency chamber pressure oscillations exhibited in some of the later thrust chamber mainstage tests. The decay rate of the chamber pressure oscillations determined from the models indicated that these low-frequency oscillations were stable. Likewise, the decay rate, determined from the model of the flight engine tests indicated that the low-frequency chamber pressure oscillations would be stable.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF CS/APS/PSHS/MSS Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
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    Design And Testing of The Floating Potential Probe For ISS (2001)
    In:  CASI
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    Publication Date: 2019-08-13
    Description: Flight 4A was an especially critical mission for the International Space Station (ISS). For the first time, the high voltage solar arrays generated significant amounts of power and long predicted environmental interactions (high negative floating potential and concomitant dielectric charging) became serious concerns. Furthermore, the same flight saw the Plasma Contacting Unit (PCU) deployed and put into operation to mitigate and control these effects. The ISS program office has recognized the critical need to verify, by direct measurement, that ISS does not charge to unacceptable levels. A Floating Potential Probe (FPP) was therefore deployed on ISS to measure ISS floating potential relative to the surrounding plasma and to measure relevant plasma parameters. The primary objective of FPP is to verify that ISS floating potential does not exceed the specified level of 40 volts with respect to the ambient. Since it is expected that in normal operations the PCU will maintain ISS within this specification, it is equivalent to say that the objective of FPP is to monitor the functionality of the PCU. In this paper, we report on the design and testing of the ISS FPP. In a separate paper, the operations and results obtained so far by the FPP will be presented.
    Keywords: Spacecraft Propulsion and Power
    Type: 7th Spacecraft Charging Technology Conference; Noordwijk; Netherlands
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  • 90
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    Development of a Low-Cost, Subscale Test System to Evaluate Particle Impingement Erosion in Nozzle Ablative Materials (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This viewgraph presentation gives an overview on the development of a low-cost, subscale test system to evaluate particle impingement erosion in nozzle ablative materials. Details are given on the need for a new test bed, solid fuel torch components, solid fuel torch test, additional uses for the solid fuel torch, the development of a supersonic blast tube (SSBT), and particle impingement material discrimination.
    Keywords: Spacecraft Propulsion and Power
    Type: 50th JANNAF Propulsion Meeting; Jul 11, 2001; Salt Lake City, UT; United States
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  • 91
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    Improving the Total Impulse Capability of the NSTAR Ion Thruster With Thick-Accelerator-Grid Ion Optics (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The results of performance tests with thick-accelerator-grid (TAG) ion optics are presented. TAG ion optics utilize a 50 percent thicker accelerator grid to double ion optics' service life. NSTAR ion optics were also tested to provide a baseline performance for comparison. Impingement-limited total voltages for the TAG ion optics were only 0 to 15 V higher than those of the NSTAR ion optics. Electron backstreaming limits for the TAG ion optics were 3 to 9 V higher than those for the NSTAR optics due to the increased accelerator grid thickness for the TAG ion optics. Screen grid ion transparencies for the TAG ion optics were only about 2 percent lower than those for the NSTAR optics, reflecting the lower physical screen grid open area fraction of the TAG ion optics. Accelerator currents for the TAG ion optics were 19 to 43 percent greater than those for the NSTAR ion optics due, in part, to a sudden increase in accelerator current during TAG ion optics' performance tests for unknown reasons and to the lower-than-nominal accelerator aperture diameters. Beam divergence half-angles that enclosed 95 percent of the total beam current and beam divergence thrust correction factors for the TAG ion optics were within 2 degrees and 1 percent, respectively, of those for the NSTAR ion optics.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211276 , NAS 1.15:211276 , E-13075 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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  • 92
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    Life Model of Hollow Cathodes Using a Barium Calcium Aluminate Impregnated Tungsten Emitter (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Hollow cathodes with barium calcium aluminate impregnated tungsten emitters for thermionic emission are widely used in electric propulsion. These high current, low power cathodes are employed in ion thrusters, Hall thrusters, and on the International Space Station in plasma contactors. The requirements on hollow cathode life are growing more stringent with the increasing use of electric propulsion technology. The life limiting mechanism that determines the entitlement lifetime of a barium impregnated thermionic emission cathode is the evolution and transport of barium away from the emitter surface. A model is being developed to study the process of barium transport and loss from the emitter insert in hollow cathodes. The model accounts for the production of barium through analysis of the relevant impregnate chemistry. Transport of barium through the approximately static gas is also being treated. Finally, the effect of temperature gradients within the cathode are considered.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2001-211290 , E-13094 , NAS 1.26:211290 , IEPC-01-276 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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  • 93
