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Sample records for ion propulsion engine

  1. A segmented ion engine design for solar electric propulsion systems

    NASA Technical Reports Server (NTRS)

    Brophy, John R.

    1992-01-01

    A new ion engine design, called a segmented ion engine, is described which is capable of reducing the required ion source life time for small body rendezvous missions from 18,000 h to about 8,000 h. The use of SAND ion optics for the engine accelerator system makes it possible to substantially reduce the cost of demonstrating the required engine endurance. It is concluded that a flight test of a 5-kW xenon ion propulsion system on the ELITE spacecraft would enormously reduce the cost and risk of using ion propulsion on a planetary vehicle by addressing systems level issues associated with flying a spacecraft radically different from conventional planetary vehicles.

  2. A segmented ion engine design for solar electric propulsion systems

    NASA Technical Reports Server (NTRS)

    Brophy, John R.

    1992-01-01

    A new ion engine design, called a segmented ion engine, is described which is capable of reducing the required ion source life time for small body rendezvous missions from 18,000 h to about 8,000 h. The use of SAND ion optics for the engine accelerator system makes it possible to substantially reduce the cost of demonstrating the required engine endurance. It is concluded that a flight test of a 5-kW xenon ion propulsion system on the ELITE spacecraft would enormously reduce the cost and risk of using ion propulsion on a planetary vehicle by addressing systems level issues associated with flying a spacecraft radically different from conventional planetary vehicles.

  3. Ion engine auxiliary propulsion applications and integration study

    NASA Technical Reports Server (NTRS)

    Zafran, S. (Editor)

    1977-01-01

    The benefits derived from application of the 8-cm mercury electron bombardment ion thruster were assessed. Two specific spacecraft missions were studied. A thruster was tested to provide additional needed information on its efflux characteristics and interactive effects. A Users Manual was then prepared describing how to integrate the thruster for auxiliary propulsion on geosynchronous satellites. By incorporating ion engines on an advanced communications mission, the weight available for added payload increases by about 82 kg (181 lb) for a 100 kg (2200 lb) satellite which otherwise uses electrothermal hydrazine. Ion engines can be integrated into a high performance propulsion module that is compatible with the multimission modular spacecraft and can be used for both geosynchronous and low earth orbit applications. The low disturbance torques introduced by the ion engines permit accurate spacecraft pointing with the payload in operation during thrusting periods. The feasibility of using the thruster's neutralizer assembly for neutralization of differentially charged spacecraft surfaces at geosynchronous altitude was demonstrated during the testing program.

  4. Ion propulsion engine installed on Deep Space 1 at CCAS

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), maneuver the ion propulsion engine into place before installation on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight- tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October.

  5. Ion propulsion engine installed on Deep Space 1 at CCAS

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), make adjustments while installing the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight- tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October.

  6. Ion propulsion engine installed on Deep Space 1 at CCAS

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), attach a strap during installation of the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October.

  7. Ion propulsion engine installed on Deep Space 1 at CCAS

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), install an ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October.

  8. Ion propulsion engine installed on Deep Space 1 at CCAS

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Workers in the Defense Satellite Communications Systems Processing Facility (DPF) at Cape Canaveral Air Station (CCAS) make adjustments while installing the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight- tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched Oct. 25 aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS.

  9. Ion propulsion engine installed on Deep Space 1 at CCAS

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Workers in the Defense Satellite Communications Systems Processing Facility (DPF) at Cape Canaveral Air Station (CCAS) finish installing the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched Oct. 25 aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS.

  10. Engineer Examines Cluster of Ion Engines in the Electric Propulsion Laboratory

    NASA Image and Video Library

    1963-01-21

    New staff member Paul Margosian inspects a cluster of ion engines in the Electric Propulsion Laboratory’s 25-foot diameter vacuum tank at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis researchers had been studying different methods of electric rocket propulsion since the mid-1950s. Harold Kaufman created the first successful engine, the electron bombardment ion engine, in the early 1960s. These engines used electric power to create and accelerate small particles of propellant material to high exhaust velocities. Electric engines have a very small thrust, and but can operate for long periods of time. The ion engines are often clustered together to provide higher levels of thrust. The Electric Propulsion Laboratory contained two large vacuum tanks capable of simulating the space environment. The tanks were designed especially for testing ion and plasma thrusters and spacecraft. The larger 25-foot diameter tank was intended for testing electric thrusters with condensable propellants. The tank’s test compartment, seen here, was 10 feet in diameter. Margosian joined Lewis in late 1962 during a major NASA hiring phase. The Agency reorganized in 1961 and began expanding its ranks through a massive recruiting effort. Lewis personnel increased from approximately 2,700 in 1961 to over 4,800 in 1966. Margosian, who worked with Bill Kerslake in the Electromagnetic Propulsion Division’s Propulsion Systems Section, wrote eight technical reports on mercury and electron bombardment thrusters, thermoelectrostatic generators, and a high voltage insulator.

  11. Three-grid accelerator system for an ion propulsion engine

    NASA Technical Reports Server (NTRS)

    Brophy, John R. (Inventor)

    1994-01-01

    An apparatus is presented for an ion engine comprising a three-grid accelerator system with the decelerator grid biased negative of the beam plasma. This arrangement substantially reduces the charge-exchange ion current reaching the accelerator grid at high tank pressures, which minimizes erosion of the accelerator grid due to charge exchange ion sputtering, known to be the major accelerator grid wear mechanism. An improved method for life testing ion engines is also provided using the disclosed apparatus. In addition, the invention can also be applied in materials processing.

  12. Computer controlled operation of a two-engine xenon ion propulsion system

    NASA Technical Reports Server (NTRS)

    Brophy, John R.

    1987-01-01

    The development and testing of a computer control system for a two-engine xenon ion propulsion module is described. The computer system controls all aspects of the propulsion module operation including: start-up, steady-state operation, throttling and shutdown of the engines; start-up, operation and shutdown of the central neutralizer subsystem; control of the gimbal system for each engine; and operation of the valves in the propellant storage and distribution system. The most important engine control algorithms are described in detail. These control algorithms provide flexibility in the operation and throttling of ion engines which has never before been possible. This flexibility is made possible in large part through the use of flow controllers which maintain the total flow rate of propellant into the engine at the proper level. Data demonstrating the throttle capabilities of the engine and control system are presented.

  13. Computer controlled operation of a two-engine xenon ion propulsion system

    NASA Technical Reports Server (NTRS)

    Brophy, John R.

    1987-01-01

    The development and testing of a computer control system for a two-engine xenon ion propulsion module is described. The computer system controls all aspects of the propulsion module operation including: start-up, steady-state operation, throttling and shutdown of the engines; start-up, operation and shutdown of the central neutralizer subsystem; control of the gimbal system for each engine; and operation of the valves in the propellant storage and distribution system. The most important engine control algorithms are described in detail. These control algorithms provide flexibility in the operation and throttling of ion engines which has never before been possible. This flexibility is made possible in large part through the use of flow controllers which maintain the total flow rate of propellant into the engine at the proper level. Data demonstrating the throttle capabilities of the engine and control system are presented.

  14. Electric propulsion. [pulsed plasma thruster and electron bombardment ion engine for MSAT attitude control and stationkeeping

    NASA Technical Reports Server (NTRS)

    1982-01-01

    An alternative propulsion subsystem for MSAT is presented which has a potential of reducing the satellite weight by more than 15%. The characteristics of pulsed plasma and ion engines are described and used to estimate of the mass of the propellant and thrusters for attitude control and stationkeeping functions for MSAT. Preliminary estimates indicate that the electric propulsion systems could also replace the large momentum wheels necessary to counteract the solar pressure; however, the fine pointing wheels would be retained. Estimates also show that either electric propulsion system can save approximately 18% to 20% of the initial 4,000 kg mass. The issues that require further experimentation are mentioned.

  15. PROPELLANTS FOR ELECTRICAL PROPULSION ENGINES OF THE CONTACT OR BOMBARDMENT ION TYPE

    DTIC Science & Technology

    ALKALI METALS, *ANTHRACENES, *COLLOIDS, *ELECTRIC PROPULSION, * FERROCENES , *ION BEAMS, *IONIZATION, *MOLECULES, *PHENANTHRENES, *PLASMA JETS... PROPELLANTS , ACCELERATION, DECOMPOSITION, ELECTRON IRRADIATION, ELECTROSTATIC ACCELERATORS, IONIZATION POTENTIALS, LABORATORY EQUIPMENT, LITHIUM

  16. Ion Beam Propulsion Study

    NASA Technical Reports Server (NTRS)

    2008-01-01

    The Ion Beam Propulsion Study was a joint high-level study between the Applied Physics Laboratory operated by NASA and ASRC Aerospace at Kennedy Space Center, Florida, and Berkeley Scientific, Berkeley, California. The results were promising and suggested that work should continue if future funding becomes available. The application of ion thrusters for spacecraft propulsion is limited to quite modest ion sources with similarly modest ion beam parameters because of the mass penalty associated with the ion source and its power supply system. Also, the ion source technology has not been able to provide very high-power ion beams. Small ion beam propulsion systems were used with considerable success. Ion propulsion systems brought into practice use an onboard ion source to form an energetic ion beam, typically Xe+ ions, as the propellant. Such systems were used for steering and correction of telecommunication satellites and as the main thruster for the Deep Space 1 demonstration mission. In recent years, "giant" ion sources were developed for the controlled-fusion research effort worldwide, with beam parameters many orders of magnitude greater than the tiny ones of conventional space thruster application. The advent of such huge ion beam sources and the need for advanced propulsion systems for exploration of the solar system suggest a fresh look at ion beam propulsion, now with the giant fusion sources in mind.

  17. Ion propulsion and Comet Halley rendezvous

    NASA Technical Reports Server (NTRS)

    Atkins, K. L.

    1979-01-01

    Cometary rendezvous missions using ion propulsion is considered. The characteristics of the ion engine are discussed including the fuel efficiency and acceleration, and the design of the ion engine is described. The operation of the ion drive engine and an overview of its applications are presented.

  18. Ion propulsion cost effectivity

    NASA Technical Reports Server (NTRS)

    Zafran, S.; Biess, J. J.

    1978-01-01

    Ion propulsion modules employing 8-cm thrusters and 30-cm thrusters were studied for Multimission Modular Spacecraft (MMS) applications. Recurring and nonrecurring cost elements were generated for these modules. As a result, ion propulsion cost drivers were identified to be Shuttle charges, solar array, power processing, and thruster costs. Cost effective design approaches included short length module configurations, array power sharing, operation at reduced thruster input power, simplified power processing units, and power processor output switching. The MMS mission model employed indicated that nonrecurring costs have to be shared with other programs unless the mission model grows. Extended performance missions exhibited the greatest benefits when compared with monopropellant hydrazine propulsion.

  19. Nuclear propulsion systems engineering

    SciTech Connect

    Madsen, W.W.; Neuman, J.E.: Van Haaften, D.H.

    1992-12-31

    The Nuclear Energy for Rocket Vehicle Application (NERVA) program of the 1960`s and early 1970`s was dramatically successful, with no major failures during the entire testing program. This success was due in large part to the successful development of a systems engineering process. Systems engineering, properly implemented, involves all aspects of the system design and operation, and leads to optimization of theentire system: cost, schedule, performance, safety, reliability, function, requirements, etc. The process must be incorporated from the very first and continued to project completion. This paper will discuss major aspects of the NERVA systems engineering effort, and consider the implications for current nuclear propulsion efforts.

  20. Nuclear propulsion systems engineering

    SciTech Connect

    Madsen, W.W.; Neuman, J.E.: Van Haaften, D.H.

    1992-01-01

    The Nuclear Energy for Rocket Vehicle Application (NERVA) program of the 1960's and early 1970's was dramatically successful, with no major failures during the entire testing program. This success was due in large part to the successful development of a systems engineering process. Systems engineering, properly implemented, involves all aspects of the system design and operation, and leads to optimization of theentire system: cost, schedule, performance, safety, reliability, function, requirements, etc. The process must be incorporated from the very first and continued to project completion. This paper will discuss major aspects of the NERVA systems engineering effort, and consider the implications for current nuclear propulsion efforts.

  1. Electrostatic propulsion beam divergence effects on spacecraft surfaces, volume 3. [effects of ion engine experiment on subsystems of ATS 6 satellite

    NASA Technical Reports Server (NTRS)

    Kemp, R. F.; Hall, D. F.; Luedke, E. E.

    1973-01-01

    Tests were conducted to determine the effects of electrostatic propulsion beam divergence effects on spacecraft surfaces. The subjects discussed are: (1) sensitive surfaces on the ATS 6 spacecraft, (2) the cesium ion source and testing facility, (3) cesium ion effects on thermophysical properties, and (4) simulated charge-exchange ion exposure. The compatibility of the ATS 6 ion engine experiment with the engineering subsystems and other experiments aboard the ATS 6 spacecraft was analyzed.

  2. Electromagnetic interference assessment of an ion drive electric propulsion system

    NASA Technical Reports Server (NTRS)

    Whittlesey, A. C.

    1981-01-01

    An electric propulsion thrust system has the capability of providing a high specific impulse for long duration scientific missions in space. The EMI from the elements of an ion engine was characterized. The compatibility of ion drive electric propulsion systems with typical interplanetary spacecraft engineering was predicted.

  3. Electromagnetic interference assessment of an ion drive electric propulsion system

    NASA Technical Reports Server (NTRS)

    Whittlesey, A. C.

    1981-01-01

    An electric propulsion thrust system has the capability of providing a high specific impulse for long duration scientific missions in space. The EMI from the elements of an ion engine was characterized. The compatibility of ion drive electric propulsion systems with typical interplanetary spacecraft engineering was predicted.

  4. Next-Generation Ion Propulsion Being Developed

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Soulas, George C.; Foster, John E.; Haag, Thomas W.; Pinero, Luis R.; Rawlin, Vincent K.; Doehne, S. Michelle

    2001-01-01

    The NASA Glenn Research Center ion-propulsion program addresses the need for high specific-impulse systems and technology across a broad range of mission applications and power levels. One activity is the development of the next-generation ion-propulsion system as a follow-on to the successful Deep Space 1 system. The system is envisioned to incorporate a lightweight ion engine that can operate over 1 to 10 kW, with a 550-kg propellant throughput capacity. The engine concept under development has a 40-cm beam diameter, twice the effective area of the Deep Space 1 engine. It incorporates mechanical features and operating conditions to maximize the design heritage established by the Deep Space 1 engine, while incorporating new technology where warranted to extend the power and throughput capability. Prototype versions of the engine have been fabricated and are under test at NASA, with an engineering model version in manufacturing. Preliminary performance data for the prototype engine have been documented over 1.1- to 7.3-kW input power. At 7.3 kW, the engine efficiency is 0.68, at 3615-sec specific impulse. Critical component temperatures, including those of the discharge cathode assembly and magnets, have been documented and are within established limits, with significant margins relative to the Deep Space 1 engine. The 1- to 10-kW ion thruster approach described here was found to provide the needed power and performance improvement to enable important NASA missions. The Integrated In-Space Transportation Planning (IISTP) studies compared many potential technologies for various NASA, Government, and commercial missions. These studies indicated that a high-power ion propulsion system is the most important technology for development because of its outstanding performance versus perceived development and recurring costs for interplanetary solar electric propulsion missions. One of the best applications of a highpower electric propulsion system was as an integral part

  5. NSTAR Ion Propulsion System Power Electronics

    NASA Technical Reports Server (NTRS)

    1996-01-01

    The NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) program, managed by the Jet Propulsion Laboratory (JPL), is currently developing a high performance, simplified ion propulsion system. This propulsion system, which is throttleable from 0.5- to 2.3-kW output power to the thruster, targets primary propulsion applications for planetary and Earth-space missions and has been baselined as the primary propulsion system for the first New Millennium spacecraft. The NASA Lewis Research Center is responsible for the design and delivery of a breadboard power processing unit (PPU) and an engineering model thruster (EMT) for this system and will manage the contract for the delivery of the flight hardware to JPL. The PPU requirements, which dictate a mass of less than 12 kg with an efficiency of 0.9 or greater at a 2.3-kW output, forced a departure from the state-of-the-art ion thruster PPU design. Several innovations--including dual-use topologies, simplified thruster control, and the use of ferrite magnetic materials--were necessary to meet these requirements.

  6. Solar-Powered Electric Propulsion Systems: Engineering and Applications

    NASA Technical Reports Server (NTRS)

    Stearns, J. W.; Kerrisk, D. J.

    1966-01-01

    Lightweight, multikilowatt solar power arrays in conjunction with electric propulsion offer potential improvements to space exploration, extending the usefulness of existing launch vehicles to higher-energy missions. Characteristics of solar-powered electric propulsion missions are outlined, and preliminary performance estimates are shown. Spacecraft system engineering is discussed with respect to parametric trade-offs in power and propulsion system design. Relationships between mission performance and propulsion system performance are illustrated. The present state of the art of electric propulsion systems is reviewed and related to the mission requirements identified earlier. The propulsion system design and test requirements for a mission spacecraft are identified and discussed. Although only ion engine systems are currently available, certain plasma propulsion systems offer some advantages in over-all system design. These are identified, and goals are set for plasma-thrustor systems to make them competitive with ion-engine systems for mission applications.

  7. Ion electric propulsion unit

    DOEpatents

    Light, Max E; Colestock, Patrick L

    2014-01-28

    An electron cyclotron resonance (ECR) thruster is disclosed having a plasma chamber which is electrically biased with a positive voltage. The chamber bias serves to efficiently accelerate and expel the positive ions from the chamber. Electrons follow the exiting ions, serving to provide an electrically neutral exhaust plume. In a further embodiment, a downstream shaping magnetic field serves to further accelerate and/or shape the exhaust plume.

  8. Ion Propulsion Technology Programs at NASA Glenn Research Center

    NASA Technical Reports Server (NTRS)

    Patterson, M. J.; Oleson, S. R.

    2001-01-01

    As lead center for the agency in electric and ion propulsion, the NASA Glenn Research Center (GRC) is pursuing technology development in ion propulsion for a range of mission applications. The program goal is to develop key technologies for advanced NSTAR-derivative high-power ion propulsion, lightweight low power high-performance ion propulsion, 'micro' ion propulsion, and engine and component technologies for high-power electric propulsion for very ambitious missions. Products include: (1) a 5 kW, 400 kg throughput ion thruster and power processing technology; (2) extremely-lightweight high-efficiency sub-kilowatt ion thruster and power processor; (3) a 1-25 W high-specific impulse ion engine; and (4) engine and component technologies for high-power (30 kW class) ion and Hall engines. Identified applications include outer planetary science missions such as Europa orbiter/lander, Comet Nucleus Sample Return mission, Titan Explorer, Neptune/Triton, Pluto-Kuiper Belt Objects Mission, various second generation interplanetary Micro spacecraft, and the Interstellar Probe Mission. Additional information is contained in the original extended abstract.

