SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 18 - SEPTEMBER 09, 2005
20 SPACECRAFT PROPULSION AND POWER - PART I
Includes main propulsion systems and components, e.g., rocket engines; and spacecraft auxiliary power sources.
For related information see also 07 Aircraft Propulsion and Power, 28 Propellants and Fuels, 15 Launch Vehicles and Launch Operations, and 44 Energy Production and Conversion.
20050204019 Boeing Co., USA
Rocket Engine Nozzle Side Load Transient Analysis Methodology: A Practical Approach
Shi, John J.; [2005]; 3 pp.; In English; AIAA Conference, 18 Apr. 2005, Austin, TX, USA Contract(s)/Grant(s): NAS8-01140; No Copyright; Avail: CASI; A01, Hardcopy
At the sea level, a phenomenon common with all rocket engines, especially for a highly over-expanded nozzle, during ignition and shutdown is that of flow separation as the plume fills and empties the nozzle, Since the flow will be separated randomly. it will generate side loads, i.e. non-axial forces. Since rocket engines are designed to produce axial thrust to power the vehicles, it is not desirable to be excited by non-axial input forcing functions, In the past, several engine failures were attributed to side loads. During the development stage, in order to design/size the rocket engine components and to reduce the risks, the local dynamic environments as well as dynamic interface loads have to be defined. The methodology developed here is the way to determine the peak loads and shock environments for new engine components. In the past it is not feasible to predict the shock environments, e.g. shock response spectra, from one engine to the other, because it is not scaleable. Therefore, the problem has been resolved and the shock environments can be defined in the early stage of new engine development. Additional information is included in the original extended abstract. Author (revised)
Rocket Nozzles; Sea Level; Low Altitude; Rocket Exhaust; Nozzle Flow; Exhaust Flow Simulation
20050205653 NASA Glenn Research Center, Cleveland, OH, USA, ZIN Technologies, Inc., Brook Park, OH, USA, Sest, Inc., OH, USA
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Advanced Controller for the Free-Piston Stirling Convertor
Gerber, Scott S.; Jamison, Mike; Roth, Mary Ellen; Regan, Timothy F.; [2004]; 1 pp.; In English; 2nd International Energy Conversion Engineering Conference (IECEC), 16-19 Aug. 2004, Providence, RI, USA; No Copyright; Avail: Other Sources; Abstract Only
The free-piston Stirling power convertor is being considered as an advanced power conversion technology to be used for future NASA deep space missions requiring long life radioisotope power systems. This technology has a conversion efficiency of over 25%, which is significantly higher than the efficiency of the Radioisotope Thermal-electric Generators (RTG) now in use. The NASA Glenn Research Center has long been recognized as a leader in Stirling technology and is responsible for the development of advanced technologies that are intended to significantly improve key characteristics of the Stirling convertor. The advanced technologies identified for development also consider the requirements of potential future missions and the new capabilities that have become available in the associated technical areas. One of the key areas identified for technology development is the engine controller. To support this activity, an advanced controller is being developed for the Stirling power convertor. This controller utilizes active power factor correction electronics and microcontroller-based controls. The object of this paper is to present an overview of the advanced controller concept with modeling, simulation and hardware test data. Author
Stirling Engines; Controllers; Control Systems Design
20050205838 NASA Marshall Space Flight Center, Huntsville, AL, USA
Free Re-boost Electrodynamic Tether on the International Space Station
Bonometti, Joseph A.; Sorenson, Kirk F.; Jansen, Ralph H.; Dankanich, John W.; Frame, Kyle L.; July 05, 2005; 7 pp.; In English; 41st AlAA Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA Report No.(s): AIAA Paper 2005-4545; No Copyright; Avail: CASI; A02, Hardcopy