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    A Review of Testing of Hollow Cathodes for the International Space Station Plasma Contactor (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: Since October 2000, two plasma contactors have been providing charge control on the International Space Station (ISS). At the heart of each of the two plasma contactors is a hollow cathode assembly (HCA) that produces the contacting xenon plasma. The HCA is the result of 9 years of design and testing at the NASA Glenn Research Center. This paper summarizes HCA testing that has been performed to date. As of this time, one cathode has demonstrated approximately 28,000 hr of lifetime during constant, high current use. Another cathode, HCA.014. has demonstrated 42,000 ignitions before cathode heater failure. In addition to these cathodes, four cathodes. HCA.006, HCA.003, HCA.010, and HCA.013 have undergone cyclic testing to simulate the variable current demand expected on the ISS. HCA.006 accumulated 8,000 hr of life test operation prior to being voluntarily stopped for analysis before the flight units were fabricated. HCA.010 has accumulated 15,876 hr of life testing, and 4,424 ignitions during ignition testing. HCA.003 and HCA.0 13 have accumulated 12,415 and 18,823 hr of life testing respectively.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211291 , E-13096 , NAS 1.15:211291 , IEPC-01-271 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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  • 94
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    Investigation of Keeper Erosion in the NSTAR Ion Thruster (2001)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The goal of the present investigation was to determine the cause for the difference in the observed discharge keeper erosion between the 8200 hr wear test of a NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) engineering model thruster and the ongoing extended life test (ELT) of the NSTAR flight spare thruster. During the ELT, the NSTAR flight spare ion thruster experienced unanticipated erosion of the discharge cathode keeper. Photographs of the discharge keeper show that the orifice has enlarged to slightly more than twice the original diameter. Several differences between the ELT and the 8200 hr wear test were initially identified to determine any effects which could lead to the erosion in the ELT. In order to identify the cause of the ELT erosion, emission spectra from an engineering model thruster were collected to assess the dependence of keeper erosion on operating conditions. Keeper ion current was measured to estimate wear. Additionally, post-test inspection of both a copper keeper-cap was conducted, and the results are presented. The analysis indicated that the bulk of the ion current was collected within 2-mm radially of the orifice. The estimated volumetric wear in the ELT was comparable to previous wear tests. Redistribution of the ion current on the discharge keeper was determined to be the most likely cause of the ELT erosion. The change in ion current distribution was hypothesized to caused by the modified magnetic field of the flight assemblies.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2001-211280 , E-13080 , IEPC-01-308 , NAS 1.15:211280 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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    SUPREM-DSMC: A New Scalable, Parallel, Reacting, Multidimensional Direct Simulation Monte Carlo Flow Code (2000)
    In:  Other Sources
    Details
    Publication Date: 2019-08-13
    Description: An AFRL/NRL team has recently been selected to develop a scalable, parallel, reacting, multidimensional (SUPREM) Direct Simulation Monte Carlo (DSMC) code for the DoD user community under the High Performance Computing Modernization Office (HPCMO) Common High Performance Computing Software Support Initiative (CHSSI). This paper will introduce the JANNAF Exhaust Plume community to this three-year development effort and present the overall goals, schedule, and current status of this new code.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF 25th Exhaust Plume Technology Subcommittee and 7th SPIRITS User Group Joint Meeting; 1; 53-60; CPIA-Publ-700-Vol-1|SPIRITS User Group; May 15, 2000 - May 18, 2000; Las Vegas, NV; United States|Exhaust Plume Technology Subcommittee; May 15, 2000 - May 18, 2000; Las Vegas, NV; United States
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  • 96
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    Rocket Engine Plume Diagnostics at Stennis Space Center (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The Stennis Space Center has been at the forefront of development and application of exhaust plume spectroscopy to rocket engine health monitoring since 1989. Various spectroscopic techniques, such as emission, absorption, FTIR, LIF, and CARS, have been considered for application at the engine test stands. By far the most successful technology h a been exhaust plume emission spectroscopy. In particular, its application to the Space Shuttle Main Engine (SSME) ground test health monitoring has been invaluable in various engine testing and development activities at SSC since 1989. On several occasions, plume diagnostic methods have successfully detected a problem with one or more components of an engine long before any other sensor indicated a problem. More often, they provide corroboration for a failure mode, if any occurred during an engine test. This paper gives a brief overview of our instrumentation and computational systems for rocket engine plume diagnostics at SSC. Some examples of successful application of exhaust plume spectroscopy (emission as well as absorption) to the SSME testing are presented. Our on-going plume diagnostics technology development projects and future requirements are discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: SE-2003-05-00029 , JANNAF 27th Exhaust Plume Technology Subcommittee; May 05, 2003 - May 09, 2003; Stennis Space Center, MS; United States
