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 II
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.
20050207391 NASA Marshall Space Flight Center, Huntsville, AL, USA
NASA’s In-Space Propulsion Technology Program: A Step Toward Interstellar Exploration
Johnson, Les; James, Bonnie; Baggett, Randy; Montgomery, Sandy; April 19, 2005; 5 pp.; In English; 41st Symposiumon Realistic Near-Term Advanced Scientific Space Missions
4-6 Jul. 2005, Aosta, Italy; No Copyright; Avail: CASI; A01, Hardcopy
NASA’s In-Space Propulsion Technology Program is investing in technologies that have the potential to revolutionize the robotic exploration of deep space. For robotic exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs and, in some cases, enable missions previously considered impossible. Continued reliance on conventional chemical propulsion alone will not enable the robust exploration of deep space. The maximum theoretical efficiencies have almost been reached and are insufficient to meet needs for many ambitious science missions currently being considered. By developing the capability to support mid-term robotic mission needs, the program is laying the technological foundation for travel to nearby interstellar space. The In-Space Propulsion Technology Program s technology portfolio includes many advanced propulsion systems. From the next-generation ion propulsion systems operating in the 5-10 kW range, to solar sail propulsion, substantial advances in spacecraft propulsion performance are anticipated. Some of the most promising technologies for achieving these goals use the environment of space itself for energy and propulsion and are generically called ‘propellantless’ because they do not require onboard fuel to achieve thrust. Propellantless propulsion technologies include scientific innovations, such as solar sails, electrodynamic and momentum transfer tethers, and aerocapture. This paper will provide an overview of those propellantless and propellant-based advanced propulsion technologies that will most significantly advance our exploration of deep space. Author
Spacecraft Propulsion; Robotics; Interstellar Space; Deep Space; Life Cycle Costs; Propulsion System Configurations; Propulsion System Performance; Space Exploration
20050207420 NASA Marshall Space Flight Center, Huntsville, AL, USA
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Electromagnetic Pumps for Conductive-Propellant Feed Systems
Markusic, T. E.; Polzin, K. A.; January 2005; 1 pp.; In English; Joint Propulsion Conference, 11-13 Jul. 2005, Tucson, AZ, USA; No Copyright; Avail: CASI; A01, Hardcopy
There has been a recent, renewed interest in high-power electric thrusters for application in nuclear-electric propulsion systems. Two of the most promising thrusters utilize liquid metal propellants: the lithium-fed magnetoplasmadynamic thruster and the bismuth-fed Hall thruster. An important element of part of the maturation of these thrusters will be the development of compact, reliable conductive-propellant feed system components. In the present paper we provide design considerations and experimental calibration data for electromagnetic (EM) pumps. The role of an electromagnetic pump in a liquid metal feed system is to establish a pressure gradient between the propellant reservoir and the thruster - to establish the requisite mass flow rate. While EM pumps have previously been used to a limited extent in nuclear reactor cooling loops, they have never been implemented in electric propulsion (EP) systems. The potential benefit of using EM pumps for EP are reliability (no moving parts) and the ability to precisely meter the propellant flow rate. We have constructed and tested EM pumps that use gallium, lithium, and bismuth propellants. Design details, test results (pressure developed versus current), and material compatibility issues are reported. It is concluded that EM pumps are a viable technology for application in both laboratory and flight EP conductive-propellant feed systems. Author
Electromagnetic Pumps; Propellants; Feed Systems; Propulsion System Configurations; Propulsion System Performance; Electric Propulsion
20050207431 Aerojet-General Corp., Redmond, WA, USA, NASA Glenn Research Center, Cleveland, OH, USA
Direct Drive Hall Thruster System Development
Hoskins, W. Andrew; Homiak, Daniel; Cassady, R. Joseph; Kerslake, Tom; Peterson, Todd; Ferguson, Dale; Snyder, Dave; Mikellides, Ioannis; Jongeward, Gary; Schneider, Todd; [2003]; 16 pp.; In English; 2003 Joint Propulsion Conference, 21-23 Jul. 2003, Huntsville, AL, USA Contract(s)/Grant(s): NAS3-01100 Report No.(s): AIAA Paper 2003-4726; Copyright; Avail: CASI; A03, Hardcopy
The status of development of a Direct Drive Hall Thruster System is presented. 13 the first part. a study of the impacts to spacecraft systems and mass benefits of a direct-drive architecture is reviewed.
The study initially examines four cases of SPT-100 and BPT-4000 Hall thrusters used for north-south station keeping on an EXPRESS-like geosynchronous spacecraft and for primary propulsion for a Deep Space- 1 based science spacecraft. The study is also extended the impact of direct drive on orbit raising for higher power geosynchronous spacecraft and on other deep space missions as a function of power and delta velocity. The major system considerations for accommodating a direct drive Hall thruster are discussed, including array regulation, system grounding, distribution of power to the spacecraft bus, and interactions between current-voltage characteristics for the arrays and thrusters.