  9. Ion Engine Test Firing

    NASA Technical Reports Server (NTRS)

    1999-01-01

    This image of a xenon ion engine, photographed through a port of the vacuum chamber where it was being tested at NASA's Jet Propulsion Laboratory, shows the faint blue glow of charged atoms being emitted from the engine. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century. Ion propulsion was first proposed in the 1950s and NASA performed experiments on this highly efficient propulsion system in the 1960s, but it was not used aboard an American spacecraft until the 1990s. Deep Space 1 was launched in October 1998 as part of NASA's New Millennium Program, which is managed by JPL for NASA's Office of Space Science, Washington, DC. The California Institute of Technology in Pasadena manages JPL for NASA. The almost imperceptible thrust from the ion propulsion system is equivalent to the pressure exerted by a sheet of paper held in the palm of your hand. The ion engine is very slow to pick up speed, but over the long haul it can deliver 10 times as much thrust per pound of fuel as more traditional rockets. Unlike the fireworks of most chemical rockets using solid or liquid fuels, the ion drive emits only an eerie blue glow as ionized (electrically charged) atoms of xenon are pushed out of the engine. Xenon is the same gas found in photo flash tubes and many lighthouse bulbs.

  10. NEXT Ion Propulsion System Development Status and Performance

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Benson, Scott W.

    2008-01-01

    NASA s Evolutionary Xenon Thruster (NEXT) project is developing next generation ion propulsion technologies to provide future NASA science missions with enhanced mission performance benefit at a low total development cost. The objective of the NEXT project is to advance next generation ion propulsion technology by producing engineering model and prototype model system components, validating these through qualification-level and integrated system testing, and ensuring preparedness for transitioning to flight system development. This paper describes the NEXT ion propulsion system development status, characteristics and performance. A review of mission analyses results conducted to date using the NEXT system is also provided.

  11. Deep Space 1 Ion Engine

    NASA Image and Video Library

    2002-12-21

    This image of a xenon ion engine, photographed through a port of the vacuum chamber where it was being tested at NASA's Jet Propulsion Laboratory, shows the faint blue glow of charged atoms being emitted from the engine. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. Though the thrust of the ion propulsion is about the same as the downward pressure of a single sheet of paper, by the end of the mission, the ion engine will have changed the spacecraft speed by about 13,700 kilometers/hour (8500 miles/hour). Even then, it will have expended only about 64 kg of its 81.5 kg supply of xenon propellant. http://photojournal.jpl.nasa.gov/catalog/PIA04247

  12. Spare Ion Engine Being Checked

    NASA Technical Reports Server (NTRS)

    2003-01-01

    July 21, 2003

    An ion thruster is removed from a vacuum chamber at NASA's Jet Propulsion Laboratory, Pasadena, Calif., its job done following almost five years of testing. Engineers John Anderson and Keith Goodfellow, from left, are part of JPL's Advanced Propulsion Technology Group. The thruster, a spare engine from NASA's Deep Space 1 mission, ran for a record 30,352 hours, giving researchers the ability to observe its performance and wear at different power levels throughout the test. This information will be vital to future missions that use ion propulsion.

    Ion propulsion systems can be very lightweight, running on just a few grams of xenon gas a day. This fuel efficiency can lower launch vehicle costs. Xenon is the same gas that is found in photo flash bulbs. The very successful Deep Space 1 mission featured the first use of an ion engine as the primary means of propulsion on a NASA spacecraft.

    NASA's next-generation ion propulsion efforts are led by the In-Space Propulsion Program, managed by the Office of Space Science at NASA Headquarters, Washington, D.C., and implemented by the Marshall Space Flight Center, Huntsville, Ala.. The program seeks to develop advanced propulsion technologies that will help near and mid-term NASA science missions by significantly reducing cost, mass or travel times. JPL is managed by the California Institute of Technology, Pasadena, Calif., for NASA.

  13. Implementation of the dawn ion propulsion system

    NASA Technical Reports Server (NTRS)

    Brophy, John; Marcucci, Michael G.; Ganapath, Gani B.; Gates, Jason; Garner, Charles E.; Klatte, Marlin; Lo, John; Nakazono, Barry; Pixler, Greg

    2005-01-01

    The Dawn ion propulsion system (IPS) was intended to be simply a larger version of the ion propulsion system that flew on Deep Space 1 (DS1). Implementation of this system to meet the needs of the Dawn mission, however, required modification of some IPS components and completely new developments of others.

  14. NASA Propulsion Engineering Research Center, Volume 2

    NASA Technical Reports Server (NTRS)

    1994-01-01

    This is the second volume in the 1994 annual report for the NASA Propulsion Engineering Research Center's Sixth Annual Symposium. This conference covered: (1) Combustors and Nozzles; (2) Turbomachinery Aero- and Hydro-dynamics; (3) On-board Propulsion systems; (4) Advanced Propulsion Applications; (5) Vaporization and Combustion; (6) Heat Transfer and Fluid Mechanics; and (7) Atomization and Sprays.

  15. NASA Propulsion Engineering Research Center, volume 2

    NASA Technical Reports Server (NTRS)

    1993-01-01

    On 8-9 Sep. 1993, the Propulsion Engineering Research Center (PERC) at The Pennsylvania State University held its Fifth Annual Symposium. PERC was initiated in 1988 by a grant from the NASA Office of Aeronautics and Space Technology as a part of the University Space Engineering Research Center (USERC) program; the purpose of the USERC program is to replenish and enhance the capabilities of our Nation's engineering community to meet its future space technology needs. The Centers are designed to advance the state-of-the-art in key space-related engineering disciplines and to promote and support engineering education for the next generation of engineers for the national space program and related commercial space endeavors. Research on the following areas was initiated: liquid, solid, and hybrid chemical propulsion, nuclear propulsion, electrical propulsion, and advanced propulsion concepts.

  16. Advanced ion propulsion systems for affordable deep-space missions

    NASA Astrophysics Data System (ADS)

    Brophy, John

    2003-01-01

    A key feature of future deep-space science missions will be the need for significantly greater on-board propulsion capability. To meet this need, ion propulsion based on the technology that flew on NASA's Deep Space 1 spacecraft has now entered the mainstream of propulsion options available for deep-space missions. The next most likely science mission to use ion propulsion is the comet nucleus sample return (CNSR) mission. CNSR has recently been identified by the Solar System Exploration Subcommittee as the highest priority new mission for NASA's Exploration of the Solar System theme. Ion propulsion for CNSR enables the use of a smaller, less expensive launch vehicle, and significantly shortens the overall trip time. A trade study for CNSR was performed to identify engine and system technology improvements, which provide the greatest mission benefits for the lowest additional risk. This trade study indicated that the maximum specific impulse of the ion engine should be increased from 3100 to 3800 s and that the maximum engine input power should be increased from 2.3 to 3.2 kW. Simultaneously the engine total propellant throughput capability must be increased from the 80-kg NSTAR design point to approximately 180 kg. A focused technology program to make these advances is underway.

  17. Xenon ion propulsion for orbit transfer

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.; Patterson, M. J.; Gruber, R. P.

    1990-01-01

    For more than 30 years, NASA has conducted an ion propulsion program which has resulted in several experimental space flight demonstrations and the development of many supporting technologies. Technologies appropriate for geosynchronous stationkeeping, earth-orbit transfer missions, and interplanetary missions are defined and evaluated. The status of critical ion propulsion system elements is reviewed. Electron bombardment ion thrusters for primary propulsion have evolved to operate on xenon in the 5 to 10 kW power range. Thruster efficiencies of 0.7 and specific impulse values of 4000 s were documented. The baseline thruster currently under development by NASA LeRC includes ring-cusp magnetic field plasma containment and dished two-grid ion optics. Based on past experience and demonstrated simplifications, power processors for these thrusters should have approximately 500 parts, a mass of 40 kg, and an efficiency near 0.94. Thrust vector control, via individual thruster gimbals, is a mature technology. High pressure, gaseous xenon propellant storage and control schemes, using flight qualified hardware, result in propellant tankage fractions between 0.1 and 0.2. In-space and ground integration testing has demonstrated that ion propulsion systems can be successfully integrated with their host spacecraft. Ion propulsion system technologies are mature and can significantly enhance and/or enable a variety of missions in the nation's space propulsion program.

  18. NEXT Ion Propulsion System Development Status and Capabilities

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Benson, Scott W.

    2008-01-01

    NASA s Evolutionary Xenon Thruster (NEXT) project is developing next generation ion propulsion technologies to provide future NASA science missions with enhanced mission performance benefit at a low total development cost. The objective of the NEXT project is to advance next generation ion propulsion technology by producing engineering model system components, validating these through qualification-level and integrated system testing, and ensuring preparedness for transitioning to flight system development. As NASA s Evolutionary Xenon Thruster technology program completes advanced development activities, it is advantageous to review the existing technology capabilities of the system under development. This paper describes the NEXT ion propulsion system development status, characteristics and performance. A review of mission analyses results conducted to date using the NEXT system is also provided.

  19. NASA Propulsion Engineering Research Center, volume 1

    NASA Technical Reports Server (NTRS)

    1993-01-01

    Over the past year, the Propulsion Engineering Research Center at The Pennsylvania State University continued its progress toward meeting the goals of NASA's University Space Engineering Research Centers (USERC) program. The USERC program was initiated in 1988 by the Office of Aeronautics and Space Technology to provide an invigorating force to drive technology advancements in the U.S. space industry. The Propulsion Center's role in this effort is to provide a fundamental basis from which the technology advances in propulsion can be derived. To fulfill this role, an integrated program was developed that focuses research efforts on key technical areas, provides students with a broad education in traditional propulsion-related science and engineering disciplines, and provides minority and other under-represented students with opportunities to take their first step toward professional careers in propulsion engineering. The program is made efficient by incorporating government propulsion laboratories and the U.S. propulsion industry into the program through extensive interactions and research involvement. The Center is comprised of faculty, professional staff, and graduate and undergraduate students working on a broad spectrum of research issues related to propulsion. The Center's research focus encompasses both current and advanced propulsion concepts for space transportation, with a research emphasis on liquid propellant rocket engines. The liquid rocket engine research includes programs in combustion and turbomachinery. Other space transportation modes that are being addressed include anti-matter, electric, nuclear, and solid propellant propulsion. Outside funding supports a significant fraction of Center research, with the major portion of the basic USERC grant being used for graduate student support and recruitment. The remainder of the USERC funds are used to support programs to increase minority student enrollment in engineering, to maintain Center

  20. Micro turbine engines for drones propulsion

    NASA Astrophysics Data System (ADS)

    Dutczak, J.

    2016-09-01

    Development of micro turbine engines began from attempts of application of that propulsion source by group of enthusiasts of aviation model making. Nowadays, the domain of micro turbojet engines is treated on a par with “full size” aviation constructions. The dynamic development of these engines is caused not only by aviation modellers, but also by use of micro turbojet engines by army to propulsion of contemporary drones, i.e. Unmanned Aerial Vehicles (UAV) or Unmanned Aerial Systems (UAS). On the base of selected examples the state of art in the mentioned group of engines has been presented in the article.

  1. Xenon ion propulsion for orbit transfer

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.; Patterson, M. J.; Gruber, R. P.

    1990-01-01

    The status of critical ion propulsion system elements is reviewed. Electron bombardment ion thrusters for primary propulsion have evolved to operate on xenon in the 5-10 kW power range. Thruster efficiencies of 0.7 and specific impulse values of 4000 s have been documented. The baseline thruster currently under development by NASA LeRC includes ring-cusp magnetic field plasma containment and dished two-grid ion optics. Based on past experience and demonstrated simplifications, power processors for these thrusters should have approximately 500 parts, a mass of 40 kg, and an efficiency near 0.94. Thrust vector control, via individual thruster gimbals, is a mature technology. High pressure, gaseous xenon propellant storage and control schemes, using flight qualified hardware, result in propellant tankage fractions between 0.1 and 0.2. In-space and ground integration testing has demonstrated that ion propulsion systems can be successfully integrated with their host spacecraft.

  2. Gasdynamic Mirror (GDM) Fusion Propulsion Engine Experiment

    NASA Technical Reports Server (NTRS)

    1999-01-01

    The Gasdynamic Mirror, or GDM, is an example of a magnetic mirror-based fusion propulsion system. Its design is primarily consisting of a long slender solenoid surrounding a vacuum chamber that contains plasma. The bulk of the fusion plasma is confined by magnetic field generated by a series of toroidal-shaped magnets in the center section of the device. the purpose of the GDM Fusion Propulsion Experiment is to confirm the feasibility of the concept and to demonstrate many of the operational characteristics of a full-size plasma can be confined within the desired physical configuration and still reman stable. This image shows an engineer from Propulsion Research Technologies Division at Marshall Space Flight Center inspecting solenoid magnets-A, an integrate part of the Gasdynamic Mirror Fusion Propulsion Engine Experiment.

  3. Gasdynamic Mirror (GDM) Fusion Propulsion Engine Experiment

    NASA Technical Reports Server (NTRS)

    1999-01-01

    The Gasdynamic Mirror, or GDM, is an example of a magnetic mirror-based fusion propulsion system. Its design is primarily consisting of a long slender solenoid surrounding a vacuum chamber that contains plasma. The bulk of the fusion plasma is confined by magnetic field generated by a series of toroidal-shaped magnets in the center section of the device. the purpose of the GDM Fusion Propulsion Experiment is to confirm the feasibility of the concept and to demonstrate many of the operational characteristics of a full-size plasma can be confined within the desired physical configuration and still reman stable. This image shows an engineer from Propulsion Research Technologies Division at Marshall Space Flight Center inspecting solenoid magnets-A, an integrate part of the Gasdynamic Mirror Fusion Propulsion Engine Experiment.

  4. 46 CFR 121.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... 46 Shipping 4 2011-10-01 2011-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and engine...

  5. 46 CFR 121.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... 46 Shipping 4 2010-10-01 2010-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and engine...

  6. ARTEMIS orbit raising inflight experience with ion propulsion

    NASA Astrophysics Data System (ADS)

    Killinger, Rainer; Kukies, Ralf; Surauer, Michael; Tomasetto, Angeo; van Holtz, Leo

    2003-08-01

    To demonstrate and promote North/South station keeping (inclination control) using ion propulsion, ESA on July 12, 2001 onboard Ariane 510 launched its most advanced telecommunication satellite: ARTEMIS. Due to a launcher failure the satellite was injected into a useless too low elliptic orbit. The ARTEMIS mission was salvaged by an Alenia Spazio / Astrium / ESA team at Telespazio (Fucino) using in novel modes to operate the on-board chemical and ion propulsion systems provided by Astrium. Using the chemical propulsion_ system provided by Astrium GmbH - Lampoldshausen - the inital orbit, having an apogee of half the targeted altitude. was quickly upgraded to a safe circular parking orbit at 31000 km altitude. The Liquid Apogee Engine was fired in total 8 times to achieve apogee as well as perigee raising. The final orbit raising to geostationary altitude is being performed by means of the ion propulsion system (IPP) applied in a newly designed spacecraft attitude control mode. Alenia Spazio and Astrium, in close cooperation, quickly redesigned all control and data handling software modules affected since the original spacecraft configuration was designed for inclination control only and not to generate thrust with the ion engines in a direction tangential to the orbit. The flexibility of the IPP system consisting of 4 thruster assemblies, provided in its totality by Astrium including the 2 alignment mechanisms for precision thrust direction control, had proven invaluable. To demonstrate the technologies available in Europe and to enhanced reliability, Astrium implemented two different technologies: a Kaufmann type system (EITA) provided by Astrium Ltd. - Portsmouth; and a Radiofrequency Ion Thruster Assembly (RITA) provided by Astrium GmbH - Ottobrunn. Two ion engines of different technology were mounted side by side on one ITAM (Ion Thruster Alignment Mechanism) provided by Austrian Aerospace. Artemis, after EURECA launched on 31 July 1992 and retrieved on 1 July

  7. The MAUS nuclear space reactor with ion propulsion system

    NASA Astrophysics Data System (ADS)

    Mainardi, Enrico

    2006-06-01

    MAUS (Moltiplicatore Avanzato Ultracompatto Spaziale) is a nuclear reactor concept design capable to ensure a reliable, long-lasting, low-mass, compact energy supply needed for advanced, future space missions. The exploration of the solar system and the space beyond requires the development of nuclear energy generators for supplying electricity to space-bases, spacecrafts, probes or satellites, as well as for propelling ships in long space missions. For propulsion, the MAUS nuclear reactor could be used to power electric ion drive engines. An ion engine is able to build up to very high velocities, far greater than chemical propulsion systems, but has high power and long service requirements. The MAUS concept is described, together with the ion propulsion engine and together with the reference thermoionic process used to convert the thermal power into electricity. The design work has been performed at the Nuclear Engineering and Energy Conversion Department of the University of Rome "La Sapienza" starting from 1992 on an issue submitted by the Italian Space Agency (ASI), in cooperation with the research laboratories of ENEA.

  8. Power processing technology for spacecraft primary ion propulsion

    NASA Technical Reports Server (NTRS)

    Biess, J. J.; Inouye, L. Y.; Frye, R. J.

    1980-01-01

    Advanced technologies developed in support of Ion Propulsion power processing, including the power circuitry portion of the Series L-C Resonant Inverter, Beam Supply, power components, packaging and heat pipe cooling of the 30 cm Ion Engine Power Processor are described. Both the transistorized and SCR versions of the Series L-C Resonant Inverter Beam Supply are discussed. A BIMOD Ion Thruster/Power Processor Prototype Assembly is undergoing environmental and life testing. These advanced technologies can be applied advantageously to other applications of future high power space power processing equipment.

  9. Propulsion

    ERIC Educational Resources Information Center

    Air and Space, 1978

    1978-01-01

    An introductory discussion of aircraft propulsion is included along with diagrams and pictures of piston, turbojet, turboprop, turbofan, and jet engines. Also, a table on chemical propulsion is included. (MDR)

  10. Propulsion

    ERIC Educational Resources Information Center

    Air and Space, 1978

    1978-01-01

    An introductory discussion of aircraft propulsion is included along with diagrams and pictures of piston, turbojet, turboprop, turbofan, and jet engines. Also, a table on chemical propulsion is included. (MDR)

  11. 46 CFR 121.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... 46 Shipping 4 2012-10-01 2012-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and...

  12. 46 CFR 121.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... 46 Shipping 4 2013-10-01 2013-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and...

  13. 46 CFR 121.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... 46 Shipping 4 2014-10-01 2014-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and...

  14. 46 CFR 184.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... 46 Shipping 7 2010-10-01 2010-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

  15. 46 CFR 184.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... 46 Shipping 7 2013-10-01 2013-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

  16. 46 CFR 184.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... 46 Shipping 7 2011-10-01 2011-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

  17. 46 CFR 184.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... 46 Shipping 7 2014-10-01 2014-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

  18. 46 CFR 184.620 - Propulsion engine control systems.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... 46 Shipping 7 2012-10-01 2012-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

  19. Nuclear thermal propulsion engine cost trade studies

    SciTech Connect

    Paschall, R.K. )

    1993-01-10

    The NASA transportation strategy for the Mars Exploration architecture includes the use of nuclear thermal propulsion as the primary propulsion system for Mars transits. It is anticipated that the outgrowth of the NERVA/ROVER programs will be a nuclear thermal propulsion (NTP) system capable of providing the propulsion for missions to Mars. The specific impulse (Isp) for such a system is expected to be in the 870 s range. Trade studies were conducted to investigate whether or not it may be cost effective to invest in a higher performance (Isp[gt]870 s) engine for nuclear thermal propulsion for missions to Mars. The basic cost trades revolved around the amount of mass that must be transported to low-earth orbit prior to each Mars flight and the cost to launch that mass. The mass required depended on the assumptions made for Mars missions scenarios including piloted/cargo flights, number of Mars missions, and transit time to Mars. Cost parameters included launch cost, program schedule for development and operations, and net discount rate. The results were very dependent on the assumptions that were made. Under some assumptions, higher performance engines showed cost savings in the billions of dollars; under other assumptions, the additional cost to develop higher performance engines was not justified.