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    Hall Effect Thruster Interactions Data From the Russian Express-A2 and Express-A3 Satellites (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This 12-part report documents the data obtained from various sensor measurements taken aboard the Russian Express-A2 and Express-A3 spacecraft in Geosynchronous Earth Orbit (GEO). These GEO communications satellites, which were designed and built by NPO Prikladnoy Mekhaniki (NPO PM) of Zheleznogorsk, Russia, utilize Hall thruster propulsion systems for north-south and east-west stationkeeping and as of June 2002, were still operating at 80 E. and 11 W., respectively. Express-A2 was launched on March 12, 2000, while Express-A3 was launched on June 24, 2000. The diagnostic equipment from which these data were taken includes electric field strength sensors, ion current and energy sensors, and pressure sensors. The diagnostics and the Hall thruster propulsion systems are described in detail along with lists of tabular data from those diagnostics and propulsion system and other satellite systems. Space Power, Inc., now part of Pratt & Whitney's Chemical Systems Division, under contract NAS3 99151 to the NASA Glenn Research Center, obtained these data over several periods from March 12, 2000, through September 30, 2001. Each of the 12 individual reports describe, in detail, the propulsion systems as well as the diagnostic sensors utilized. Finally, parts 11 and 12 include the requirements to which NPO PM prepared and delivered these data.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/CR-2003-212005/Pt6 , NAS 1.26:212005/Pt6 , E-13691-6/Pt6
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  • 98
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    Reduced Toxicity Fuel Satellite Propulsion System Including Plasmatron (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply for consumption in an axial class thruster and an ACS class thruster. The system includes suitable valves and conduits for supplying the reduced toxicity propellant to the ACS decomposing element of an ACS thruster. The ACS decomposing element is operative to decompose the reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits for supplying the reduced toxicity propellant to an axial decomposing element of the axial thruster. The axial decomposing element is operative to decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits for supplying a second propellant to a combustion chamber of the axial thruster. whereby the hot gases and the second propellant auto-ignite and begin the combustion process for producing thrust.
    Keywords: Spacecraft Propulsion and Power
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  • 99
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    Status of the RBCC Direct-Connect Mixer Combustor Experiment (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The NASA Glenn Research Center is developing hydrogen based combined cycle propulsion technology for a single-stage-to-orbit launch vehicle application under a project called GTX. Rocket Based Combined Cycle (RBCC) propulsion systems incorporate one or more rocket engines into an airbreathing flow path to increase specific impulse as compared to an all rocket-powered vehicle. In support of this effort, an RBCC direct-connect test capability was established at the Engine Components Research Laboratory to investigate low speed, ejector ramjet, and initial ramjet operations and performance. The facility and test article enables the evaluation of two candidate low speed operating schemes; the simultaneous mixing and combustion (SMC) and independent ramjet stream (IRS). The SMC operating scheme is based on the fuel rich operations of the rocket where performance depends upon mixing between the rocket plume and airstream. In contrast, the IRS scheme fuels the airstream separately and uses the rocket plume to ignite the fuel-air mixture. This paper describes the test hardware and facility upgrades installed to support the RBCC tests. It also defines and discusses low speed technical challenges being addressed by the experiments. Finally, preliminary test results, including rocket risk mitigating tests, unfueled airflow tests, and the integrated system hot fire test will be presented.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211555 , NAS 1.15:211555 , E-13334 , Combustion, Airbreathing Propulsion, Propulsion Systems Hazards, and Modelling and Simulation Subcommittes Joint Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
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  • 100
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    Liquid-Metal-Fed Pulsed Electromagnetic Thrusters For In-Space Propulsion (2004)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: We describe three pulsed electromagnetic thruster concepts, which span four orders of magnitude in power processing capability (100 W to 〉100 kW), for in-space propulsion applications. The primary motivation for using a pulsed system is to is to enable high (instantaneous) power operation, which provides high acceleration efficiency, while using considerably less (continuous) power from the spacecraft power system. Unfortunately, conventional pulsed thrusters require failure-prone electrical switches and gas-puff valves. The series of thrusters described here directly address this problem, through the use of liquid metal propellant, by either eliminating both components or providing less taxing operational requirements, thus yielding a path toward both efficient and reliable pulsed electromagnetic thrusters. The emphasis of this paper is to conceptually describe each of the thruster concepts; however, initial test results with gallium propellant in one thruster geometry are presented. These tests reveal that a greater understanding of gallium material compatibility, contamination, and wetting behavior will be necessary before a completely functional thruster can be developed. Initial experimental results aimed at providing insight into these issues are presented.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF Conference; May 10, 2004 - May 13, 2004; Las Vegas, NV; United States
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