The mass benefit analysis shows that, for the initial cases, up to 42 kg of dry mass savings is attributable directly to changes in the propulsion hardware. When projected mass impacts of operating the arrays and the electric power system at 300V are included, up to 63 kg is saved for the four initial cases. Adoption of high voltage lithium ion battery technology is projected to further improve these savings. Orbit raising of higher powered geosynchronous spacecraft, is the mission for which direct drive provides the most benefit, allowing higher efficiency electric orbit raising to be accomplished in a limited period of time, as well as nearly eliminating significant power processing heat rejection mass. The total increase in useful payload to orbit ranges up to 278 kg for a 25 kWspacecraft, launched from an Atlas IIA. For deep space missions, direct drive is found to be most applicable to higher power missions with delta velocities up to several km/s, typical of several Discovery-class missions.
In the second part, the status of development of direct drive propulsion power electronics is presented. The core of this hardware is the heater-keeper-magnet supply being qualified for the BPT-4000 by Aerojet. A breadboard propulsion power unit is in fabrication and is scheduled for delivery late in 2003. Author
Hall Thrusters; Geosynchronous Orbits; Spacecraft Propulsion; Mechanical Drives; Deep Space 1 Mission; Payloads; Systems Engineering
20050207448 Aerojet-General Corp., Huntsville, AL, USA
High-Flux, High Performance H2O2 Catalyst Bed for ISTAR
Ponzo, J.; [2005]; 1 pp.; In English; 40th JANNAF Combustion Meeting, 13-17 Jun. 2005, Charleston, NC, USA Contract(s)/Grant(s): NAS8-02001; No Copyright; Avail: Other Sources; Abstract Only
On NASA’s ISTAR RBCC program packaging and performance requirements exceeded traditional H2O2 catalyst bed capabilities. Aerojet refined a high performance, monolithic 90% H202 catalyst bed previously developed and demonstrated. This approach to catalyst bed design and fabrication was an enabling technology to the ISTAR tri-fluid engine. The catalyst bed demonstrated 55 starts at throughputs greater than 0.60 lbm/s/sq in for a duration of over 900 seconds in a physical envelope approximately 114 of traditional designs. The catalyst bed uses photoetched plates of metal bonded into a single piece monolithic structure. The precise control of the geometry and complete mixing results in repeatable, quick starting, high performing catalyst bed. Three different beds were designed and tested, with the best performing bed used for tri-fluid engine testing. Author
Catalysts; Engine Tests; Fabrication; Packaging
20050207455 Boeing Co., Huntington Beach, CA, USA, NASA Marshall Space Flight Center, Huntsville, AL, USA
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The Importance of Detailed Component Simulations in the Feedsystem Development for a Two-Stage-to Orbit Reusable Launch Vehicle
Mazurkivich, Pete; Chandler, Frank; Grayson, Gary; April 18, 2005; 17 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NAS8-01099 Report No.(s): AIAA Paper 2005-4370; No Copyright; Avail: CASI; A03, Hardcopy
To meet the requirements for the 2nd Generation Reusable Launch Vehicle (RLV), a unique propulsion feed system concept was identified using crossfeed between the booster and orbiter stages that could reduce the Two-Stage-to-Orbit (TSTO) vehicle weight and development cost by approximately 25%. A Main Propulsion System (MPS) crossfeed water demonstration test program was configured to address all the activities required to reduce the risks for the MPS crossfeed system. A transient, one-dimensional system simulation was developed for the subscale crossfeed water flow tests. To ensure accurate representation of the crossfeed valve’s dynamics in the system model, a high-fidelity, three-dimensional, computational fluid-dynamics (CFD) model was employed. The results from the CFD model were used to specify the valve’s flow characteristics in the system simulation. This yielded a crossfeed system model that was anchored to the specific valve hardware and achieved good agreement with the measured test data. These results allowed the transient models to be correlated and validated and used for full scale mission predictions. The full scale model simulations indicate crossfeed is ‘ viable with the system pressure disturbances at the crossfeed transition being less than experienced by the propulsion system during engine start and shutdown transients. Author
Reusable Launch Vehicles; Feed Systems; Computational Fluid Dynamics; Booster Rocket Engines; Flow Characteristics; Simulation; Dimensional Analysis; Propulsion System Configurations
20050207459 NASA Marshall Space Flight Center, Huntsville, AL, USA
Test Results for a Non-toxic, Dual Thrust Reaction Control Engine
Robinson, Philip J.; Veith, Eric M.; Turpin, Alicia A.; [2005]; 17 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 11-13 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NAS8-01109 Report No.(s): AIAA Paper 2005-4457; Copyright; Avail: CASI; A03, Hardcopy