  20. ARTEMIS Orbit Raising Inflight Experience with Ion Propulsion

    NASA Astrophysics Data System (ADS)

    Killinger, Rainer

    2002-01-01

    To demonstrate and promote North/South station keeping (inclination control) using ion propulsion, ESA on July 12, 2001 onboard Ariane 510 launched its most advanced telecommunication satellite: ARTEMIS. Due to a launcher failure the satellite was injected into a useless too low elliptic orbit. The ARTEMIS mission was salvaged by the ALTEL/Astrium/ESA team at Telespazio (Fucino) using in novel modes of operation the on-board chemical and ion propulsion systems provided by Astrium. Using the chemical propulsion system provided by Astrium GmbH - Lampoldshausen - the inital orbit, having an apogee of half the targeted altitude. was quickly upgraded to a safe circular parking orbit at 31000 km altitude. The Liquid Apogee Engine was fired in total 8 times to achieve perigee as well as apogee raising. The final orbit raising to geostationary altitude is being performed by means of the ion propulsion system (IPP) applied in a newly designed spacecraft attitude control mode. Alenia Spazio and Astrium, in close cooperation, quickly redesigned all control and data handling software modules affected since the original spacecraft configuration was designed for inclination control only and not to generate thrust with the ion engines in a direction tangential to the orbit. The flexibility of the IPP system consisting of 4 thruster assemblies, provided in its totality by Astrium including the 2 alignment mechanisms for precision thrust direction control, had proven invaluable. To demonstrate the technologies available in Europe and to enhanced reliability, Astrium implemented two different technologies: a Kaufmann type system (EITA) provided by Astrium Ltd. - Portsmouth, and a Radiofrequency Ion Thruster Assembly (RITA) provided by Astrium GmbH - Ottobrunn. Two ion engines of different technology were mounted side by side on one ITAM (Ion Thruster Alignment Mechanism) provided by Austrian Aerospace. This paper, after a brief description of the ion propulsion system, will

  1. Ion Propulsion Module design and mission performance

    NASA Technical Reports Server (NTRS)

    Graf, J. E.; Boain, R. J.; Pawlik, E. V.; Pless, L. C.

    1978-01-01

    This paper describes the design options, processes and tradeoffs that occur during the establishment of viable Ion Drive vehicle and mission designs. The options identify those internal vehicle design alternatives which are being considered for future Ion Drive missions, such as sunlight concentrating arrays and direct drive thrust subsystems, and their effect on mission performance. Also, the highly interactive nature of the Ion Drive design process, which occurs between the spacecraft and mission designers, is described. The results of design tradeoffs, performed for three Ion Drive comet rendezvous missions, are presented. These results include the following: (1) the power profile is determined primarily by the trajectory while second order effects include the solar cell characteristics and array concentration factor and degradation; and (2) the dominant parameter in mission performance determination, Ion Propulsion Module (IPM) mass, and IPM design, is the total cell power evaluated without concentration, at the beginning of life and at 1 AU.

  2. The ion propulsion system for Dawn

    NASA Technical Reports Server (NTRS)

    Brophy, J. R.; Marcucci, M. G.; Ganapathi, G. B.; Garner, C. E.; Henry, M. D.; Nakazono, B.; Noon, D.

    2003-01-01

    The Dawn Project's mission is to rendezvous and map the two heaviest main belt asteroids Vesta and Ceres. The Ion Propulsion System (IPS) for Dawn will be used for the heliocentric transfer from the Earth to Vesta, orbit capture at Vesta, transfer to a low Vesta orbit, departure and escape from Vesta, the heliocentric transfer from Vesta to Ceres, orbit capture at Ceres, and transfer to a low Ceres orbit.

  3. An ion propulsion orbit manoeuvring vehicle

    NASA Astrophysics Data System (ADS)

    Holdaway, R.; Flaherty, M.

    1990-10-01

    It is shown that the concept of a reusable ion propulsion orbit maneuvering vehicle (OMV) can fulfill two particular orbital dynamics needs for the coming decades. On is the boosting of nonoperational satellites out of GEO in order to reduce GEO crowding and frequency transmission crosstalk. The other is the ability to maneuver certain satellites and other orbital debris out of the area in which the Space Station is to be located. The vehicle is flexible in operation, far less expensive than a chemical OMV, and requires very little fuel and maintenance.

  4. Fusion Propulsion Z-Pinch Engine Concept

    NASA Technical Reports Server (NTRS)

    Miernik, J.; Statham, G.; Fabisinski, L.; Maples, C. D.; Adams, R.; Polsgrove, T.; Fincher, S.; Cassibry, J.; Cortez, R.; Turner, M.; hide

    2011-01-01

    Fusion-based nuclear propulsion has the potential to enable fast interplanetary transportation. Due to the great distances between the planets of our solar system and the harmful radiation environment of interplanetary space, high specific impulse (Isp) propulsion in vehicles with high payload mass fractions must be developed to provide practical and safe vehicles for human spaceflight missions. The Z-Pinch dense plasma focus method is a Magneto-Inertial Fusion (MIF) approach that may potentially lead to a small, low cost fusion reactor/engine assembly1. Recent advancements in experimental and theoretical understanding of this concept suggest favorable scaling of fusion power output yield 2. The magnetic field resulting from the large current compresses the plasma to fusion conditions, and this process can be pulsed over short timescales (10(exp -6 sec). This type of plasma formation is widely used in the field of Nuclear Weapons Effects testing in the defense industry, as well as in fusion energy research. A Decade Module 2 (DM2), approx.500 KJ pulsed-power is coming to the RSA Aerophysics Lab managed by UAHuntsville in January, 2012. A Z-Pinch propulsion concept was designed for a vehicle based on a previous fusion vehicle study called "Human Outer Planet Exploration" (HOPE), which used Magnetized Target Fusion (MTF) 3 propulsion. The reference mission is the transport of crew and cargo to Mars and back, with a reusable vehicle.

  5. Progress in Technology Validation of the Next Ion Propulsion System

    NASA Technical Reports Server (NTRS)

    Benson, Scott W.; Patterson, Michael J.

    2007-01-01

    The NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system has been in advanced technology development under the NASA In-Space Propulsion Technology project. The highest fidelity hardware planned has now been completed by the government/industry team, including a flight prototype model (PM) thruster, an engineering model (EM) power processing unit, EM propellant management assemblies, a breadboard gimbal, and control unit simulators. Subsystem and system level technology validation testing is in progress. To achieve the objective Technology Readiness Level 6, environmental testing is being conducted to qualification levels in ground facilities simulating the space environment. Additional tests have been conducted to characterize the performance range and life capability of the NEXT thruster. This paper presents the status and results of technology validation testing accomplished to date, the validated subsystem and system capabilities, and the plans for completion of this phase of NEXT development.

  6. Engineering Challenges in Antiproton Triggered Fusion Propulsion

    SciTech Connect

    Cassenti, Brice; Kammash, Terry

    2008-01-21

    During the last decade antiproton triggered fusion propulsion has been investigated as a method for achieving high specific impulse, high thrust in a nuclear pulse propulsion system. In general the antiprotons are injected into a pellet containing fusion fuel with a small amount of fissionable material (i.e., an amount less than the critical mass) where the products from the fission are then used to trigger a fusion reaction. Initial calculations and simulations indicate that if magnetically insulated inertial confinement fusion is used that the pellets should result in a specific impulse of between 100,000 and 300,000 seconds at high thrust. The engineering challenges associated with this propulsion system are significant. For example, the antiprotons must be precisely focused. The pellet must be designed to contain the fission and initial fusion products and this will require strong magnetic fields. The fusion fuel must be contained for a sufficiently long time to effectively release the fusion energy, and the payload must be shielded from the radiation, especially the excess neutrons emitted, in addition to many other particles. We will review the recent progress, possible engineering solutions and the potential performance of these systems.

  7. Refan Engine in the Propulsion Systems Laboratory

    NASA Image and Video Library

    1974-10-21

    A refanned Pratt and Whitney JT-8D-109 turbofan engine installed in Cell 4 of the Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. NASA Lewis’ Refan Program sought to demonstrate that noise reduction modifications could be applied to existing aircraft engines with minimal costs and without diminishing the engine’s performance or integrity. At the time, Pratt and Whitney’s JT-8D turbofans were one of the most widely used engines in the commercial airline industry. The engines powered Boeing’s 727 and 737 and McDonnell Douglas’ DC-9 aircraft. Pratt and Whitney worked with the airline manufacturers on a preliminary study that verified feasibility of replacing the JT-8D’s two-stage fan with a larger single-stage fan. The new fan slowed the engine’s exhaust, which significantly reduced the amount of noise it generated. Booster stages were added to maintain the proper level of airflow through the engine. Pratt and Whitney produced six of the modified engines, designated JT-8D-109, and performed the initial testing. One of the JT-8D-109 engines, seen here, was tested in simulated altitude conditions in NASA Lewis’ Propulsion Systems Laboratory. The Refan engine was ground-tested on an actual aircraft before making a series of flight tests on 727 and DC-9 aircraft in early 1976. The Refan Program reduced the JT-8D’s noise by 50 percent while increasing the fuel efficiency. The retro-fit kits were estimated to cost between $1 million and $1.7 million per aircraft.

  8. Propulsion Mechanism of Catalytic Microjet Engines

    PubMed Central

    Fomin, Vladimir M.; Hippler, Markus; Magdanz, Veronika; Soler, Lluís; Sanchez, Samuel; Schmidt, Oliver G.

    2014-01-01

    We describe the propulsion mechanism of the catalytic microjet engines that are fabricated using rolled-up nanotech. Microjets have recently shown numerous potential applications in nanorobotics but currently there is a lack of an accurate theoretical model that describes the origin of the motion as well as the mechanism of self-propulsion. The geometric asymmetry of a tubular microjet leads to the development of a capillary force, which tends to propel a bubble toward the larger opening of the tube. Because of this motion in an asymmetric tube, there emerges a momentum transfer to the fluid. In order to compensate this momentum transfer, a jet force acting on the tube occurs. This force, which is counterbalanced by the linear drag force, enables tube velocities of the order of 100 μm/s. This mechanism provides a fundamental explanation for the development of driving forces that are acting on bubbles in tubular microjets. PMID:25177214

  9. Gravity-assist engine for space propulsion

    NASA Astrophysics Data System (ADS)

    Bergstrom, Arne

    2014-06-01

    As a possible alternative to rockets, the present article describes a new type of engine for space travel, based on the gravity-assist concept for space propulsion. The new engine is to a great extent inspired by the conversion of rotational angular momentum to orbital angular momentum occurring in tidal locking between astronomical bodies. It is also greatly influenced by Minovitch's gravity-assist concept, which has revolutionized modern space technology, and without which the deep-space probes to the outer planets and beyond would not have been possible. Two of the three gravitating bodies in Minovitch's concept are in the gravity-assist engine discussed in this article replaced by an extremely massive ‘springbell' (in principle a spinning dumbbell with a powerful spring) incorporated into the spacecraft itself, and creating a three-body interaction when orbiting around a gravitating body. This makes gravity-assist propulsion possible without having to find suitably aligned astronomical bodies. Detailed numerical simulations are presented, showing how an actual spacecraft can use a ca 10-m diameter springbell engine in order to leave the earth's gravitational field and enter an escape trajectory towards interplanetary destinations.

  10. Technology Readiness of the NEXT Ion Propulsion System

    NASA Technical Reports Server (NTRS)

    Benson, Scott W.; Patterson, Michael J.

    2008-01-01

    The NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system has been in advanced technology development under the NASA In-Space Propulsion Technology project. The highest fidelity hardware planned has now been completed by the government/industry team, including: a flight prototype model (PM) thruster, an engineering model (EM) power processing unit, EM propellant management assemblies, a breadboard gimbal, and control unit simulators. Subsystem and system level technology validation testing is in progress. To achieve the objective Technology Readiness Level 6, environmental testing is being conducted to qualification levels in ground facilities simulating the space environment. Additional tests have been conducted to characterize the performance range and life capability of the NEXT thruster. This paper presents the status and results of technology validation testing accomplished to date, the validated subsystem and system capabilities, and the plans for completion of this phase of NEXT development. The next round of competed planetary science mission announcements of opportunity, and directed mission decisions, are anticipated to occur in 2008 and 2009. Progress to date, and the success of on-going technology validation, indicate that the NEXT ion propulsion system will be a primary candidate for mission consideration in these upcoming opportunities.

  11. Plasma simulation in a hybrid ion electric propulsion system

    NASA Astrophysics Data System (ADS)

    Jugroot, Manish; Christou, Alex

    2015-04-01

    An exciting possibility for the next generation of satellite technology is the microsatellite. These satellites, ranging from 10-500 kg, can offer advantages in cost, reduced risk, and increased functionality for a variety of missions. For station keeping and control of these satellites, a suitable compact and high efficiency thruster is required. Electrostatic propulsion provides a promising solution for microsatellite thrust due to their high specific impulse. The rare gas propellant is ionized into plasma and generates a beam of high speed ions by electrostatic processes. A concept explored in this work is a hybrid combination of dc ion engines and hall thrusters to overcome space-charge and lifetime limitations of current ion thruster technologies. A multiphysics space and time-dependent formulation was used to investigate and understand the underlying physical phenomena. Several regions and time scales of the plasma have been observed and will be discussed.

  12. Deep Space 1 Ion Engine

    NASA Image and Video Library

    2002-12-21

    This image of a xenon ion engine prototype, photographed through a port of the vacuum chamber where it was being tested at NASA's Jet Propulsion Laboratory, shows the faint blue glow of charged atoms being emitted from the engine. The engine is now in an ongoing extended- life test, in a vacuum test chamber at JPL, and has run for almost 500 days (12,000 hours) and is scheduled to complete nearly 625 days (15,000 hours) by the end of 2001. A similar engine powers the New Millennium Program's flagship mission, Deep Space 1, which uses the ion engine in a trip through the solar system. The engine, weighing 17.6 pounds (8 kilograms), is 15.7 inches (40 centimeters) in diameter and 15.7 inches long. The actual thrust comes from accelerating and expelling positively charged xenon atoms, or ions. While the ions are fired in great numbers out the thruster at more than 110,000 kilometers (68,000 miles) per hour, their mass is so low that the engine produces a gentle thrust of only 90 millinewtons (20-thousandths of a pound). http://photojournal.jpl.nasa.gov/catalog/PIA04238

  13. CFD applications in chemical propulsion engines

    NASA Technical Reports Server (NTRS)

    Merkle, Charles L.

    1991-01-01

    The present research is aimed at developing analytical procedures for predicting the performance and stability characteristics of chemical propulsion engines. Specific emphasis is being placed on understanding the physical and chemical processes in the small engines that are used for applications such as spacecraft attitude control and drag make-up. The small thrust sizes of these engines lead to low nozzle Reynolds numbers with thick boundary layers which may even meet at the nozzle centerline. For this reason, the classical high Reynolds number procedures that are commonly used in the industry are inaccurate and of questionable utility for design. A complete analysis capability for the combined viscous and inviscid regions as well as for the subsonic, transonic, and supersonic portions of the flowfield is necessary to estimate performance levels and to enable tradeoff studies during design procedures.

  14. Simulation Based on Negative ion pair Techniques of Electric propulsion In Satellite Mission Using Chlorine Gas

    NASA Astrophysics Data System (ADS)

    Bakkiyaraj, R.

    R.Bakkiyaraj,Assistant professor,Government college of Engineering ,Bargur,Tamilnadu. *C.Sathiyavel, PG Student and Department of Aeronautical Engineering/Branch of Avionics, PSN college of Engineering and Technology,Tirunelveli,India. Abstract: Ion propulsion rocket system is expected to become popular with the development of ion-ion pair techniques because of their stimulated of low propellant, Design of repulsive between negative ions with low electric power and high efficiency. A Negative ion pair of ion propulsion rocket system is proposed in this work .Negative Ion Based Rocket system consists of three parts 1.ionization chamber 2. Repulsion force and ion accelerator 3. Exhaust of Nozzle. The Negative ions from electro negatively gas are produced by attachment of the gas ,such as chlorine with electron emitted from a Electron gun ionization chamber. The formulate of large stable negative ion is achievable in chlorine gas with respect to electron affinity (∆E). When a neutral chlorine atom in the gaseous form picks up an electron to form a cl- ion, it releases energy of 349 kJ/mol or 3.6 eV/atom. It is said to have an electron affinity of -349 kJ/mol ,the negative sign indicating that energy is released during this process .The distance between negative ions pair is important for the evaluation of the rocket thrust and is also determined by the exhaust velocity of the propellant. The mass flow rate of ions is related to the ion beam current. Accelerate the Negative ions to a high velocity in the thrust vector direction with a significantly intense grids and the exhaust of negative ions through Nozzle. The simulation of the ion propulsion system has been carried out by MATLAB. By comparing the simulation results with the theoretical and previous results, we have found that the proposed method is achieved of thrust value with low electric power for simulating the ion propulsion rocket system

  15. Deep Space 1 Ion Engine

    NASA Image and Video Library

    2002-12-21

    Kennedy Space Center, Florida. - Deep Space 1 is lifted from its work platform, giving a closeup view of the experimental solar-powered ion propulsion engine. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century. Another onboard experiment includes software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, Cape Canaveral Air Station, in October. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches. http://photojournal.jpl.nasa.gov/catalog/PIA04232

  16. 40 CFR 87.62 - Test procedure (propulsion engines).

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 21 2012-07-01 2012-07-01 false Test procedure (propulsion engines). 87.62 Section 87.62 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) Definitions. Test Procedures § 87.62 Test procedure (propulsion engines). Link to...

  17. Ion Propulsion Development Projects in US: Space Electric Rocket Test I to Deep Space 1

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Rawlin, Vincent K.; Patterson, Michael J.

    2001-01-01

    The historical background and characteristics of the experimental flights of ion propulsion systems and the major ground-based technology demonstrations are reviewed. The results of the first successful ion engine flight in 1964, Space Electric Rocket Test (SERT) I, which demonstrated ion beam neutralization, are discussed along with the extended operation of SERT II starting in 1970. These results together with the technologies employed on the early cesium engine flights, the applications technology satellite series, and the ground-test demonstrations, have provided the evolutionary path for the development of xenon ion thruster component technologies, control systems, and power circuit implementations. In the 1997-1999 period, the communication satellite flights using ion engine systems and the Deep Space 1 flight confirmed that these auxiliary and primary propulsion systems have advanced to a high level of flight readiness.

  18. The Ion Propulsion System for the Solar Electric Propulsion Technology Demonstration Mission

    NASA Technical Reports Server (NTRS)

    Herman, Daniel A.; Santiago, Walter; Kamhawi, Hani; Polk, James E.; Snyder, John Steven; Hofer, Richard R.; Parker, J. Morgan

    2015-01-01

    The Asteroid Redirect Robotic Mission is a candidate Solar Electric Propulsion Technology Demonstration Mission whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a companion human-crewed mission. The ion propulsion system must be capable of operating over an 8-year time period and processing up to 10,000 kg of xenon propellant. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as a critical part of an affordable, beyond-low-Earth-orbit, manned-exploration architecture. Under the NASA Space Technology Mission Directorate the critical electric propulsion and solar array technologies are being developed. The ion propulsion system being co-developed by the NASA Glenn Research Center and the Jet Propulsion Laboratory for the Asteroid Redirect Vehicle is based on the NASA-developed 12.5 kW Hall Effect Rocket with Magnetic Shielding (HERMeS0 thruster and power processing technologies. This paper presents the conceptual design for the ion propulsion system, the status of the NASA in-house thruster and power processing activity, and an update on flight hardware.