The International Space Station (ISS) currently experiences significant orbital drag that requires constant make up propulsion or the Station will quickly reenter the Earth’s Atmosphere. The reboost propulsion is presently achieved through the firing of hydrazine rockets at the cost of considerable propellant mass. The problem will inevitably grow much worse as station components continue to be assembled, particularly when the full solar panel arrays are deployed. This paper discusses many long established themes on electrodynamic propulsion in the context of Exploration relevance, shows how to couple unique ISS electrical power system characteristics and suggests a way to tremendously impact ISS’s sustainability. Besides allowing launch mass and volume presently reserved for reboost propellant to be reallocated for science experiments and other critically needed supplies, there are a series of technology hardware demonstrations steps that can be accomplished on ISS, which are helpful to NASA s Exploration mission. The suggested ElectroDynamic (ED) tether and flywheel approach is distinctive in its use of free energy currently unusable, yet presently available from the existing solar array panels on ISS. The ideas presented are intended to maximize the utility of Station and radically increase orbital safety. Author
International Space Station; Tethering; Electrodynamics; Acceleration (Physics); Spacecraft Propulsion
20050205865 NASA Glenn Research Center, Cleveland, OH, USA
Flywheel Energy Storage System Designed for the International Space Station
Delventhal, Rex A.; Research and Technology 2001; March 2002; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
Following successful operation of a developmental flywheel energy storage system in fiscal year 2000, researchers at the NASA Glenn Research Center began developing a flight design of a flywheel system for the International Space Station (ISS). In such an application, a two-flywheel system can replace one of the nickel-hydrogen battery strings in the ISS power system. The development unit, sized at approximately one-eighth the size needed for ISS was run at 60,000 rpm. The design point for the flight unit is a larger composite flywheel, approximately 17 in. long and 13 in. in diameter, running at 53,000 rpm when fully charged. A single flywheel system stores 2.8 kW-hr of useable energy, enough to light a 100-W light bulb for over 24 hr. When housed in an ISS orbital replacement unit, the flywheel would provide energy storage with approximately 3 times the service life of the nickel-hydrogen battery currently in use. Derived from text
Flywheels; Service Life; Nickel Hydrogen Batteries
20050206333 NASA Marshall Space Flight Center, Huntsville, AL, USA
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Development of Ionic Liquid Monopropellants for In-Space Propulsion
Blevins, John A.; Drake, Gregory W.; Osborne, Robin J.; [2005]; 2 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA; No Copyright; Avail: CASI; A01, Hardcopy
A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines. These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine.
The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster. Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner, shown schematically in Figure 1 and similar in design to apparatuses used by other researchers to study solid and liquid monopropellants, will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows. Author
Monopropellants; Rocket Engines; Temperature Control; Spark Ignition; Nitrogen Tetroxide; High Temperature; Combustion Temperature; Burning Rate
20050206336 Gray Research, Inc., Huntsville, AL, USA
Direct Drive for Low Power Hall Thrusters
Dankanich, JohnW.; [2005]; 10 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA; No Copyright; Avail: CASI; A02, Hardcopy
Due to recent studies, NASA has initiated the development of a low power Hall thruster for discovery class missions. The potential advantages of a low power Hall thruster is primarily due to its high efficiency operation at low power and its lower complexity compared to ion engines. Direct drive is another method of reducing the complexity of a Hall thruster system while improving its efficiency. The technical challenges associated with this technology are reported. Additionally, the benefits of this technology are discussed based on parametric studies and mission analysis. Author
Hall Thrusters; Space Missions; Technology Utilization; Solar Electric Propulsion
20050206360 NASA Glenn Research Center, Cleveland, OH, USA
18th Space Photovoltaic Research and Technology Conference