A non-toxic, dual thrust reaction control engine (RCE) was successfully tested over a broad range of operating conditions at the Aerojet Sacramento facility. The RCE utilized LOX/Ethanol propellants; and was tested in steady state and pulsing modes at 25-lbf thrust (vernier) and at 870-lbf thrust (primary). Steady state vernier tests vaned chamber pressure (Pc) from 0.78 to 5.96 psia, and mixture ratio (MR) from 0.73 to 1.82, while primary steady state tests vaned Pc from 103 to 179 psia and MR from 1.33 to 1.76. Pulsing tests explored EPW from 0.080 to 10 seconds and DC from 5 to 50 percent at both thrust levels. Vernier testing accumulated a total of 6,670 seconds of firing time, and 7,215 pulses, and primary testing accumulated a total of 2,060 seconds of firing time and 3,646 pulses. Author
Thrust Control; Steady State; Testing Time; Ethyl Alcohol; Liquid Oxygen
20050207460 NASA Marshall Space Flight Center, Huntsville, AL, USA
Verification of a Multiphysics Toolkit Against the Magnetized Target Fusion Concept
Thomas, Scott; Perrell, Eric; Liron, Caroline; Chiroux, Robert; Cassibry, Jason; Adams, Robert B.; July 10, 2005; 11 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA Report No.(s): AIAA Paper 2005-4 141; Copyright; Avail: CASI; A03, Hardcopy
In the spring of 2004 the Advanced Concepts team at MSFC embarked on an ambitious project to develop a suite of modeling routines that would interact with one another. The tools would each numerically model a portion of any advanced propulsion system. The tools were divided by physics categories, hence the name multiphysics toolset. Currently most of the anticipated modeling tools have been created and integrated. Results are given in this paper for both a quarter nozzle with chemically reacting flow and the interaction of two plasma jets representative of a Magnetized Target Fusion device. The results have not been calibrated against real data as of yet, but this paper demonstrates the current capability of the multiphysics tool and planned future enhancements. Author
Fusion; Plasma Jets; Propulsion; Jet Flow
20050207461 NASA Marshall Space Flight Center, Huntsville, AL, USA
Future Directions for Fusion Propulsion Research at NASA
Adams, Robert B.; Cassibry, Jason T.; July 10, 2005; 6 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA Report No.(s): AIAA Paper 2005-4140; Copyright; Avail: CASI; A02, Hardcopy
Fusion propulsion is inevitable if the human race remains dedicated to exploration of the solar system. There are fundamental reasons why fusion surpasses more traditional approaches to routine crewed missions to Mars, crewed missions to the outer planets, and deep space high speed robotic missions, assuming that reduced trip times, increased payloads, and higher available power are desired. A recent series of informal discussions were held among members from government, academia, and industry concerning fusion propulsion. We compiled a sufficient set of arguments for utilizing fusion in space..If the U.S. is to lead the effort and produce a working system in a reasonable amount of time, NASA must take the initiative, relying on, but not waiting for, DOE guidance. Arguments for fusion propulsion are presented, along with fusion enabled mission examples, fusion technology trade space, and a proposed outline for future efforts. Author
Fusion Propulsion; Space Missions; Mars Missions; Payloads
20050207462 L’Garde, Inc., USA
Vacuum Deployment and Testing of a 4-Quadrant Scalable Inflatable Solar Sail System
Lichodziejewski, David; Derbes, Billy; Galena, Daisy; Friese, Dave; June 1, 2005; 9 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 11-15 Jul. 2005, Tucson, AZ, USA Report No.(s): AIAA Paper 2005-3927; No Copyright; Avail: CASI; A02, Hardcopy
Solar sails reflect photons streaming from the sun and transfer momentum to the sail. The thrust, though small, is continuous and acts for the life of the mission without the need for propellant. Recent advances in materials and ultra-low mass gossamer structures have enabled a host of useful missions utilizing solar sail propulsion.
The team of L’Garde, Jet Propulsion Laboratories, Ball Aerospace, and Langley Research Center, under the direction of the NASA In-Space Propulsion office, has been developing a scalable solar sail configuration to address NASA s future space propulsion needs. The baseline design currently in development and testing was optimized around the 1 AU solar sentinel mission. Featuring inflatably deployed sub-T(sub g), rigidized beam components, the 10,000 sq m sail and support structure weighs only 47.5 kg, including margin, yielding an areal density of 4.8 g/sq m. Striped sail architecture, net/membrane sail design, and L’Garde’s conical boom deployment technique allows scalability without high mass penalties. This same structural concept can be scaled to meet and exceed the requirements of a number of other useful NASA missions.