  19. Explorations of Psyche and Callisto Enabled by Ion Propulsion

    NASA Technical Reports Server (NTRS)

    Wenkert, Daniel D.; Landau, Damon F.; Bills, Bruce G.; Elkins-Tanton, Linda T.

    2013-01-01

    Recent developments in ion propulsion (specifically solar electric propulsion - SEP) have the potential for dramatically reducing the transportation cost of planetary missions. We examine two representative cases, where these new developments enable missions which, until recently, would have required resouces well beyond those allocated to the Discovery program. The two cases of interest address differentiation of asteroids and large icy satellites

  20. Explorations of Psyche and Callisto Enabled by Ion Propulsion

    NASA Technical Reports Server (NTRS)

    Wenkert, Daniel D.; Landau, Damon F.; Bills, Bruce G.; Elkins-Tanton, Linda T.

    2013-01-01

    Recent developments in ion propulsion (specifically solar electric propulsion - SEP) have the potential for dramatically reducing the transportation cost of planetary missions. We examine two representative cases, where these new developments enable missions which, until recently, would have required resouces well beyond those allocated to the Discovery program. The two cases of interest address differentiation of asteroids and large icy satellites

  1. The Ion Propulsion System on NASA's Space Technology 4/Champollion Comet Rendezvous Mission

    NASA Technical Reports Server (NTRS)

    Brophy, John R.; Garner, Charles E.; Weiss, Jeffery M.

    1999-01-01

    The ST4/Champollion mission is designed to rendezvous with and land on the comet Tempel 1 and return data from the first-ever sampling of a comet surface. Ion propulsion is an enabling technology for this mission. The ion propulsion system on ST4 consists of three ion engines each essentially identical to the single engine that flew on the DS1 spacecraft. The ST4 propulsion system will operate at a maximum input power of 7.5 kW (3.4 times greater than that demonstrated on DS1), will produce a maximum thrust of 276 mN, and will provide a total (Delta)V of 11.4 km/s. To accomplish this the propulsion system will carry 385 kg of xenon. All three engines will be operated simultaneously for the first 168 days of the mission. The nominal mission requires that each engine be capable of processing 118 kg. If one engine fails after 168 days, the remaining two engines can perform the mission, but must be capable of processing 160 kg of xenon, or twice the original thruster design requirement. Detailed analyses of the thruster wear-out failure modes coupled with experience from long-duration engine tests indicate that the thrusters have a high probability of meeting the 160-kg throughput requirement.

  2. Development Efforts Expanded in Ion Propulsion: Ion Thrusters Developed With Higher Power Levels

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Rawlin, Vincent K.; Sovey, James S.

    2003-01-01

    The NASA Glenn Research Center was the major contributor of 2-kW-class ion thruster technology to the Deep Space 1 mission, which was successfully completed in early 2002. Recently, NASA s Office of Space Science awarded approximately $21 million to Glenn to develop higher power xenon ion propulsion systems for large flagship missions such as outer planet explorers and sample return missions. The project, referred to as NASA's Evolutionary Xenon Thruster (NEXT), is a logical follow-on to the ion propulsion system demonstrated on Deep Space 1. The propulsion system power level for NEXT is expected to be as high as 25 kW, incorporating multiple ion thrusters, each capable of being throttled over a 1- to 6-kW power range. To date, engineering model thrusters have been developed, and performance and plume diagnostics are now being documented. The project team-Glenn, the Jet Propulsion Laboratory, General Dynamics, Boeing Electron Dynamic Devices, the Applied Physics Laboratory, the University of Michigan, and Colorado State University-is in the process of developing hardware for a ground demonstration of the NEXT propulsion system, which comprises a xenon feed system, controllers, multiple thrusters, and power processors. The development program also will include life assessments by tests and analyses, single-string tests of ion thrusters and power systems, and finally, multistring thruster system tests in calendar year 2005. In addition, NASA's Office of Space Science selected Glenn to lead the development of a 25-kW xenon thruster to enable NASA to conduct future missions to the outer planets of Jupiter and beyond, under the High Power Electric Propulsion (HiPEP) program. The development of a 100-kW-class ion propulsion system and power conversion systems are critical components to enable future nuclear-electric propulsion systems. In fiscal year 2003, a team composed of Glenn, the Boeing Company, General Dynamics, the Applied Physics Laboratory, the Naval Research

  3. Development Efforts Expanded in Ion Propulsion: Ion Thrusters Developed With Higher Power Levels

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Rawlin, Vincent K.; Sovey, James S.

    2003-01-01

    The NASA Glenn Research Center was the major contributor of 2-kW-class ion thruster technology to the Deep Space 1 mission, which was successfully completed in early 2002. Recently, NASA s Office of Space Science awarded approximately $21 million to Glenn to develop higher power xenon ion propulsion systems for large flagship missions such as outer planet explorers and sample return missions. The project, referred to as NASA's Evolutionary Xenon Thruster (NEXT), is a logical follow-on to the ion propulsion system demonstrated on Deep Space 1. The propulsion system power level for NEXT is expected to be as high as 25 kW, incorporating multiple ion thrusters, each capable of being throttled over a 1- to 6-kW power range. To date, engineering model thrusters have been developed, and performance and plume diagnostics are now being documented. The project team-Glenn, the Jet Propulsion Laboratory, General Dynamics, Boeing Electron Dynamic Devices, the Applied Physics Laboratory, the University of Michigan, and Colorado State University-is in the process of developing hardware for a ground demonstration of the NEXT propulsion system, which comprises a xenon feed system, controllers, multiple thrusters, and power processors. The development program also will include life assessments by tests and analyses, single-string tests of ion thrusters and power systems, and finally, multistring thruster system tests in calendar year 2005. In addition, NASA's Office of Space Science selected Glenn to lead the development of a 25-kW xenon thruster to enable NASA to conduct future missions to the outer planets of Jupiter and beyond, under the High Power Electric Propulsion (HiPEP) program. The development of a 100-kW-class ion propulsion system and power conversion systems are critical components to enable future nuclear-electric propulsion systems. In fiscal year 2003, a team composed of Glenn, the Boeing Company, General Dynamics, the Applied Physics Laboratory, the Naval Research

  4. The Ion Propulsion System for the Solar Electric Propulsion Technology Demonstration Mission

    NASA Technical Reports Server (NTRS)

    Herman, Daniel A.; Santiago, Walter; Kamhawi, Hani; Polk, James E.; Snyder, John Steven; Hofer, Richard; Parker, J. Morgan

    2015-01-01

    The Asteroid Redirect Robotic Mission is a candidate Solar Electric Propulsion Technology Demonstration Mission whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a subsequent human-crewed mission. The ion propulsion subsystem must be capable of operating over an 8-year time period and processing up to 10,000 kg of xenon propellant. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as an enabling element of an affordable beyond low-earth orbit human-crewed exploration architecture. Under the NASA Space Technology Mission Directorate the critical electric propulsion and solar array technologies are being developed. The ion propulsion system for the Asteroid Redirect Vehicle is based on the NASA-developed 12.5 kW Hall Effect Rocket with Magnetic Shielding thruster and power processing technologies. This paper presents the conceptual design for the ion propulsion system, a status on the NASA in-house thruster and power processing is provided, and an update on acquisition for flight provided.

  5. Optimization of Air-Breathing Propulsion Engine Concepts

    NASA Technical Reports Server (NTRS)

    Patnaik, Surya N.; Hopkins, Dale A.

    1997-01-01

    Air-breathing propulsion engines play an important role in the development of both civil and military aircraft Design optimization of such engines can lead to higher power, or more thrust for less fuel consumption. A multimission propulsion engine design can be modeled mathematically as a multivariable global optimization problem, with a sequence of subproblems, which are specific to the mission events defined through Mach number, altitude, and power setting combinations.

  6. Engineering of the Magnetized Target Fusion Propulsion System

    NASA Technical Reports Server (NTRS)

    Statham, G.; White, S.; Adams, R. B.; Thio, Y. C. F.; Santarius, J.; Alexander, R.; Chapman, J.; Fincher, S.; Philips, A.; Polsgrove, T.

    2003-01-01

    Engineering details are presented for a magnetized target fusion (MTF) propulsion system designed to support crewed missions to the outer solar system. Basic operation of an MTF propulsion system is introduced. Structural, thermal, radiation-management and electrical design details are presented. The propellant storage and supply system design is also presented. A propulsion system mass estimate and associated performance figures are given. The advantages of helium-3 as a fusion fuel for an advanced MTF system are discussed.

  7. Propulsion of ripples on glass by ion bombardment.

    PubMed

    Alkemade, P F A

    2006-03-17

    The propulsion of surface ripples on SiO(2) by an ion beam was investigated by in situ electron microscopy. The observed propagation of the ripples contradicts existing models for ion-beam-induced ripple development. A new model based on the Navier-Stokes relations for viscous flow in a thin layer is introduced. It includes inhomogeneous viscous flow, driven by spatial variations in the deposition of the energy of the ion beam. The model explains the observed reversed propagation. The hitherto unknown propulsion mechanism is important for understanding nanoscale pattern formation by ion bombardment.

  8. Propulsion of Ripples on Glass by Ion Bombardment

    SciTech Connect

    Alkemade, P.F.A.

    2006-03-17

    The propulsion of surface ripples on SiO{sub 2} by an ion beam was investigated by in situ electron microscopy. The observed propagation of the ripples contradicts existing models for ion-beam-induced ripple development. A new model based on the Navier-Stokes relations for viscous flow in a thin layer is introduced. It includes inhomogeneous viscous flow, driven by spatial variations in the deposition of the energy of the ion beam. The model explains the observed reversed propagation. The hitherto unknown propulsion mechanism is important for understanding nanoscale pattern formation by ion bombardment.

  9. Engineering of the Magnetized Target Fusion Propulsion System

    NASA Technical Reports Server (NTRS)

    Statham, G.; White, S.; Adams, R. B.; Thio, Y. C. F.; Santarius, J.; Alexander, R.; Fincher, S.; Polsgrove, T.; Chapman, J.; Philips, A.

    2002-01-01

    Engineering details are presented for a magnetized target fusion (MTF) propulsion system designed to support crewed missions to the outer solar system. Structural, thermal and radiation-management design details are presented. Propellant storage and supply options are also discussed and a propulsion system mass estimate is given.

  10. An overview of the Penn State Propulsion Engineering Research Center

    NASA Technical Reports Server (NTRS)

    Merkle, Charles L.

    1991-01-01

    An overview of the Penn State Propulsion Engineering Research Center is presented. The following subject areas are covered: research objectives and long term perspective of the Center; current status and operational philosophy; and brief description of Center projects (combustion, fluid mechanics and heat transfer, materials compatibility, turbomachinery, and advanced propulsion concepts).

  11. Antiproton powered propulsion with magnetically confined plasma engines

    NASA Technical Reports Server (NTRS)

    Lapointe, Michael R.

    1989-01-01

    The reaction of the matter-antimatter annihilation, with its specific energy being over 250 times the specific energy released in nuclear fusion, is considered as an energy source for spacecraft propulsion. A concept of a magnetically confined pulsed plasma engine is described. In this concept, antiproton beams are injected axially into a pulsed magnetic mirror system, where they annihilate with an initially neutral hydrogen gas; the resulting charge annihilation products transfer energy to the hydrogen propellant, which is then exhausted through one end of the pulsed mirror system to provide thrust. Numerical simulations were developed to calculate the annihilation rate of antiprotons in hydrogen and to follow the resulting ion, muon, and electron/positron number density evolutions.

  12. Antiproton powered propulsion with magnetically confined plasma engines

    SciTech Connect

    Lapointe, M.R.

    1989-01-01

    The reaction of the matter-antimatter annihilation, with its specific energy being over 250 times the specific energy released in nuclear fusion, is considered as an energy source for spacecraft propulsion. A concept of a magnetically confined pulsed plasma engine is described. In this concept, antiproton beams are injected axially into a pulsed magnetic mirror system, where they annihilate with an initially neutral hydrogen gas; the resulting charge annihilation products transfer energy to the hydrogen propellant, which is then exhausted through one end of the pulsed mirror system to provide thrust. Numerical simulations were developed to calculate the annihilation rate of antiprotons in hydrogen and to follow the resulting ion, muon, and electron/positron number density evolutions. 22 refs.

  13. Antiproton powered propulsion with magnetically confined plasma engines

    NASA Technical Reports Server (NTRS)

    Lapointe, Michael R.

    1989-01-01

    The reaction of the matter-antimatter annihilation, with its specific energy being over 250 times the specific energy released in nuclear fusion, is considered as an energy source for spacecraft propulsion. A concept of a magnetically confined pulsed plasma engine is described. In this concept, antiproton beams are injected axially into a pulsed magnetic mirror system, where they annihilate with an initially neutral hydrogen gas; the resulting charge annihilation products transfer energy to the hydrogen propellant, which is then exhausted through one end of the pulsed mirror system to provide thrust. Numerical simulations were developed to calculate the annihilation rate of antiprotons in hydrogen and to follow the resulting ion, muon, and electron/positron number density evolutions.

  14. The Ion Propulsion System for the Asteroid Redirect Robotic Mission

    NASA Technical Reports Server (NTRS)

    Herman, Daniel A.; Santiago, Walter; Kamhawi, Hani; Polk, James E.; Snyder, John Steven; Hofer, Richard R.; Sekerak, Michael J.

    2016-01-01

    The Asteroid Redirect Robotic Mission is a Solar Electric Propulsion Technology Demonstration Mission (ARRM) whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a companion human-crewed mission. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as a critical part of NASA'a future beyond-low-Earth-orbit, human-crewed exploration plans. Under the NASA Space Technology Mission Directorate the critical electric propulsion and solar array technologies are being developed. This paper presents the conceptual design of the ARRM ion propulsion system, the status of the NASA in-house thruster and power processing development activities, the status of the planned technology maturation for the mission through flight hardware delivery, and the status of the mission formulation and spacecraft acquisition.

  15. Energy efficient engine: Propulsion system-aircraft integration evaluation

    NASA Technical Reports Server (NTRS)

    Owens, R. E.

    1979-01-01

    Flight performance and operating economics of future commercial transports utilizing the energy efficient engine were assessed as well as the probability of meeting NASA's goals for TSFC, DOC, noise, and emissions. Results of the initial propulsion systems aircraft integration evaluation presented include estimates of engine performance, predictions of fuel burns, operating costs of the flight propulsion system installed in seven selected advanced study commercial transports, estimates of noise and emissions, considerations of thrust growth, and the achievement-probability analysis.

  16. Evolution of engine cycles for STOVL propulsion concepts

    NASA Technical Reports Server (NTRS)

    Bucknell, R. L.; Frazier, R. H.; Giulianetti, D. J.

    1990-01-01

    Short Take-off, Vertical Landing (STOVL) demonstrator concepts using a common ATF engine core are discussed. These concepts include a separate fan and core flow engine cycle, mixed flow STOVL cycles, separate flow cycles convertible to mixed flow, and reaction control system engine air bleed. STOVL propulsion controls are discussed.

  17. Advanced electrostatic ion thruster for space propulsion

    NASA Technical Reports Server (NTRS)

    Masek, T. D.; Macpherson, D.; Gelon, W.; Kami, S.; Poeschel, R. L.; Ward, J. W.

    1978-01-01

    The suitability of the baseline 30 cm thruster for future space missions was examined. Preliminary design concepts for several advanced thrusters were developed to assess the potential practical difficulties of a new design. Useful methodologies were produced for assessing both planetary and earth orbit missions. Payload performance as a function of propulsion system technology level and cost sensitivity to propulsion system technology level are among the topics assessed. A 50 cm diameter thruster designed to operate with a beam voltage of about 2400 V is suggested to satisfy most of the requirements of future space missions.

  18. Simulation Based on Ion-Ion Plasma Techniques of Electric propulsion In Mars Mission Using Chlorine Gas

    NASA Astrophysics Data System (ADS)

    Sathiyavel, C.

    Abstract:The recently(Nov-5/2013) launched Mangalyan by the Indian space Research Organization (ISRO) to Mars orbit with Mankalyan contained by small liquid engine(MMH+N2O4).This will take long time to reach the Mars orbit that is around the 9 Months. Bi-Propellant rocket system has good thrust but low specific impulse and velocity. In future we need a rocket with good high specific impulse and high velocity of rocket system, to reduce the trip time to Mars. Electric propulsion rocket system is expected to become popular with the development of ion-ion pair techniques because this needs low propellant, Design thrust range is 1.5 N with high efficiency. An ion - ion pair of Electric propulsion rocket system is proposed in this work. Ion-Ion(positive ion- negative ion) Based Rocket system consists of three parts 1.The negative ionization stage with electro negative propellant 2. Ion-Ion plasma formation and ion accelerator 3. Exhaust of Nozzle. The Negative ions from electro negative gas are produced by adding up the gas, such as chlorine with electron emitted from an Electron gun ionization chamber. The formulate of large stable negative ion is achievable in chlorine gas with respect to electron affinity (∆E). When a neutral chlorine atom in the gaseous form picks up an electron to form a Cl- ion, it releases energy of 3.6eV. The negative ion density becomes several orders of magnitude larger than that of the electrons, hence forming ion-ion (positive ion - negative ion) plasma at the periphery of the discharge. The distance between ion- ions is important for the evaluate the rocket thrust and it also that the distance is determined by the exhaust velocity of the propellant. Accelerate the ion-ion plasma to a high velocity in the thrust vector direction via electron gun and the exhaust of ions through Nozzle. The simulation of the ion propulsion system has been carried out by MATLAB. By comparing the simulation results with the theoretical and previous results, we

  19. Heavy Ion Propulsion in the Megadalton Range

    DTIC Science & Technology

    2006-11-01

    atomizacidn electrostdtica, Universidad Carlos III, Madrid, Spain (2006) 15. D. Garoz, "Sintesis, estudio y mezclas de nuevos combustibles basados en...propellants for electrical propulsion from Taylor cones in vacuo), Proyecto fin de carrera (Senior Thesis), Universidad Politecnica de Madrid, Marzo 2004

  20. Performance and Controllability of Pulsed Ion Beam Ablation Propulsion

    SciTech Connect

    Yazawa, Masaru; Buttapeng, Chainarong; Harada, Nobuhiro; Suematsu, Hisayuki; Jiang Weihua; Yatsui, Kiyoshi

    2006-05-02

    We propose novel propulsion driven by ablation plasma pressures produced by the irradiation of pulsed ion beams onto a propellant. The ion beam ablation propulsion demonstrates by a thin foil (50 {mu}mt), and the flyer velocity of 7.7 km/s at the ion beam energy density of 2 kJ/cm2 adopted by using the Time-of-flight method is observed numerically and experimentally. We estimate the performance of the ion beam ablation propulsion as specific impulse of 3600 s and impulse bit density of 1700 Ns/m2 obtained from the demonstration results. In the numerical analysis, a one-dimensional hydrodynamic model with ion beam energy depositions is used. The control of the ion beam kinetic energy is only improvement of the performance but also propellant consumption. The spacecraft driven by the ion beam ablation provides high performance efficiency with short-pulsed ion beam irradiation. The numerical results of the advanced model explained latent heat and real gas equation of state agreed well with experimental ones over a wide range of the incident ion beam energy density.