Morton, Thomas L., Compiler; April 2005; 14 pp.; In English; 18th Space Photovoltaic Research and Technology Conference, 16-18 Sep. 2003, Cleveland, OH, USA; See also 20050206361 - 20050206396 Contract(s)/Grant(s): WBS 22-103-06-10 Report No.(s): NASA/CP-2005-213431; E-14965; No Copyright; Avail: CASI; A03, Hardcopy
The 18th Space Photovoltaic Research and Technology (SPRAT XVIII) Conference was held September 16 to 18, 2003, at the Ohio Aerospace Institute (OAI) in Brook Park, Ohio. The SPRAT conference, hosted by the Photovoltaic and Space Environments Branch of the NASA Glenn Research Center, brought together representatives of the space photovoltaic community from around the world to share the latest advances in space solar cell technology. This year s conference continued to build on many of the trends shown in SPRAT XVII-the continued advances of thin-film and multijunction solar cell technologies and the new issues required to qualify those types of cells for space applications. Author
Aerospace Engineering; Solar Cells; Thin Films; Aerospace Environments; Radiation Damage; Photovoltaic Cells
20050206415 ATK-Thiokol Propulsion, Brigham City, UT, USA
NARC Rayon Replacement Program for the RSRM Nozzle, Phase IV Qualification and Implementation Status
Haddock, M. Reed; Wendel, Gary M.; Cook, Roger V.; [2005]; 10 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA; Original contains color illustrations Contract(s)/Grant(s): NAS8-97238; Copyright; Avail: CASI; A02, Hardcopy
The Space Shuttle NARC Rayon Replacement Program has down-selected Enka rayon as a replacement for the obsolete NARC rayon in the nozzle carbon cloth phenolic (CCP) ablative insulators. Full qualification testing of the Enka rayon-based carbon cloth phenolic is underway, including processing, thmal/structural properties, and hot-fire subscale tests. Required thermal-structural capabilities, together with confidence in erosio/char performance in simulated and subscale hot fire tests such as Wright-Patterson Air Force Base Laser Hardened Materials Evaluation Laboratory testing, NASA-MSFC 24-inch motor tests, NASA-MSFC Solid Fuel Torch - Super Sonic Blast Tube, NASA-MSFC Plasma Torch Test Bed, ATK Thiokol Forty Pound Charge and NASA-MSFC MNASA justified the testing of the new Enka-rayon candidate on full-scale static test motors. The first RSRM full-scale static test motor nozzle, fabricated using the new Enka rayon-based CCP, was successfully demonstrated in June 2004. Two additional static test motors are planned with the new Enka rayon in the next two years along with additional A-basis property characterization. Process variation or ‘corner-of-the-box’ testing together with cured and uncured aging studies are also planned as some of the pre-flight implementation activities with 5-year cured aging studies over-lapping flight hardware fabrication. Author
Carbon Fibers; Rayon; Full Scale Tests; Performance Tests; Laser Materials
20050207310 NASA Glenn Research Center, Cleveland, OH, USA
Solar Electric Propulsion Technologies Being Designed for Orbit Transfer Vehicle Applications
Sarver-Verhey, Timothy R.; Hoffman, David J.; Kerslake, Thomas W.; Oleson, Steven R.; Falck, Robert D.; Research and Technology 2001; March 2002; 4 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
There is increasing interest in employing Solar Electric Propulsion (SEP) for new missions requiring transfer from low Earth orbit to the Earth-Moon Lagrange point, L1. Mission architecture plans place the Gateway Habitat at L1 in the 2011 to 2016 timeframe. The Gateway Habitat is envisioned to be used for Lunar exploration, space telescopes, and planetary mission staging. In these scenarios, an SEP stage, or ‘tug,’ is used to transport payloads to L1--such as the habitat module, lunar excursion and return vehicles, and chemical propellant for return crew trips. SEP tugs are attractive because they are able to efficiently transport large (less than 10,000 kg) payloads while minimizing propellant requirements. To meet the needs of these missions, a preliminary conceptual design for a general-purpose SEP tug was developed that incorporates several of the advanced space power and in-space propulsion technologies (such as high-power gridded ion and Hall thrusters, high-performance thin-film photovoltaics, lithium-ion batteries, and advanced high-voltage power processing) being developed at the NASA Glenn Research Center. A spreadsheet-based vehicle system model was developed for component sizing and is currently being used for mission planning. This model incorporates a low-thrust orbit transfer algorithm to make preliminary determinations of transfer times and propellant requirements. Results from this combined tug mass estimation and orbit transfer model will be used in a higher fidelity trajectory model to refine the analysis. Author