This paper discusses the interim accomplishments of phase 3 of a 3-phase NASA program to advance the technology readiness level (TRL) of the solar sail system from 3 toward a technology readiness level of 6 in 2005. Under earlier phases of the program many test articles have been fabricated and tested successfully. Most notably an unprecedented 4-quadrant 10 m solar sail ground test article was fabricated, subjected to launch environment tests, and was successfully deployed under simulated space conditions at NASA Plum Brook s 30m vacuum facility.
Phase 2 of the program has seen much development and testing of this design validating assumptions, mass estimates, and predicted mission scalability. Under Phase 3 a much larger 20 m square test article including subscale vane has been fabricated and tested. A 20 m system ambient deployment has been successfully conducted after enduring Delta-2 launch environment testing.
The program will culminate in a vacuum deployment of a 20 m subscale test article at the NASA Glenn s Plum Brook 30 m vacuum test facility to bring the TRL level as close to 6 as possible in 1 g. This focused program will pave the way for a flight experiment of this highly efficient space propulsion technology. Author
Solar Sails; Technology Assessment; Propulsion; Aerospace Engineering; Launching
20050207466 Alabama Univ., Huntsville, AL, USA
Characterization of an 8-cm Diameter Ion Source System
Li, Zhongmin; Hawk, C. W.; Hawk, Clark W.; Buttweiler, Mark S.; Williams, John D.; Buchholtz, Brett; June 16, 2005; 14 pp.; In English; AIAA 41st Joint Propulsion Conference and Exhibit, 10-14 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NCC8-200 Report No.(s): AlAA Paper 2005-3882; No Copyright; Avail: CASI; A03, Hardcopy
Results of tests characterizing an 8-cm diameter ion source are presented. The tests were conducted in three separate vacuum test facilities at the University of Alabama-Huntsville, Colorado State University, and L3 Communications’ ETI division. Standard ion optics tests describing electron backstreaming and total-voltage-limited impingement current behavior as a function of beam current were used as guidelines for selecting operating conditions where more detailed ion beam measurements were performed. The ion beam was profiled using an in-vacuum actuating probe system to determine the total ion current density and the ion charge state distribution variation across the face of the ion source. Both current density and ExB probes were utilized. The ion current density data were used to obtain integrated beam current, beam flatness parameters, and general beam profile shapes. The ExB probe data were used to determine the ratio of doubly to singly charged ion current. The ion beam profile tests were performed at over six different operating points that spanned the expected operating range of the DAWN thrusters being developed at L3. The characterization tests described herein reveal that the 8-cm ion source is suitable for use in (a) validating plasma diagnostic equipment, (b) xenon ion sputtering and etching studies of spacecraft materials, (c) plasma physics research, and (d) the study of ion thruster optics at varying conditions. Author
Characterization; Ion Optics; Ion Sources; Test Facilities; Vacuum Tests
20050207495 Boeing Aerospace Co., Huntington Beach, CA, USA
Pressurization System Modeling for a Generic Bimese Two- Stage-to-Orbit Reusable Launch Vehicle
Mazurkivich, Pete; Chandler, Frank; Nguyen, Han; April 09, 2005; 8 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA; Original contains color illustrations Contract(s)/Grant(s): NAS8-01099 Report No.(s): AIAA Paper 2005-437 I; No Copyright; Avail: CASI; A02, Hardcopy
A pressurization system model was developed for a generic bimese Two-Stage-to-orbit Reusable Launch Vehicle using a cross-feed system and operating with densified propellants. The model was based on the pressurization system model for a crossfeed subscale water test article and was validated with test data obtained from the test article. The model consists of the liquid oxygen and liquid hydrogen pressurization models, each made up of two submodels, Booster and Orbiter tank pressurization models. The tanks are controlled within a 0.2-psi band and pressurized on the ground with ambient helium and autogenously in flight with gaseous oxygen and gaseous hydrogen. A 15-psi pressure difference is maintained between the Booster and Orbiter tanks to ensure crossfeed check valve closure before Booster separation. The analysis uses an ascent trajectory generated for a generic bimese vehicle and a tank configuration based on the Space Shuttle External Tank. It determines the flow rates required to pressurize the tanks on the ground and in flight, and demonstrates the model’s capability to analyze the pressurization system performance of a full-scale bimese vehicle with densified propellants. Author
Pressurizing; Reusable Launch Vehicles; Fuel Tank Pressurization; Liquid Hydrogen; Propellant Tanks; Booster Rocket Engines
Source: NASA.
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