  1. Electromagnetic interference assessment of an ion drive electric propulsion system

    NASA Technical Reports Server (NTRS)

    Whittlesey, A. C.

    1979-01-01

    The electromagnetic interference (EMI) form elements of an ion drive electric propulsion system was analyzed, and the effects of EMI interaction with a typical interplanetary spacecraft engineering and scientific subsystems were predicted. SEMCAP, a computerized electromagnetic compatibility assessment code, was used to analyze the impact of EMI noise sources on 65 engineering/telemetry circuits and 48 plasma wave and planetary radio astronomy channels measuring over the range of 100 Hz to 40 MHz in a spacecraft of the Voyager type; manual methods were used to evaluate electrostatics, magnetics, and communications effects. Results indicate that some conducted and radiated spectra are in excess of electromagnetic compatibility specification limits; direct design changes may be required for filtering and shielding of thrust system elements. The worst source of broadband radiated noise appears to be the power processor. The magnetic field necessary to thruster operation is equivalent to about 18 amp-sq m per amp of beam current at right angles to the axis caused by the neutralizer/plume loop.

  2. Electromagnetic interference assessment of an ion drive electric propulsion system

    NASA Technical Reports Server (NTRS)

    Whittlesey, A. C.

    1979-01-01

    The electromagnetic interference (EMI) form elements of an ion drive electric propulsion system was analyzed, and the effects of EMI interaction with a typical interplanetary spacecraft engineering and scientific subsystems were predicted. SEMCAP, a computerized electromagnetic compatibility assessment code, was used to analyze the impact of EMI noise sources on 65 engineering/telemetry circuits and 48 plasma wave and planetary radio astronomy channels measuring over the range of 100 Hz to 40 MHz in a spacecraft of the Voyager type; manual methods were used to evaluate electrostatics, magnetics, and communications effects. Results indicate that some conducted and radiated spectra are in excess of electromagnetic compatibility specification limits; direct design changes may be required for filtering and shielding of thrust system elements. The worst source of broadband radiated noise appears to be the power processor. The magnetic field necessary to thruster operation is equivalent to about 18 amp-sq m per amp of beam current at right angles to the axis caused by the neutralizer/plume loop.

  3. The Ion Propulsion System for the Asteroid Redirect Robotic Mission

    NASA Technical Reports Server (NTRS)

    Herman, Daniel A.; Santiago, Walter; Kamhawi, Hani; Polk, James E.; Snyder, John Steven; Hofer, Richard; Sekerak, Michael

    2016-01-01

    The Asteroid Redirect Robotic Mission is a Solar Electric Propulsion Technology Demonstration Mission (ARRM) whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a companion human-crewed mission. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as a critical part of NASA's future beyond-low-Earth-orbit, human-crewed exploration plans. This presentation presents the conceptual design of the ARRM ion propulsion system, the status of the NASA in-house thruster and power processing development activities, the status of the planned technology maturation for the mission through flight hardware delivery, and the status of the mission formulation and spacecraft acquisition.

  4. Simulation Based on Ion Propulsion Rocket System with Using Negative ion - Negative Ion Pair Techniques

    NASA Astrophysics Data System (ADS)

    Sathiyavel, C.

    2016-07-01

    Ion propulsion rocket system is expected to become popular with the development of ion-ion pair techniques because of their stimulated of low propellant, Design of Thrust range is 1N with low electric power and high efficiency. A Negative ion-Negative ion pair of ion propulsion rocket system is proposed in this work .Negative Ion Based Rocket system consists of three parts 1.ionization chamber 2. Repulsion force and ion accelerator 3. Exhaust of Nozzle. The Negative ions from electro negatively gas are produced by attachment of the gas ,such as chlorine with electron emitted from a Electron gun ionization chamber. The formulate of large stable negative ion is achievable in chlorine gas with respect to electron affinity (∆E). The electron affinity is a measure of the energy change when an electron is added to a neutral atom to form a negative ion. When a neutral chlorine atom in the gaseous form picks up an electron to form a Cl- ion, it releases energy of 349 kJ/mol or 3.6 ev/atom. It is said to have an electron affinity of -349 kJ/mol ,the negative sign indicating that energy is released during this process .The mechanisms of attachment involve the formation of intermediate states. In that reason for , the highly repulsive force created between the same negative ions. The distance between same negative ions is important for the evaluate of the rocket thrust and is also determined by the exhaust velocity of the propellant. The mass flow rate of propellant is achieved by the ratio of total mass of the propellant (Kg) needed for operation to time period(s). Accelerate the Negative ions to a high velocity in the thrust vector direction with a significantly intense Magnetic field and the exhaust of negative ions through Nozzle. The simulation of the ion propulsion system has been carried out by MATLAB. By comparing the simulation results with the theoretical and previous results, we have found that the proposed method is achieved of thrust value with estimated

  5. General Aviation Propulsion (GAP) Program, Turbine Engine System Element

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The goal of the General Aviation Propulsion (GAP) Program Turbine Engine System Elements is to conduct a shared resource project to develop an affordable gas turbine engine for use on 4 to 6 place, light aircraft that will lead to revitalization of the general aviation industry in the United States, creating many new, high-quality jobs.

  6. Nuclear electric propulsion mission engineering study. Volume 2: Final report

    NASA Technical Reports Server (NTRS)

    1973-01-01

    Results of a mission engineering analysis of nuclear-thermionic electric propulsion spacecraft for unmanned interplanetary and geocentric missions are summarized. Critical technologies associated with the development of nuclear electric propulsion (NEP) are assessed, along with the impact of its availability on future space programs. Outer planet and comet rendezvous mission analysis, NEP stage design for geocentric and interplanetary missions, NEP system development cost and unit costs, and technology requirements for NEP stage development are studied.

  7. CVD Rhenium Engines for Solar-Thermal Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Williams, Brian E.; Fortini, Arthur J.; Tuffias, Robert H.; Duffy, Andrew J.; Tucker, Stephen P.

    1999-01-01

    Solar-thermal upper-stage propulsion systems have the potential to provide specific impulse approaching 900 seconds, with 760 seconds already demonstrated in ground testing. Such performance levels offer a 100% increase in payload capability compared to state-of-the-art chemical upper-stage systems, at lower cost. Although alternatives such as electric propulsion offer even greater performance, the 6- to 18- month orbital transfer time is a far greater deviation from the state of the art than the one to two months required for solar propulsion. Rhenium metal is the only material that is capable of withstanding the predicted thermal, mechanical, and chemical environment of a solar-thermal propulsion device. Chemical vapor deposition (CVD) is the most well-established and cost-effective process for the fabrication of complex rhenium structures. CVD rhenium engines have been successfully constructed for the Air Force ISUS program (bimodal thrust/electricity) and the NASA Shooting Star program (thrust only), as well as under an Air Force SBIR project (thrust only). The bimodal engine represents a more long-term and versatile approach to solar-thermal propulsion, while the thrust-only engines provide a potentially lower weight/lower cost and more near-term replacement for current upper-stage propulsion systems.

  8. CVD Rhenium Engines for Solar-Thermal Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Williams, Brian E.; Fortini, Arthur J.; Tuffias, Robert H.; Duffy, Andrew J.; Tucker, Stephen P.

    1999-01-01

    Solar-thermal upper-stage propulsion systems have the potential to provide specific impulse approaching 900 seconds, with 760 seconds already demonstrated in ground testing. Such performance levels offer a 100% increase in payload capability compared to state-of-the-art chemical upper-stage systems, at lower cost. Although alternatives such as electric propulsion offer even greater performance, the 6- to 18- month orbital transfer time is a far greater deviation from the state of the art than the one to two months required for solar propulsion. Rhenium metal is the only material that is capable of withstanding the predicted thermal, mechanical, and chemical environment of a solar-thermal propulsion device. Chemical vapor deposition (CVD) is the most well-established and cost-effective process for the fabrication of complex rhenium structures. CVD rhenium engines have been successfully constructed for the Air Force ISUS program (bimodal thrust/electricity) and the NASA Shooting Star program (thrust only), as well as under an Air Force SBIR project (thrust only). The bimodal engine represents a more long-term and versatile approach to solar-thermal propulsion, while the thrust-only engines provide a potentially lower weight/lower cost and more near-term replacement for current upper-stage propulsion systems.

  9. The stirling engine for vehicle propulsion

    NASA Technical Reports Server (NTRS)

    Kuhlman, P.

    1978-01-01

    The performance data of experimental Stirling engines are considered along with questions of exhaust-gas composition, engine noise, engine volume and weight, engine control, and the engine-starting process. The Stirling engine can use practically any liquid or gaseous fuel for its operation. It is found that technically a use of the Stirling engine in motor vehicles is feasible. Economic questions related to an introduction of the Stirling engine are discussed along with possible new developments which could improve the economic situation in favor of a use of Stirling engine.

  10. SMART-1 ion engine fired successfully

    NASA Astrophysics Data System (ADS)

    2003-10-01

    Close-up view of SMART-1's stationary plasma thruster hi-res Size hi-res: 257 kb Credits: ESA 2002. Illustration by Medialab. Close-up view of SMART-1's stationary plasma thruster Electrons attracted into the discharge chamber collide with xenon atoms from the propellant gas supply, making charged atoms (ions). Current-carrying coils, inside and outside the doughnut-shaped discharge chamber, sustain a magnetic field oriented like the spokes of a wheel. By the Hall effect, ions and electrons swerving in opposite directions in the magnetic field create an electric field. This expels the xenon ions in a propulsive jet. Other emitted electrons then neutralize the xenon, producing the blue jet. Engineers at ESOC, the European Space Agency's control centre in Darmstadt, Germany, sent a command to begin the firing test, which lasted for one hour. This was similar to a trial performed on Earth before SMART-1 was launched. Several months ago, the ion engine, or Solar Electric Primary Propulsion (SEPP) system, had been placed in a vacuum chamber on the ground and its functions and operation were measured. Now in space and in a true vacuum, the ion engine actually worked better than in the test on ground and has nudged SMART-1 a little closer to the Moon. This is the first time that Europe flies an electric primary propulsion in space, and also the first European use of this particular type of ion engine, called a 'Hall-effect' thruster. The SEPP consists of a single ion engine fuelled by xenon gas and powered by solar energy. The ion engine will accelerate SMART-1 very gradually to cause the spacecraft to travel in a series of spiralling orbits - each revolution slightly further away from the Earth - towards the Moon. Once captured by the Moon's gravity, SMART-1 will move into ever-closer orbits of the Moon. As part of one of the overall mission objectives to test this new SEPP technology, the data will now be analysed to see how much acceleration was achieved and how

  11. NASA's Evolutionary Xenon Thruster (NEXT) Ion Propulsion System Information Summary

    NASA Technical Reports Server (NTRS)

    Pencil, Eirc S.; Benson, Scott W.

    2008-01-01

    This document is a guide to New Frontiers mission proposal teams. The document describes the development and status of the NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system (IPS) technology, its application to planetary missions, and the process anticipated to transition NEXT to the first flight mission.

  12. Engine Propeller Research Building at the Lewis Flight Propulsion Laboratory

    NASA Image and Video Library

    1955-02-21

    The Engine Propeller Research Building, referred to as the Prop House, emits steam from its acoustic silencers at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. In 1942 the Prop House became the first completed test facility at the new NACA laboratory in Cleveland, Ohio. It contained four test cells designed to study large reciprocating engines. After World War II, the facility was modified to study turbojet engines. Two of the test cells were divided into smaller test chambers, resulting in a total of six engine stands. During this period the NACA Lewis Materials and Thermodynamics Division used four of the test cells to investigate jet engines constructed with alloys and other high temperature materials. The researchers operated the engines at higher temperatures to study stress, fatigue, rupture, and thermal shock. The Compressor and Turbine Division utilized another test cell to study a NACA-designed compressor installed on a full-scale engine. This design sought to increase engine thrust by increasing its airflow capacity. The higher stage pressure ratio resulted in a reduction of the number of required compressor stages. The last test cell was used at the time by the Engine Research Division to study the effect of high inlet densities on a jet engine. Within a couple years of this photograph the Prop House was significantly altered again. By 1960 the facility was renamed the Electric Propulsion Research Building to better describe its new role in electric propulsion.

  13. ION ROCKET ENGINE

    DOEpatents

    Ehlers, K.W.; Voelker, F. III

    1961-12-19

    A thrust generating engine utilizing cesium vapor as the propellant fuel is designed. The cesium is vaporized by heat and is passed through a heated porous tungsten electrode whereby each cesium atom is fonized. Upon emergfng from the tungsten electrode, the ions are accelerated rearwardly from the rocket through an electric field between the tungsten electrode and an adjacent accelerating electrode grid structure. To avoid creating a large negative charge on the space craft as a result of the expulsion of the positive ions, a source of electrons is disposed adjacent the ion stream to neutralize the cesium atoms following acceleration thereof. (AEC)

  14. Single ion heat engine

    NASA Astrophysics Data System (ADS)

    Singer, Kilian

    2015-03-01

    An experimental realization of a heat engine with a single ion is presented, which will allow for work extraction even with non-classical thermal reservoirs. To this goal a custom designed linear Paul trap with a single ion performing an Otto cycle is presented. The radial state of the ion is used as the working gas analogous to the gas in a conventional heat engine. The conventional piston is realized by the axial degrees of freedom and the axial motional excitation stores the generated work, just like a conventional fly-wheel. The heat baths can be realized by tailored laser radiation. Alternatively electrical noise can be used to control the state of the ion. The presented system possesses advantageous properties, as the working parameters can be tuned over a broad range and the motional degrees of freedom of the ion can be accurately determined. Dark resonances allow for fast stroboscopic thermometry during the entire working cycle. Monte Carlo simulations are performed to predict the efficiency and the gained work of the working cycle. We have also shown how the equations for the Carnot limit have to be modified if a squeezed thermal reservoir is employed. Furthermore structural phase transitions with laser cooled linear ion crystals are induced verifying the Kibble-Zurek mechanism.

  15. Propulsion engineering study for small-scale Mars missions

    SciTech Connect

    Whitehead, J.

    1995-09-12

    Rocket propulsion options for small-scale Mars missions are presented and compared, particularly for the terminal landing maneuver and for sample return. Mars landing has a low propulsive {Delta}v requirement on a {approximately}1-minute time scale, but at a high acceleration. High thrust/weight liquid rocket technologies, or advanced pulse-capable solids, developed during the past decade for missile defense, are therefore more appropriate for small Mars landers than are conventional space propulsion technologies. The advanced liquid systems are characterize by compact lightweight thrusters having high chamber pressures and short lifetimes. Blowdown or regulated pressure-fed operation can satisfy the Mars landing requirement, but hardware mass can be reduced by using pumps. Aggressive terminal landing propulsion designs can enable post-landing hop maneuvers for some surface mobility. The Mars sample return mission requires a small high performance launcher having either solid motors or miniature pump-fed engines. Terminal propulsion for 100 kg Mars landers is within the realm of flight-proven thruster designs, but custom tankage is desirable. Landers on a 10 kg scale also are feasible, using technology that has been demonstrated but not previously flown in space. The number of sources and the selection of components are extremely limited on this smallest scale, so some customized hardware is required. A key characteristic of kilogram-scale propulsion is that gas jets are much lighter than liquid thrusters for reaction control. The mass and volume of tanks for inert gas can be eliminated by systems which generate gas as needed from a liquid or a solid, but these have virtually no space flight history. Mars return propulsion is a major engineering challenge; earth launch is the only previously-solved propulsion problem requiring similar or greater performance.

  16. A Synopsis of Ion Propulsion Development Projects in the United States: SERT 1 to Deep Space I

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Rawlin, Vincent K.; Patterson, Michael J.

    1999-01-01

    The historical background and characteristics of the experimental flights of ion propulsion systems and the major ground-based technology demonstrations were reviewed. The results of the first successful ion engine flight in 1964, SERT I which demonstrated ion beam neutralization, are discussed along with the extended operation of SERT II starting in 1970. These results together with the technology employed on the early cesium engine flights. the Applications Technology Satellite (ATS) series, and the ground-test demonstrations, have provided the evolutionary path for the development of xenon ion thruster component technologies, control systems, and power circuit implementations. In the 1997-1999 period, the communication satellite flights using ion engine systems and the Deep Space I flight confirmed that these auxiliary and primary propulsion systems have advanced to a high-level of flight-readiness.

  17. An engineering model 30 cm ion thruster

    NASA Technical Reports Server (NTRS)

    Poeschel, R. L.; King, H. J.; Schnelker, D. E.

    1973-01-01

    Thruster development at Hughes Research Laboratories and NASA Lewis Research Center has brought the 30-cm mercury bombardment ion thruster to the state of an engineering model. This thruster has been designed to have sufficient internal strength for direct mounting on gimbals, to weigh 7.3 kg, to operate with a corrected overall efficiency of 71%, and to have 10,000 hours lifetime. Subassemblies, such as the ion optical system, isolators, etc., have been upgraded to meet launch qualification standards. This paper presents a summary of the design specifications and performance characteristics which define the interface between the thruster module and the remainder of the propulsion system.

  18. Development of arcjet and ion propulsion for spacecraft stationkeeping

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Curran, Francis M.; Haag, Thomas W.; Patterson, Michael J.; Pencil, Eric J.; Rawlin, Vincent K.; Sankovic, John M.

    1992-01-01

    Near term flight applications of arc jet and ion thruster satellite station-keeping systems as well as development activities in Europe, Japan, and the United States are reviewed. At least two arc jet and three ion propulsion flights are scheduled during the 1992-1995 period. Ground demonstration technology programs are focusing on the development of kW-class hydrazine and ammonia arc jets and xenon ion thrusters. Recent work at NASA LeRC on electric thruster and system integration technologies relating to satellite station keeping and repositioning will also be summarized.

  19. Development of arcjet and ion propulsion for spacecraft stationkeeping

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Curran, Francis M.; Haag, Thomas W.; Patterson, Michael J.; Pencil, Eric J.; Rawlin, Vincent K.; Sankovic, John M.

    1992-01-01

    Near term flight applications of arcjet and ion thruster satellite station-keeping systems as well as development activities in Europe, Japan, and the United States are reviewed. At least two arcjet and three ion propulsion flights are scheduled during the 1992 - 1995 period. Ground demonstration technology programs are focusing on the development of kW-class hydrazine and ammonia arcjets and xenon ion thrusters. Recent work at NASA Lewis Research Center on electric thruster and system integration technologies relating to satellite stationkeeping and repositioning will also be summarized.

  20. Nuclear powered Mars cargo transport mission utilizing advanced ion propulsion

    SciTech Connect

    Galecki, D.L.; Patterson, M.J.

    1987-01-01

    Nuclear-powered ion propulsion technology was combined with detailed trajectory analysis to determine propulsion system and trajectory options for an unmanned cargo mission to Mars in support of manned Mars missions. A total of 96 mission scenarios were identified by combining two power levels, two propellants, four values of specific impulse per propellant, three starting altitudes, and two starting velocities. Sixty of these scenarios were selected for a detailed trajectory analysis; a complete propulsion system study was then conducted for 20 of these trajectories. Trip times ranged from 344 days for a xenon propulsion system operating at 300 kW total power and starting from lunar orbit with escape velocity, to 770 days for an argon propulsion system operating at 300 kW total power and starting from nuclear start orbit with circular velocity. Trip times for the 3 MW cases studied ranged from 356 to 413 days. Payload masses ranged from 5700 to 12,300 kg for the 300 kW power level, and from 72,200 to 81,500 kg for the 3 MW power level.

  1. Annular Ion Engine Concept and Development Status

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.