Solar Electric Propulsion; Hall Thrusters; High Voltages; Ion Engines; Photovoltaic Conversion
20050207337 NASA Kennedy Space Center, Cocoa Beach, FL, USA
STS-114: Discovery Propulsion System Modification Briefing
July 10, 2005; In English; 1 hr., 37 sec. playing time, in color, with sound; No Copyright; Avail: CASI; V04, Videotape-VHS; B04, Videotape-Beta
A briefing on the propulsion system modification of the STS-114 Discovery is presented. June Malone, NASA Public Affairs, introduces the panel who consists of: Sandy Coleman, External Tank Project Manager, Neil Otte, External Tank Chief Engineer, and Tom Williams, Solid Rocket Booster, Deputy Project Manager. Neil Otte presents charts on new requirements for foam debris reduction on the external tank. He also presents charts describing the Forward Bipod Redesign, LO2 Feedline Bellows Location, LH2 Intertank Flange Location, and In-Flight Imagery. Tom Williams presents charts describing Solid Rocket Booster Activities and Return to Flight efforts. CASI
Space Transportation System; NASA Space Programs; Discovery (Orbiter); Spacecraft Propulsion; Systems Engineering
20050207370 NASA Marshall Space Flight Center, Huntsville, AL, USA
Chemical Propulsion Technology Challenges for Exploration
Leopard, Larry; May 9, 2005; 30 pp.; In English; Topics in Engineering (TE20), NASA Training Workshop Technologies for Space Exploration, 17-18 May 2005, Hampton, VA, USA; No Copyright; Avail: CASI; A03, Hardcopy
Contents include the following: Introduction. Assumptions and scope. Reference mission (National). Propulsion elements amd opportunities for technology insertion. Issues. Summary and conclusion. CASI
Chemical Propulsion; Technology Utilization
20050207374 NASA Glenn Research Center, Cleveland, OH, USA
Thruster Plume Plasma Diagnostics: A Ground Chamber Experiment for a 2-Kilowatt Arcjet
Galofaro, Joel T.; Vayner, Boris V.; Hillard, G. Barry; Chornak, Michael T.; July 2005; 11 pp.; In English; Sixth Frontiers in Low-Temperature Plasma Diagnostics Workshop, 17-21 Apr. 2005, Les Houches, France Contract(s)/Grant(s): WBS 22-612-50-81-07 Report No.(s): NASA/TM-2005-213837; E-15207; No Copyright; Avail: CASI; A03, Hardcopy
Although detailed near field (0 to 3 cm) information regarding the exhaust plume of a two kilowatt arc jet is available (refs. 1 to 6), there is virtually little or no information (outside of theoretical extrapolations) available concerning the far field (2.6 to 6.1 m). Furthermore, real information about the plasma at distances between (3 to 6 m) is of critical importance to high technology satellite companies in understanding the effect of arc jet plume exhausts on space based power systems. It is therefore of utmost importance that one understands the exact nature of the interaction between the arc jet plume, the spacecraft power system and the surrounding electrical plasma environment.
A good first step in understanding the nature of the interactions lies in making the needed plume parameter measurements in the far field. All diagnostic measurements are performed inside a large vacuum system (12 m diameter by 18 m high) with a full scale arc jet and solar array panel in the required flight configuration geometry. Thus, necessary information regarding the plume plasma parameters in the far field is obtained.
Measurements of the floating potential, the plasma potential, the electron temperature, number density, density distribution, debye length, and plasma frequency are obtained at various locations about the array (at vertical distances from the arc jet nozzle: 2.6, 2.7, 2.8, 3.2, 3.6, 4.0, 4.9, 5.0, 5.4, 5.75, and 6.14 m). Plasma diagnostic parameters are measured for both the floating and grounded configurations of the arc jet anode and array. Spectroscopic optical measurements are then acquired in close proximity to the nozzle, and contamination measurements are made in the vicinity of the array utilizing a mass spectrometer and two Quartz Crystal Microbalances (QCM’s). Author
Jet Exhaust; Plumes; Spacecraft Power Supplies; Electron Energy; Optical Measurement; Plasma Frequencies
Source: NASA.
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