    2016-01-01

    The Annular Ion Engine (AIE) concept represents an evolutionary development in gridded ion thruster technology with the potential for delivering revolutionary capabilities. It has this potential because the AIE concept: (a) enables scaling of ion thruster technology to high power at specific impulse (Isp) values of interest for near-term mission applications, 5000 sec; and (b) it enables an increase in both thrust density and thrust-to-power (FP) ratio exceeding conventional ion thrusters and other electric propulsion (EP) technology options, thereby yielding the highest performance over a broad range in Isp. The AIE concept represents a natural progression of gridded ion thruster technology beyond the capabilities embodied by NASAs Evolutionary Xenon Thruster (NEXT) [1]. The AIE would be appropriate for: (a) applications which require power levels exceeding NEXTs capabilities (up to about 14 kW [2]), with scalability potentially to 100s of kW; and/or (b) applications which require FP conditions exceeding NEXTs capabilities.

  2. Plasma Propulsion Testing Capabilities at Arnold Engineering Development Center

    NASA Technical Reports Server (NTRS)

    Polzin, Kurt A.; Dawbarn, Albert; Moeller, Trevor

    2007-01-01

    This paper describes the results of a series of experiments aimed at quantifying the plasma propulsion testing capabilities of a 12-ft diameter vacuum facility (12V) at USAF-Arnold Engineering Development Center (AEDC). Vacuum is maintained in the 12V facility by cryogenic panels lining the interior of the chamber. The pumping capability of these panels was shown to be great enough to support plasma thrusters operating at input electrical power >20 kW. In addition, a series of plasma diagnostics inside the chamber allowed for measurement of plasma parameters at different spatial locations, providing information regarding the chamber's effect on the global plasma thruster flowfield. The plasma source used in this experiment was Hall thruster manufactured by Busek Co. The thruster was operated at up to 20 kW steady-state power in both a lower current and higher current mode. The vacuum level in the chamber never rose above 9 x 10(exp -6) torr during the course of testing. Langmuir probes, ion flux probes, and Faraday cups were used to quantify the plasma parameters in the chamber. We present the results of these measurements and estimates of pumping speed based on the background pressure level and thruster propellant mass flow rate.

  3. Transport Processes in Beamed Energy Propulsion Systems

    DTIC Science & Technology

    1991-11-01

    the advanced propulsion systems currently being considered for future space missions are the resistojet, arcjet , ion engines, and laser and microwave... electrothermal propulsion systems. Studies conducted of advanced propulsion concepts for both NASA and Air Force missions, such as low-earth orbit...advanced propulsion concepts mentioned, microwave electrothermal propulsion systems best suit this range of operation. Resistojets employ electrothermal

  4. Structural Requirements for the Space Propulsion Engine Systems

    NASA Technical Reports Server (NTRS)

    Aggarwal, Pravin K.

    2006-01-01

    In January 2004, the National Aeronautics and Space Administration (NASA) was given a vision for Space Exploration by President Bush, setting our sight on a bold new path to go back to the Moon, then to Mars and beyond. As NASA gets ready to meet the vision set by President Bush, failures are not an option. Reliability of the propulsion engine systems will play an important role in establishing an overall safe and reliable operation of these new space systems. A new standard, NASA-STD-5012, Strength and Life Assessment for Space Propulsion System Engines, has been developed to provide structural requirements for assessment of the propulsion systems engine. This standard is a complement to the current NASA-wide standard NASA-STD-5001, Structural Design and Test Factors of Safety for Spaceflight Hardware, which excluded the requirement for the engine systems (rotatory structures) along with pressure vessels. As developed, this document builds on the heritage of the multiple industrial standards related to strength and life assessment of the structures. For assuring a safe and reliable operation of a product and/or mission, establishing a set of structural assessment requirements is a key ingredient. Hence, a concentrated effort was made to improve the requirements where there are known lessons learned during the design, test, and operation phases of the Space Shuttle Main Engine (SSME) and other engine development programs. Requirements delineated in this standard are also applicable for the reusable and/or human missions. It shall be noted that "reliability of a system cannot be tested and inspected but can only be achieved if it is first designed into a system." Hence, these strength and life assessment requirements for the space propulsion system engines shall be used along with other good engineering practices, requirements, and policies.

  5. Effect of Correlation on Multi-Engine Rocket Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Baker, J. R.; Breneman, J. E.

    1989-01-01

    A matter of great concern in the design and operation of multi-engine rocket propulsion systems is the effect of the premature shutdown of one engine on the vehicle. This probability that a premature shutdown will cause a vehicle loss is termed correlation. Based on airbreathing experiences as well as rocket engine data the best estimate of this correlation is made and then applied to the overall multi-engine reliability problem to demonstrate its potential effect. At this point, follow-on analyses are pointed out that illustrate how any potential failures that may cause a correlatable event can be eliminated; thus bringing this correlation to almost 0.

  6. Nuclear electric propulsion mission engineering study. Volume 1: Executive summary

    NASA Technical Reports Server (NTRS)

    1973-01-01

    Results of a mission engineering analysis of nuclear-thermionic electric propulsion spacecraft for unmanned interplanetary and geocentric missions are summarized. Critical technologies associated with the development of nuclear electric propulsion (NEP) are assessed. Outer planet and comet rendezvous mission analysis, NEP stage design for geocentric and interplanetary missions, NEP system development cost and unit costs, and technology requirements for NEP stage development are studied. The NEP stage design provides both inherent reliability and high payload mass capability. The NEP stage and payload integration was found to be compatible with the space shuttle.

  7. Propulsion Controls, 1979. [air breathing engine control

    NASA Technical Reports Server (NTRS)

    1980-01-01

    The state of the art of multivariable engine control is examined in order to determine future needs and problem areas and to establish the appropriate roles of government, industries, and universities in addressing these problems.

  8. Deep Space 1's ion engine

    NASA Technical Reports Server (NTRS)

    2002-01-01

    Kennedy Space Center, Florida. - Deep Space 1 is lifted from its work platform, giving a closeup view of the experimental solar-powered ion propulsion engine. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century. Another onboard experiment includes software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, Cape Canaveral Air Station, in October. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches.

  9. Hypersonic propulsion. [supersonic combustion ramjet engines

    NASA Technical Reports Server (NTRS)

    Beach, H. L., Jr.

    1979-01-01

    Research on hydrogen fueled scramjet engines for hypersonic flight is reviewed. Component developments, computational methods, and preliminary ground tests of subscale scramjet engine modules at Mach 4 and 7 are emphasized. Airframe integration, structures, and flow diagnostics are also discussed. It is shown that mixed-mode perpendicular and parallel fuel injection controls heat release over a wide Mach range and the fixed geometry inlet gives good performance over a wide range of Mach numbers.

  10. The High Power Electric Propulsion (HiPEP) Ion Thruster

    NASA Technical Reports Server (NTRS)

    Foster, John E.; Haag, Tom; Patterson, Michael; Williams, George J., Jr.; Sovey, James S.; Carpenter, Christian; Kamhawi, Hani; Malone, Shane; Elliot, Fred

    2004-01-01

    Practical implementation of the proposed Jupiter Icy Moon Orbiter (JIMO) mission, which would require a total delta V of approximately 38 km/s, will require the development of a high power, high specific impulse propulsion system. Initial analyses show that high power gridded ion thrusters could satisfy JIMO mission requirements. A NASA GRC-led team is developing a large area, high specific impulse, nominally 25 kW ion thruster to satisfy both the performance and the lifetime requirements for this proposed mission. The design philosophy and development status as well as a thruster performance assessment are presented.

  11. Combined Experimental and Numerical Investigations into Laser Propulsion Engineering Physics

    NASA Astrophysics Data System (ADS)

    Kenoyer, David Adam

    The RPI pulsed Laser Propulsion (LP) research effort focuses on the future application of launching nano- and micro-satellites (1-10 kg payloads) into Low Earth Orbit (LEO), using a remote Ground Based Laser (GBL) power station to supply the required energy for flight. This research program includes both experimental and numerical studies investigating the propulsive performance of several engine geometries (constituting a lightcraft family). Using the Lumonics twin K-922m TEA pulsed laser system, axial and lateral thrust, C m, Isp, and η measurements were made for these engine geometries, examining the effects of several critical factors including: engine orientation (e.g. lateral and angular offset), laser pulse energy, pulse repetition frequency, pulse duration, propellant type, and engine size-scaling effects. Investigation into the origins of lateral "beam riding" forces was of particular interest. Lateral impulse measurements and high speed Schlieren photography were utilized to provide an understanding of laser beam-riding/propulsive physics. The acquired lightcraft database was used to further develop an existing 7-Degree Of Freedom (DOF) flight dynamics model extensively calibrated against 16 actual trajectories of small scale model lightcraft flown at White Sands Missile Range, NM on a 10 kW pulsed CO2 laser called PLVTS. The full system 7-DOF model is comprised of updated individual aerodynamics, engine, laser beam propagation, variable vehicle inertia, reaction controls system, and dynamics models, integrated to represent all major phenomena in a consistent framework. This flight dynamics model and associated 7-DOF code provide a physics-based predictive tool for basic research investigations into laser launched lightcraft for suborbital and orbital missions. Simulations were performed to demonstrate the flight capabilities of each engine geometry using the updated lightcraft propulsion database, the results of which further demonstrate that autonomous

  12. Investigation of Exoskeletal Engine Propulsion System Concept

    NASA Technical Reports Server (NTRS)

    Roche, Joseph M.; Palac, Donald T.; Hunter, James E.; Myers, David E.; Snyder, Christopher A.; Kosareo, Daniel N.; McCurdy, David R.; Dougherty, Kevin T.

    2005-01-01

    An innovative approach to gas turbine design involves mounting compressor and turbine blades to an outer rotating shell. Designated the exoskeletal engine, compression (preferable to tension for high-temperature ceramic materials, generally) becomes the dominant blade force. Exoskeletal engine feasibility lies in the structural and mechanical design (as opposed to cycle or aerothermodynamic design), so this study focused on the development and assessment of a structural-mechanical exoskeletal concept using the Rolls-Royce AE3007 regional airliner all-axial turbofan as a baseline. The effort was further limited to the definition of an exoskeletal high-pressure spool concept, where the major structural and thermal challenges are represented. The mass of the high-pressure spool was calculated and compared with the mass of AE3007 engine components. It was found that the exoskeletal engine rotating components can be significantly lighter than the rotating components of a conventional engine. However, bearing technology development is required, since the mass of existing bearing systems would exceed rotating machinery mass savings. It is recommended that once bearing technology is sufficiently advanced, a "clean sheet" preliminary design of an exoskeletal system be accomplished to better quantify the potential for the exoskeletal concept to deliver benefits in mass, structural efficiency, and cycle design flexibility.

  13. Propulsion Controls Modeling for a Small Turbofan Engine

    NASA Technical Reports Server (NTRS)

    Connolly, Joseph W.; Csank, Jeffrey T.; Chicatelli, Amy; Franco, Kevin

    2017-01-01

    A nonlinear dynamic model and propulsion controller are developed for a small-scale turbofan engine. The small-scale turbofan engine is based on the Price Induction company's DGEN 380, one of the few turbofan engines targeted for the personal light jet category. Comparisons of the nonlinear dynamic turbofan engine model to actual DGEN 380 engine test data and a Price Induction simulation are provided. During engine transients, the nonlinear model typically agrees within 10 percent error, even though the nonlinear model was developed from limited available engine data. A gain scheduled proportional integral low speed shaft controller with limiter safety logic is created to replicate the baseline DGEN 380 controller. The new controller provides desired gain and phase margins and is verified to meet Federal Aviation Administration transient propulsion system requirements. In understanding benefits, there is a need to move beyond simulation for the demonstration of advanced control architectures and technologies by using real-time systems and hardware. The small-scale DGEN 380 provides a cost effective means to accomplish advanced controls testing on a relevant turbofan engine platform.

  14. Computational simulation for concurrent engineering of aerospace propulsion systems

    NASA Technical Reports Server (NTRS)

    Chamis, C. C.; Singhal, S. N.

    1993-01-01

    Results are summarized for an investigation to assess the infrastructure available and the technology readiness in order to develop computational simulation methods/software for concurrent engineering. These results demonstrate that development of computational simulation methods for concurrent engineering is timely. Extensive infrastructure, in terms of multi-discipline simulation, component-specific simulation, system simulators, fabrication process simulation, and simulation of uncertainties--fundamental to develop such methods, is available. An approach is recommended which can be used to develop computational simulation methods for concurrent engineering of propulsion systems and systems in general. Benefits and issues needing early attention in the development are outlined.

  15. Computational simulation of concurrent engineering for aerospace propulsion systems

    NASA Technical Reports Server (NTRS)

    Chamis, C. C.; Singhal, S. N.

    1992-01-01

    Results are summarized of an investigation to assess the infrastructure available and the technology readiness in order to develop computational simulation methods/software for concurrent engineering. These results demonstrate that development of computational simulations methods for concurrent engineering is timely. Extensive infrastructure, in terms of multi-discipline simulation, component-specific simulation, system simulators, fabrication process simulation, and simulation of uncertainties - fundamental in developing such methods, is available. An approach is recommended which can be used to develop computational simulation methods for concurrent engineering for propulsion systems and systems in general. Benefits and facets needing early attention in the development are outlined.

  16. Computational simulation of concurrent engineering for aerospace propulsion systems

    NASA Technical Reports Server (NTRS)

    Chamis, C. C.; Singhal, S. N.

    1992-01-01

    Results are summarized of an investigation to assess the infrastructure available and the technology readiness in order to develop computational simulation methods/software for concurrent engineering. These results demonstrate that development of computational simulations methods for concurrent engineering is timely. Extensive infrastructure, in terms of multi-discipline simulation, component-specific simulation, system simulators, fabrication process simulation, and simulation of uncertainties - fundamental in developing such methods, is available. An approach is recommended which can be used to develop computational simulation methods for concurrent engineering for propulsion systems and systems in general. Benefits and facets needing early attention in the development are outlined.

  17. Uprated OMS engine for upper stage propulsion

    NASA Technical Reports Server (NTRS)

    Boyd, William C.

    1986-01-01

    The results of a pre-development component demonstration program on the use of a gas generator-driven turbopump that increases the Space Shuttle's Orbital Maneuvering Engine (OME) operating pressure are given. Tests and analysis confirm the the capability of the concept to meet or exceed performance and life requirements. Storable propellant upper stage concepts are also discussed.

  18. Energy Efficient Engine: Flight propulsion system final design and analysis

    NASA Technical Reports Server (NTRS)

    Davis, Donald Y.; Stearns, E. Marshall

    1985-01-01

    The Energy Efficient Engine (E3) is a NASA program to create fuel saving technology for future transport engines. The Flight Propulsion System (FPS) is the engine designed to achieve E3 goals. Achieving these goals required aerodynamic, mechanical and system technologies advanced beyond that of current production engines. These technologies were successfully demonstrated in component rigs, a core engine and a turbofan ground test engine. The design and benefits of the FPS are presented. All goals for efficiency, environmental considerations, and economic payoff were met. The FPS has, at maximum cruise, 10.67 km (35,000 ft), M0.8, standard day, a 16.9 percent lower installed specific fuel consumption than a CF6-50C. It provides an 8.6 percent reduction in direct operating cost for a short haul domestic transport and a 16.2 percent reduction for an international long distance transport.

  19. High energy density propulsion systems and small engine dynamometer

    NASA Astrophysics Data System (ADS)

    Hays, Thomas

    2009-07-01

    Scope and Method of Study. This study investigates all possible methods of powering small unmanned vehicles, provides reasoning for the propulsion system down select, and covers in detail the design and production of a dynamometer to confirm theoretical energy density calculations for small engines. Initial energy density calculations are based upon manufacturer data, pressure vessel theory, and ideal thermodynamic cycle efficiencies. Engine tests are conducted with a braking type dynamometer for constant load energy density tests, and show true energy densities in excess of 1400 WH/lb of fuel. Findings and Conclusions. Theory predicts lithium polymer, the present unmanned system energy storage device of choice, to have much lower energy densities than other conversion energy sources. Small engines designed for efficiency, instead of maximum power, would provide the most advantageous method for powering small unmanned vehicles because these engines have widely variable power output, loss of mass during flight, and generate rotational power directly. Theoretical predictions for the energy density of small engines has been verified through testing. Tested values up to 1400 WH/lb can be seen under proper operating conditions. The implementation of such a high energy density system will require a significant amount of follow-on design work to enable the engines to tolerate the higher temperatures of lean operation. Suggestions are proposed to enable a reliable, small-engine propulsion system in future work. Performance calculations show that a mature system is capable of month long flight times, and unrefueled circumnavigation of the globe.

  20. Low-Power Ion Propulsion for Small Spacecraft

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Oleson, Steven R.

    1997-01-01

    Analyses were conducted which indicate that sub kW-class ion thrusters may provide performance benefits for near-Earth space commercial and science missions. Small spacecraft applications with masses ranging from 50 to 500 kg and power levels less than 0.5 kW were considered. To demonstrate the efficacy of propulsion systems of this class, two potential missions were chosen as examples; a geosynchronous north-south station keeping application, and an Earth orbit magnetospheric mapping satellite constellation. Xenon ion propulsion system solutions using small thrusters were evaluated for these missions. A payload mass increase of more than 15% is provided by a 300-W ion system for the north-south station keeping mission. A launch vehicle reduction from four to one results from using the ion thruster for the magnetospheric mapping mission. Typical projected thruster performance over the input power envelope of 100-300 W range from approximately 40% to 54% efficiency and approximately 2000 to 3000 seconds specific impulse. Thruster technologies required to achieve the mission-required performance and lifetime are identified.

  1. Hydrodynamic Efficiency of Ablation Propulsion with Pulsed Ion Beam

    SciTech Connect

    Buttapeng, Chainarong; Yazawa, Masaru; Harada, Nobuhiro; Suematsu, Hisayuki; Jiang Weihua; Yatsui, Kiyoshi

    2006-05-02

    This paper presents the hydrodynamic efficiency of ablation plasma produced by pulsed ion beam on the basis of the ion beam-target interaction. We used a one-dimensional hydrodynamic fluid compressible to study the physics involved namely an ablation acceleration behavior and analyzed it as a rocketlike model in order to investigate its hydrodynamic variables for propulsion applications. These variables were estimated by the concept of ablation driven implosion in terms of ablated mass fraction, implosion efficiency, and hydrodynamic energy conversion. Herein, the energy conversion efficiency of 17.5% was achieved. In addition, the results show maximum energy efficiency of the ablation process (ablation efficiency) of 67% meaning the efficiency with which pulsed ion beam energy-ablation plasma conversion. The effects of ion beam energy deposition depth to hydrodynamic efficiency were briefly discussed. Further, an evaluation of propulsive force with high specific impulse of 4000s, total impulse of 34mN and momentum to energy ratio in the range of {mu}N/W was also analyzed.

  2. Engines and innovation: Lewis Laboratory and American propulsion technology

    NASA Technical Reports Server (NTRS)

    Dawson, Virginia Parker

    1991-01-01

    This book is an institutional history of the NASA Lewis Research Center, located in Cleveland, Ohio, from 1940, when Congress authorized funding for a third laboratory for the National Advisory Committee for Aeronautics, through the 1980s. The history of the laboratory is discussed in relation to the development of American propulsion technology, with particular focus on the transition in the 1940s from the use of piston engines in airplanes to jet propulsion and that from air-breathing engines to rocket technology when the National Aeronautics and Space Administration was established in 1958. The personalities and research philosophies of the people who shaped the history of the laboratory are discussed, as is the relationship of Lewis Research Center to the Case Institute of Technology.

  3. 40 CFR Appendix II to Part 1045 - Duty Cycles for Propulsion Marine Engines

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 40 Protection of Environment 33 2014-07-01 2014-07-01 false Duty Cycles for Propulsion Marine...) AIR POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Pt. 1045, App. II Appendix II to Part 1045—Duty Cycles for Propulsion Marine Engines (a)...

  4. 40 CFR Appendix II to Part 1045 - Duty Cycles for Propulsion Marine Engines

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 32 2010-07-01 2010-07-01 false Duty Cycles for Propulsion Marine...) AIR POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Pt. 1045, App. II Appendix II to Part 1045—Duty Cycles for Propulsion Marine Engines (a)...

  5. 40 CFR Appendix II to Part 1045 - Duty Cycles for Propulsion Marine Engines

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 34 2013-07-01 2013-07-01 false Duty Cycles for Propulsion Marine...) AIR POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Pt. 1045, App. II Appendix II to Part 1045—Duty Cycles for Propulsion Marine Engines (a)...

  6. 40 CFR Appendix II to Part 1045 - Duty Cycles for Propulsion Marine Engines

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 33 2011-07-01 2011-07-01 false Duty Cycles for Propulsion Marine...) AIR POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Pt. 1045, App. II Appendix II to Part 1045—Duty Cycles for Propulsion Marine Engines (a)...

  7. 40 CFR Appendix II to Part 1045 - Duty Cycles for Propulsion Marine Engines

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 34 2012-07-01 2012-07-01 false Duty Cycles for Propulsion Marine...) AIR POLLUTION CONTROLS CONTROL OF EMISSIONS FROM SPARK-IGNITION PROPULSION MARINE ENGINES AND VESSELS Pt. 1045, App. II Appendix II to Part 1045—Duty Cycles for Propulsion Marine Engines (a)...

  8. Characterization of advanced electric propulsion systems

    NASA Technical Reports Server (NTRS)

    Ray, P. K.

    1982-01-01

    Characteristic parameters of several advanced electric propulsion systems are evaluated and compared. The propulsion systems studied are mass driver, rail gun, argon MPD thruster, hydrogen free radical thruster and mercury electron bombardment ion engine. Overall, ion engines have somewhat better characteristics as compared to the other electric propulsion systems.

  9. Characterization of advanced electric propulsion systems

    NASA Technical Reports Server (NTRS)

    Ray, P. K.

    1982-01-01

    Characteristic parameters of several advanced electric propulsion systems are evaluated and compared. The propulsion systems studied are mass driver, rail gun, argon MPD thruster, hydrogen free radical thruster and mercury electron bombardment ion engine. Overall, ion engines have somewhat better characteristics as compared to the other electric propulsion systems.

  10. Dawn Ion Propulsion System - Initial Checkout After Launch

    NASA Technical Reports Server (NTRS)

    Brophy, John R.; Garner, Charles E.; Mikes, Steven

    2008-01-01

    The first 80 days after launch of the Dawn mission were dedicated to the checkout of the spacecraft with a major emphasis on the ion propulsion system. All three ion thrusters, all three thruster-gimbal assemblies, both power processor units, both digital interface and control units, and the entire xenon feed system were completely checked out and every component was found to be in good health. Direct thrust measurements agreed well with preflight expected values for all three thrusters over the entire throttle range. Thruster electrical operating parameters and power processor units efficiencies also agreed well with preflight expected values based on acceptance test data. Two of the three ion thrusters were fully checked out within 30 days after launch. Checkout of all three thrusters was completed 64 days after launch. Deterministic thrusting with the IPS began on December 17, 2007.

  11. Simulation of an advanced techniques of ion propulsion Rocket system

    NASA Astrophysics Data System (ADS)

    Bakkiyaraj, R.

    2016-07-01

    The ion propulsion rocket system is expected to become popular with the development of Deuterium,Argon gas and Hexagonal shape Magneto hydrodynamic(MHD) techniques because of the stimulation indirectly generated the power from ionization chamber,design of thrust range is 1.2 N with 40 KW of electric power and high efficiency.The proposed work is the study of MHD power generation through ionization level of Deuterium gas and combination of two gaseous ions(Deuterium gas ions + Argon gas ions) at acceleration stage.IPR consists of three parts 1.Hexagonal shape MHD based power generator through ionization chamber 2.ion accelerator 3.Exhaust of Nozzle.Initially the required energy around 1312 KJ/mol is carrying out the purpose of deuterium gas which is changed to ionization level.The ionized Deuterium gas comes out from RF ionization chamber to nozzle through MHD generator with enhanced velocity then after voltage is generated across the two pairs of electrode in MHD.it will produce thrust value with the help of mixing of Deuterium ion and Argon ion at acceleration position.The simulation of the IPR system has been carried out by MATLAB.By comparing the simulation results with the theoretical and previous results,if reaches that the proposed method is achieved of thrust value with 40KW power for simulating the IPR system.

  12. Engine Power Turbine and Propulsion Pod Arrangement Study

    NASA Technical Reports Server (NTRS)

    Robuck, Mark; Zhang, Yiyi

    2014-01-01

    A study has been conducted for NASA Glenn Research Center under contract NNC10BA05B, Task NNC11TA80T to identify beneficial arrangements of the turboshaft engine, transmissions and related systems within the propulsion pod nacelle of NASA's Large Civil Tilt-Rotor 2nd iteration (LCTR2) vehicle. Propulsion pod layouts were used to investigate potential advantages, disadvantages, as well as constraints of various arrangements assuming front or aft shafted engines. Results from previous NASA LCTR2 propulsion system studies and tasks performed by Boeing under NASA contracts are used as the basis for this study. This configuration consists of two Fixed Geometry Variable Speed Power Turbine Engines and related drive and rotor systems (per nacelle) arranged in tilting nacelles near the wing tip. Entry-into-service (EIS) 2035 technology is assumed for both the engine and drive systems. The variable speed rotor system changes from 100 percent speed for hover to 54 percent speed for cruise by the means of a two speed gearbox concept developed under previous NASA contracts. Propulsion and drive system configurations that resulted in minimum vehicle gross weight were identified in previous work and used here. Results reported in this study illustrate that a forward shafted engine has a slight weight benefit over an aft shafted engine for the LCTR2 vehicle. Although the aft shafted engines provide a more controlled and centered CG (between hover and cruise), the length of the long rotor shaft and complicated engine exhaust arrangement outweighed the potential benefits. A Multi-Disciplinary Analysis and Optimization (MDAO) approach for transmission sizing was also explored for this study. This tool offers quick analysis of gear loads, bearing lives, efficiencies, etc., through use of commercially available RomaxDESIGNER software. The goal was to create quick methods to explore various concept models. The output results from RomaxDESIGNER have been successfully linked to Boeing

  13. Development Status of the NSTAR Ion Propulsion System Power Processor

    NASA Technical Reports Server (NTRS)

    Hamley, John A.; Pinero, Luis R.; Rawlin, Vincent K.; Miller, John R.; Cartier, Kevin C.; Bowers, Glen E.

    1995-01-01

    A 0.5-2.3 kW xenon ion propulsion system is presently being developed under the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) program. This propulsion system includes a 30 cm diameter xenon ion thruster, a Digital Control Interface Unit, a xenon feed system, and a power processing unit (PPU). The PPU consists of the power supply assemblies which operate the thruster neutralizer, main discharge chamber, and ion optics. Also included are recycle logic and a digital microcontroller. The neutralizer and discharge power supplies employ a dual use configuration which combines the functions of two power supplies into one, significantly simplifying the PPU. Further simplification was realized by implementing a single thruster control loop which regulates the beam current via the discharge current. Continuous throttling is possible over a 0.5-2.3 kW output power range. All three power supplies have been fabricated and tested with resistive loads, and have been combined into a single breadboard unit with the recycle logic and microcontroller. All line and load regulation test results show the power supplies to be within the NSTAR flight PPU specified power output of 1.98 kW. The overall efficiency of the PPU, calculated as the combined efficiencies of the power supplies and controller, at 2.3 kW delivered to resistive loads was 0.90. The component was 6.16 kg. Integration testing of the neutralizer and discharge power supplies with a functional model thruster revealed no issues with discharge ignition or steady state operation.

  14. Engine exhaust apparatus for water-jet propulsion boat

    SciTech Connect

    Nishida, H.

    1986-12-23

    A propulsion apparatus is described for a water jet propulsion boat having a hull, a water jet pump casing with the hull, and an impeller in the pump casing driven by an engine in the hull, the propulsion apparatus comprising: means in the pump casing defining a converging passage for increasing water flow velocity and having a smaller diameter section within which water jet pressure is reduced due to the increase in velocity; a hollow impeller shaft defining an exhaust passage within an inner wall of the impeller shaft. The exhaust passage is communicated at a forward end portion of the impeller shaft with an exhaust port of the engine; means connecting the impeller to the impeller shaft: an exhaust outlet communicating with the exhaust passage formed in the impeller shaft and located rearward of the impeller and within the smaller diameter section of the pump casing. This discharges exhaust into a water jet portion of higher velocity and reduced pressure within the section, that exhaust outlet communicating at a rear end surface of the impeller shaft, and a centrifugally operated stop valve for the outlet operated by the impeller shaft.

  15. Low Cost Propulsion Technology at the Marshall Space Flight Center: Fastrac Engine and the Propulsion Test Article

    NASA Technical Reports Server (NTRS)

    Fisher, Mark F.; Ise, Michael R.

    1998-01-01

    The need for low cost access to space has initiated the development of low cost liquid rocket engine and propulsion system hardware at the Marshall Space Flight Center (MSFC). The engine, the 60,000 lbf, RP-1 and LOX Fastrac Engine has been designed as a robust, low cost liquid rocket engine with applications for X-34 as well as future low cost booster systems. The engine is a turbopump fed, gas generator cycle, rocket motor with an ablative nozzle. The Propulsion Test Article (PTA) is a test bed for low cost propulsion system hardware including a composite RP-1 tank, flight feedlines and pressurization system, stacked in a booster configuration. A general description of the PTA and the Fastrac engine is given, with emphasis on the technical specification of the hardware including flow rates, pressures and other operating conditions. The process which has been used for the design and integration of this hardware is described.

  16. Low Cost Propulsion Technology at the Marshall Space Flight Center: Fastrac Engine and the Propulsion Test Article

    NASA Technical Reports Server (NTRS)

    Fisher, Mark F.; Ise, Michael R.

    1998-01-01

    The need for low cost access to space has initiated the development of low cost liquid rocket engine and propulsion system hardware at the Marshall Space Flight Center (MSFC). The engine, the 60,000 lbf, RP-1 and LOX Fastrac Engine has been designed as a robust, low cost liquid rocket engine with applications for X-34 as well as future low cost booster systems. The engine is a turbopump fed, gas generator cycle, rocket motor with an ablative nozzle. The Propulsion Test Article (PTA) is a test bed for low cost propulsion system hardware including a composite RP-1 tank, flight feedlines and pressurization system, stacked in a booster configuration. A general description of the PTA and the Fastrac engine is given, with emphasis on the technical specification of the hardware including flow rates, pressures and other operating conditions. The process which has been used for the design and integration of this hardware is described.

  17. Completely modular Thermionic Reactor Ion Propulsion System (TRIPS)

    NASA Technical Reports Server (NTRS)

    Peelgren, M. L.; Kikin, G. M.; Sawyer, C. D.

    1972-01-01

    The nuclear reactor powered ion propulsion system described is an advanced completely modularized system which lends itself to development of prototype and/or flight type components without the need for complete system tests until late in the development program. This modularity is achieved in all of the subsystems and components of the electric propulsion system including (1) the thermionic fuel elements, (2) the heat rejection subsystem (heat pipes), (3) the power conditioning modules, and (4) the ion thrusters. Both flashlight and external fuel type in-core thermionic reactors are considered as the power source. The thermionic fuel elements would be useful over a range of reactor power levels. Electrical heated acceptance testing in their flight configuration is possible for the external fuel case. Nuclear heated testing by sampling methods could be used for acceptance testing of flashlight fuel elements. The use of heat pipes for cooling the collectors and as a means of heat transport to the radiator allows early prototype or flight configuration testing of a small module of the heat rejection subsystem as opposed to full scale liquid metal pumps and radiators in a large vacuum chamber. The power conditioner (p/c) is arranged in modules with passive cooling.

  18. NASA's Evolutionary Xenon Thruster: The NEXT Ion Propulsion System for Solar System Exploration

    NASA Technical Reports Server (NTRS)

    Pencil, Eric J.; Benson, Scott W.

    2008-01-01

    This viewgraph presentation reviews NASA s Evolutionary Xenon Thruster (NEXT) Ion Propulsion system. The NEXT project is developing a solar electric ion propulsion system. The NEXT project is advancing the capability of ion propulsion to meet NASA robotic science mission needs. The NEXT system is planned to significantly improve performance over the state of the art electric propulsion systems, such as NASA Solar Electric Propulsion Technology Application Readiness (NSTAR). The status of NEXT development is reviewed, including information on the NEXT Thruster, the power processing unit, the propellant management system (PMS), the digital control interface unit, and the gimbal. Block diagrams NEXT system are presented. Also a review of the lessons learned from the Dawn and NSTAR systems is provided. In summary the NEXT project activities through 2007 have brought next-generation ion propulsion technology to a sufficient maturity level.

  19. Mission Benefits of Gridded Ion and Hall Thruster Hybrid Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Dankanich, John W.; Polsgrove, Tara

    2006-01-01

    The NASA In-Space Propulsion Technology (ISPT) Project Office has been developing the NEXT gridded ion thruster system and is planning to procure a low power Hall system. The new ion propulsion systems will join NSTAR as NASA's primary electric propulsion system options. Studies have been performed to show mission benefits of each of the stand alone systems. A hybrid ion propulsion system (IPS) can have the advantage of reduced cost, decreased flight time and greater science payload delivery over comparable homogeneous systems. This paper explores possible advantages of combining various thruster options for a single mission.

  20. Energy efficient engine: Flight propulsion system preliminary analysis and design

    NASA Technical Reports Server (NTRS)

    Johnston, R. P.; Beitler, R. S.; Bobinger, R. O.; Broman, C. L.; Gravitt, R. D.; Heineke, H.; Holloway, P. R.; Klem, J. S.; Nash, D. O.; Ortiz, P.

    1980-01-01

    The characteristics of an advanced flight propulsion system (FPS), suitable for introduction in the late 1980's to early 1990's, was more fully defined. It was determined that all goals for efficiency, environmental considerations, and economics could be met or exceeded with the possible exception of NOx emission. In evaluating the FPS, all aspects were considered including component design, performance, weight, initial cost, maintenance cost, engine system integration (including nacelle), and aircraft integration considerations. The current FPS installed specific fuel consumption was reduced 14.2% from that of the CF6-50C reference engine. When integrated into an advanced, subsonic, study transport, the FPS produced a fuel burn savings of 15 to 23% and a direct operating cost reduction of 5 to 12% depending on the mission and study aircraft characteristics relative to the reference engine.

  1. Energy efficient engine: flight propulsion system preliminary analysis and design

    SciTech Connect

    Johnston, R.P.; Beitler, R.S.; Bobinger, R.O.; Broman, C.L.; Gravitt, R.D.; Heineke, H.; Holloway, P.R.; Klem, J.S.; Nash, D.O.; Ortiz, P.

    1980-06-01

    The characteristics of an advanced flight propulsion system (FPS), suitable for introduction in the late 1980's to early 1990's, was more fully defined. It was determined that all goals for efficiency, environmental considerations, and economics could be met or exceeded with the possible exception of NOx emission. In evaluating the FPS, all aspects were considered including component design, performance, weight, initial cost, maintenance cost, engine system integration (including nacelle), and aircraft integration considerations. The current FPS installed specific fuel consumption was reduced 14.2% from that of the CF6-50C reference engine. When integrated into an advanced, subsonic, study transport, the FPS produced a fuel burn savings of 15 to 23% and a direct operating cost reduction of 5 to 12% depending on the mission and study aircraft characteristics relative to the reference engine.

  2. Energy efficient engine. Flight propulsion system preliminary analysis and design

    NASA Technical Reports Server (NTRS)

    Johnston, R. P.

    1979-01-01

    The characteristics of an advanced Flight Propulsion System (FPS) suitable for introduction in the late 1980's to early 1990's, were defined. It was determined that NASA goals for efficiency, environmental considerations, and economics could be met or exceeded with the possible exception of NOx emission. In evaluating the FPS, all aspects were considered including component design, performance, weight, initial cost, maintenance cost, engine-system integration (including nacelle), and aircraft integration considerations. In terms of the NASA goals, the current FPS installed specific fuel consumption was reduced 14.2% from that of the CF6-50C reference engine. When integrated into an advanced, subsonic, study transport, the FPS produced a fuel-burn savings of 15 to 23% and a direct operating cost reduction of 5 to 12% depending on the mission and study-aircraft characteristics relative to the reference engine.

  3. Noble gas storage and delivery system for ion propulsion

    NASA Technical Reports Server (NTRS)

    Back, Dwight Douglas (Inventor); Ramos, Charlie (Inventor)

    2001-01-01

    A method and system for storing and delivering a noble gas for an ion propulsion system where an adsorbent bearing a noble gas is heated within a storage vessel to desorb the noble gas which is then flowed through a pressure reduction device to a thruster assembly. The pressure and flow is controlled using a flow restrictor and low wattage heater which heats an adsorbent bed containing the noble gas propellant at low pressures. Flow rates of 5-60 sccm can be controlled to within about 0.5% or less and the required input power is generally less than 50 W. This noble gas storage and delivery system and method can be used for earth orbit satellites, and lunar or planetary space missions.

  4. Civil helicopter propulsion system reliability and engine monitoring technology assessments

    NASA Technical Reports Server (NTRS)

    Murphy, J. A.; Zuk, J.

    1982-01-01

    A study to reduce operating costs of helicopters, particularly directed at the maintenance of the propulsion subsystem, is presented. The tasks of the study consisted of problem definition refinement, technology solutions, diagnostic system concepts, and emergency power augmentation. Quantifiable benefits (reduced fuel consumption, on-condition engine maintenance, extended drive system overhaul periods, and longer oil change intervals) would increase the initial cost by $43,000, but the benefit of $24.46 per hour would result in breakeven at 1758 hours. Other benefits not capable of being quantified but perhaps more important include improved aircraft avilability due to reduced maintenance time, potential for increased operating limits due to continuous automatic monitoring of gages, and less time and fuel required to make engine power checks. The most important improvement is the on-condition maintenance program, which will require the development of algorithms, equipment, and procedures compatible with all operating environments.

  5. The Case for Intelligent Propulsion Control for Fast Engine Response

    NASA Technical Reports Server (NTRS)

    Litt, Jonathan S.; Frederick, Dean K.; Guo, Ten-Huei

    2009-01-01

    Damaged aircraft have occasionally had to rely solely on thrust to maneuver as a consequence of losing hydraulic power needed to operate flight control surfaces. The lack of successful landings in these cases inspired research into more effective methods of utilizing propulsion-only control. That research demonstrated that one of the major contributors to the difficulty in landing is the slow response of the engines as compared to using traditional flight control. To address this, research is being conducted into ways of making the engine more responsive under emergency conditions. This can be achieved by relaxing controller limits, adjusting schedules, and/or redesigning the regulators to increase bandwidth. Any of these methods can enable faster response at the potential expense of engine life and increased likelihood of stall. However, an example sensitivity analysis revealed a complex interaction of the limits and the difficulty in predicting the way to achieve the fastest response. The sensitivity analysis was performed on a realistic engine model, and demonstrated that significantly faster engine response can be achieved compared to standard Bill of Material control. However, the example indicates the need for an intelligent approach to controller limit adjustment in order for the potential to be fulfilled.

  6. A tissue-engineered jellyfish with biomimetic propulsion

    PubMed Central

    Nawroth, Janna C; Lee, Hyungsuk; Feinberg, Adam W; Ripplinger, Crystal M; McCain, Megan L; Grosberg, Anna; Dabiri, John O; Parker, Kevin Kit

    2014-01-01

    Reverse engineering of biological form and function requires hierarchical design over several orders of space and time. Recent advances in the mechanistic understanding of biosynthetic compound materials1–3, computer-aided design approaches in molecular synthetic biology4,5 and traditional soft robotics6,7, and increasing aptitude in generating structural and chemical microenvironments that promote cellular self-organization8–10 have enhanced the ability to recapitulate such hierarchical architecture in engineered biological systems. Here we combined these capabilities in a systematic design strategy to reverse engineer a muscular pump. We report the construction of a freely swimming jellyfish from chemically dissociated rat tissue and silicone polymer as a proof of concept. The constructs, termed ‘medusoids’, were designed with computer simulations and experiments to match key determinants of jellyfish propulsion and feeding performance by quantitatively mimicking structural design, stroke kinematics and animal-fluid interactions. The combination of the engineering design algorithm with quantitative benchmarks of physiological performance suggests that our strategy is broadly applicable to reverse engineering of muscular organs or simple life forms that pump to survive. PMID:22820316

  7. Energy efficient engine flight propulsion system: Aircraft/engine integration evaluation

    SciTech Connect

    Patt, R.F.

    1980-06-01

    Results of aircraft/engine integration studies conducted on an advanced flight propulsion system are reported. Economic evaluations of the preliminary design are included and indicate that program goals will be met. Installed sfc, DOC, noise, and emissions were evaluated. Aircraft installation considerations and growth were reviewed.

  8. Energy efficient engine flight propulsion system: Aircraft/engine integration evaluation

    NASA Technical Reports Server (NTRS)

    Patt, R. F.

    1980-01-01

    Results of aircraft/engine integration studies conducted on an advanced flight propulsion system are reported. Economic evaluations of the preliminary design are included and indicate that program goals will be met. Installed sfc, DOC, noise, and emissions were evaluated. Aircraft installation considerations and growth were reviewed.

  9. An advanced electric propulsion diagnostic (AEPD) platform for in-situ characterization of electric propulsion thrusters and ion beam sources

    NASA Astrophysics Data System (ADS)

    Bundesmann, Carsten; Eichhorn, Christoph; Scholze, Frank; Spemann, Daniel; Neumann, Horst; Pagano, Damiano; Scaranzin, Simone; Scortecci, Fabrizio; Leiter, Hans J.; Gauter, Sven; Wiese, Ruben; Kersten, Holger; Holste, Kristof; Köhler, Peter; Klar, Peter J.; Mazouffre, Stéphane; Blott, Richard; Bulit, Alexandra; Dannenmayer, Käthe

    2016-10-01

    Experimental characterization is an essential task in development, qualification and optimization process of electric propulsion thrusters or ion beam sources for material processing, because it can verify that the thruster or ion beam source fulfills the requested mission or application requirements, and it can provide parameters for thruster and plasma modeling. Moreover, there is a need for standardizing electric propulsion thruster diagnostics in order to make characterization results of different thrusters and also from measurements performed in different vacuum facilities reliable and comparable. Therefore, we have developed an advanced electric propulsion diagnostic (AEPD) platform, which allows a comprehensive in-situ characterization of electric propulsion thrusters (or ion beam sources) and could serve as a standard on-ground tool in the future. The AEPD platform uses a five-axis positioning system and provides the option to use diagnostic tools for beam characterization (Faraday probe, retarding potential analyzer, ExB probe, active thermal probe), for optical inspection (telemicroscope, triangular laser head), and for thermal characterization (pyrometer, thermocamera). Here we describe the capabilities of the diagnostic platform and provide first experimental results of the characterization of a gridded ion thruster RIT- μX.

  10. Engine System Model Development for Nuclear Thermal Propulsion

    NASA Technical Reports Server (NTRS)

    Nelson, Karl W.; Simpson, Steven P.

    2006-01-01

    In order to design, analyze, and evaluate conceptual Nuclear Thermal Propulsion (NTP) engine systems, an improved NTP design and analysis tool has been developed. The NTP tool utilizes the Rocket Engine Transient Simulation (ROCETS) system tool and many of the routines from the Enabler reactor model found in Nuclear Engine System Simulation (NESS). Improved non-nuclear component models and an external shield model were added to the tool. With the addition of a nearly complete system reliability model, the tool will provide performance, sizing, and reliability data for NERVA-Derived NTP engine systems. A new detailed reactor model is also being developed and will replace Enabler. The new model will allow more flexibility in reactor geometry and include detailed thermal hydraulics and neutronics models. A description of the reactor, component, and reliability models is provided. Another key feature of the modeling process is the use of comprehensive spreadsheets for each engine case. The spreadsheets include individual worksheets for each subsystem with data, plots, and scaled figures, making the output very useful to each engineering discipline. Sample performance and sizing results with the Enabler reactor model are provided including sensitivities. Before selecting an engine design, all figures of merit must be considered including the overall impacts on the vehicle and mission. Evaluations based on key figures of merit of these results and results with the new reactor model will be performed. The impacts of clustering and external shielding will also be addressed. Over time, the reactor model will be upgraded to design and analyze other NTP concepts with CERMET and carbide fuel cores.

  11. Key Reliability Drivers of Liquid Propulsion Engines and A Reliability Model for Sensitivity Analysis

    NASA Technical Reports Server (NTRS)

    Huang, Zhao-Feng; Fint, Jeffry A.; Kuck, Frederick M.

    2005-01-01

    This paper is to address the in-flight reliability of a liquid propulsion engine system for a launch vehicle. We first establish a comprehensive list of system and sub-system reliability drivers for any liquid propulsion engine system. We then build a reliability model to parametrically analyze the impact of some reliability parameters. We present sensitivity analysis results for a selected subset of the key reliability drivers using the model. Reliability drivers identified include: number of engines for the liquid propulsion stage, single engine total reliability, engine operation duration, engine thrust size, reusability, engine de-rating or up-rating, engine-out design (including engine-out switching reliability, catastrophic fraction, preventable failure fraction, unnecessary shutdown fraction), propellant specific hazards, engine start and cutoff transient hazards, engine combustion cycles, vehicle and engine interface and interaction hazards, engine health management system, engine modification, engine ground start hold down with launch commit criteria, engine altitude start (1 in. start), Multiple altitude restart (less than 1 restart), component, subsystem and system design, manufacturing/ground operation support/pre and post flight check outs and inspection, extensiveness of the development program. We present some sensitivity analysis results for the following subset of the drivers: number of engines for the propulsion stage, single engine total reliability, engine operation duration, engine de-rating or up-rating requirements, engine-out design, catastrophic fraction, preventable failure fraction, unnecessary shutdown fraction, and engine health management system implementation (basic redlines and more advanced health management systems).

  12. German Jumo 004 Engine at the Lewis Flight Propulsion Laboratory

    NASA Image and Video Library

    1946-03-21

    Researcher Robert Miller led an investigation into the combustor performance of a German Jumo 004 engine at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Jumo 004 powered the world's first operational jet fighter, the Messerschmitt Me 262, beginning in 1942. The Me 262 was the only jet aircraft used in combat during World War II. The eight-stage axial-flow compressor Jumo 004 produced 2000 pounds of thrust. The US Army Air Forces provided the NACA with a Jumo 004 engine in 1945 to study the compressor’s design and performance. Conveniently the engine’s designer Anselm Franz had recently arrived at Wright-Patterson Air Force Base in nearby Dayton, Ohio as part of Project Paperclip. The Lewis researchers used a test rig in the Engine Research Building to analyze one of the six combustion chambers. It was difficult to isolate a single combustor’s performance when testing an entire engine. The combustion efficiency, outlet-temperature distribution, and total pressure drop were measured. The researchers determined the Jumo 004’s maximum performance was 5000 revolutions per minute at a 27,000 foot altitude and 11,000 revolutions per minute at a 45,000 foot altitude. The setup in this photograph was created for a tour of NACA Lewis by members of the Institute of Aeronautical Science on March 22, 1945.

  13. Test bed ion engine development

    NASA Technical Reports Server (NTRS)

    Aston, G.; Deininger, W. D.

    1984-01-01

    A test bed ion (TBI) engine was developed to serve as a tool in exploring the limits of electrostatic ion thruster performance. A description of three key ion engine components, the decoupled extraction and amplified current (DE-AC) accelerator system, field enhanced refractory metal (FERM) hollow cathode and divergent line cusp (DLC) discharge chamber, whose designs and operating philosophies differ markedly from conventional thruster technology is given. Significant program achievements were: (1) high current density DE-AC accelerator system operation at low electric field stress with indicated feasibility of a 60 mA/sq cm argon ion beam; (2) reliable FERM cathode start up times of 1 to 2 secs. and demonstrated 35 ampere emission levels; (3) DLC discharge chamber plasma potentials negative of anode potential; and (4) identification of an efficient high plasma density engine operating mode. Using the performance projections of this program and reasonable estimates of other parameter values, a 1.0 Newton thrust ion engine is identified as a realizable technology goal. Calculations show that such an engine, comparable in beam area to a J series 30 cm thruster, could, operating on Xe or Hg, have thruster efficiencies as high as 0.76 and 0.78 respectively, with a 100 eV/ion discharge loss.

  14. Enabling propulsion materials for high-speed civil transport engines

    NASA Technical Reports Server (NTRS)

    Stephens, Joseph R.; Herbell, Thomas P.

    1992-01-01

    NASA Headquarters and LeRC have advocated an Enabling Propulsion Materials Program (EPM) to begin in FY-92. The High Speed Research Phase 1 program which began in FY-90 has focused on the environmental acceptability of a High Speed Civil Transport (HSCT). Studies by industry, including Boeing, McDonnell Douglas, GE Aircraft Engines, and Pratt & Whitney Aircraft, and in-house studies by NASA concluded that NO(x) emissions and airport noise reduction can only be economically achieved by revolutionary advancements in materials technologies. This is especially true of materials for the propulsion system where the combustor is the key to maintaining low emissions, and the exhaust nozzle is the key to reducing airport noise to an acceptable level. Both of these components will rely on high temperature composite materials that can withstand the conditions imposed by commercial aircraft operations. The proposed EPM program will operate in conjunction with the HSR Phase 1 Program and the planned HSR Phase 2 program slated to start in FY-93. Components and subcomponents developed from advanced materials will be evaluated in the HSR Phase 2 Program.

  15. Deep Space 1 Ion Engine Completed a 3-Year Journey

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Patterson, Michael J.; Rawlin, Vincent K.; Hamley, John A.

    2001-01-01

    A xenon ion engine and power processor system, which was developed by the NASA Glenn Research Center in partnership with the Jet Propulsion Laboratory and Boeing Electron Dynamic Devices, completed nearly 3 years of operation aboard the Deep Space 1 spacecraft. The 2.3-kW ion engine, which provided primary propulsion and two-axis attitude control, thrusted for more than 16,000 hr and consumed more than 70 kg of xenon propellant. The Deep Space 1 spacecraft was launched on October 24, 1998, to validate 12 futuristic technologies, including the ion-propulsion system. After the technology validation process was successfully completed, the Deep Space 1 spacecraft flew by the small asteroid Braille on July 29, 1999. The final objective of this mission was to encounter the active comet Borrelly, which is about 6 miles long. The ion engine was on a thrusting schedule to navigate the Deep Space 1 spacecraft to within 1400 miles of the comet. Since the hydrazine used for spacecraft attitude control was in short supply, the ion engine also provided two-axis attitude control to conserve the hydrazine supply for the Borrelly encounter. The comet encounter took place on September 22, 2001. Dr. Marc Rayman, project manager of Deep Space 1 at the Jet Propulsion Laboratory said, "Deep Space 1 plunged into the heart of the comet Borrelly and has lived to tell every detail of its spinetingling adventure! The images are even better than the impressive images of comet Halley taken by Europe's Giotto spacecraft in 1986." The Deep Space 1 mission, which successfully tested the 12 high-risk, advanced technologies and captured the best images ever taken of a comet, was voluntarily terminated on December 18, 2001. The successful demonstration of the 2-kW-class ion propulsion system technology is now providing mission planners with off-the-shelf flight hardware. Higher power, next generation ion propulsion systems are being developed for large flagship missions, such as outer planet

  16. Antiproton powered propulsion with magnetically confined plasma engines

    SciTech Connect

    Lapointe, M.R.

    1989-08-01

    Matter-antimatter annihilation releases more energy per unit mass than any other method of energy production, making it an attractive energy source for spacecraft propulsion. In the magnetically confined plasma engine, antiproton beams are injected axially into a pulsed magnetic mirror system, where they annihilate with an initially neutral hydrogen gas. The resulting charged annihilation products transfer energy to the hydrogen propellant, which is then exhausted through one end of the pulsed mirror system to provide thrust. The calculated energy transfer efficiencies for a low number density (10(14)/cu cm) hydrogen propellant are insufficient to warrant operating the engine in this mode. Efficiencies are improved using moderate propellant number densities (10(16)/cu cm), but the energy transferred to the plasma in a realistic magnetic mirror system is generally limited to less than 2 percent of the initial proton-antiproton annihilation energy. The energy transfer efficiencies are highest for high number density (10(18)/cu cm) propellants, but plasma temperatures are reduced by excessive radiation losses. Low to moderate thrust over a wide range of specific impulse can be generated with moderate propellant number densities, while higher thrust but lower specific impulse may be generated using high propellant number densities. Significant mass will be required to shield the superconducting magnet coils from the high energy gamma radiation emitted by neutral pion decay. The mass of such a radiation shield may dominate the total engine mass, and could severely diminish the performance of antiproton powered engines which utilize magnetic confinement. The problem is compounded in the antiproton powered plasma engine, where lower energy plasma bremsstrahlung radiation may cause shield surface ablation and degradation.

  17. Antiproton powered propulsion with magnetically confined plasma engines

    NASA Technical Reports Server (NTRS)

    Lapointe, Michael R.

    1989-01-01

    Matter-antimatter annihilation releases more energy per unit mass than any other method of energy production, making it an attractive energy source for spacecraft propulsion. In the magnetically confined plasma engine, antiproton beams are injected axially into a pulsed magnetic mirror system, where they annihilate with an initially neutral hydrogen gas. The resulting charged annihilation products transfer energy to the hydrogen propellant, which is then exhausted through one end of the pulsed mirror system to provide thrust. The calculated energy transfer efficiencies for a low number density (10(14)/cu cm) hydrogen propellant are insufficient to warrant operating the engine in this mode. Efficiencies are improved using moderate propellant number densities (10(16)/cu cm), but the energy transferred to the plasma in a realistic magnetic mirror system is generally limited to less than 2 percent of the initial proton-antiproton annihilation energy. The energy transfer efficiencies are highest for high number density (10(18)/cu cm) propellants, but plasma temperatures are reduced by excessive radiation losses. Low to moderate thrust over a wide range of specific impulse can be generated with moderate propellant number densities, while higher thrust but lower specific impulse may be generated using high propellant number densities. Significant mass will be required to shield the superconducting magnet coils from the high energy gamma radiation emitted by neutral pion decay. The mass of such a radiation shield may dominate the total engine mass, and could severely diminish the performance of antiproton powered engines which utilize magnetic confinement. The problem is compounded in the antiproton powered plasma engine, where lower energy plasma bremsstrahlung radiation may cause shield surface ablation and degradation.

  18. The Mission Defines the Cycle: Turbojet, Turbofan and Variable Cycle Engines for High Speed Propulsion

    DTIC Science & Technology

    2010-09-01

    RTO-EN-AVT-185 2 - 1 The Mission Defines the Cycle: Turbojet , Turbofan and Variable Cycle Engines for High Speed Propulsion Joachim Kurzke...different variable cycle engine configurations have been studied in the past. Ref. 4 gives an overview about the work done at General Electric Aircraft...Defines the Cycle: Turbojet , Turbofan and Variable Cycle Engines for High Speed Propulsion RTO-EN-AVT-185 2 - 23 • If VABI 1 is closed and VABI 2 open

  19. Ion Propulsion: An Enabling Technology for the Dawn Mission

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Mark M.; Whiffen, Greg J.; Brophy, John R.; Mikes, Steven C.

    2013-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, 4 Vesta, and the dwarf planet 1 Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 rocket that placed the 1218-kg spacecraft into an Earth-escape (heliocentic) trajectory with an escape velocity of 11 km/s. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide an additional delta-V of approximately 11 km/s for the heliocentric transfers to each body and for all orbit transfers including orbit capture/escape and transition to the various science orbits. Deterministic thrusting to Vesta began in December 2007 and concluded with orbit capture at Vesta in July 2011. The transfer to Vesta included a Mars gravity assist flyby in February 2009 that provided an additional delta-V of 2.6 km/s and was the only post-launch mission delta-V not provided by IPS. The IPS was used during the 14 months at Vesta for all science orbit transfers and then for Vesta escape. Deterministic thrusting for Ceres began in late August 2012 with a planned arrival date at Ceres in early 2015, whereupon IPS will be used for all science orbit transfers. As of January 2013 the IPS has been operated for approximately 28,000 hours, consumed approximately 280 kg of xenon, and provided a delta-V of approximately 7.5 km/s, the most post-launch delta-V of any spacecraft yet flown. IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. Use of the IPS together with a medium-priced launch vehicle enabled this high delta-V mission to be performed within the Discovery Program cost cap. This paper provides an overview of the Dawn IPS and its mission operations through departure for Ceres.

  20. A cyclic ground test of an ion auxiliary propulsion system: Description and operational considerations

    NASA Technical Reports Server (NTRS)

    Ling, Jerri S.; Kramer, Edward H.

    1988-01-01

    The Ion Auxiliary Propulsion System (IAPS) experiment is designed for launch on an Air Force Space Test Program satellite (NASA-TM-78859; AIAA Paper No. 78-647). The primary objective of the experiment is to flight qualify the 8 cm mercury ion thruster system for stationkeeping applications. Secondary objectives are measuring the interactions between operating ion thruster systems and host spacecraft, and confirming the design performance of the thruster systems. Two complete 8 cm mercury ion thruster subsystems will be flown. One of these will be operated for 2557 on and off cycles and 7057 hours at full thrust. Tests are currently under way in support of the IAPS flight experiment. In this test an IAPS thruster is being operated through a series of startup/run/shut-down cycles which simulate thruster operation during the planned flight experiment. A test facility description and operational considerations of this testing using an engineering model 8 cm thruster (S/N 905) is the subject of this paper. Final results will be published at a later date when the ground test has been concluded.

  1. Energy Efficient Engine flight propulsion system preliminary analysis and design report

    SciTech Connect

    Bisset, J.W.; Howe, D.C.

    1983-09-01

    The final design and analysis of the flight propulsion system is presented. This system is the conceptual study engine defined to meet the performance, economic and environmental goals established for the Energy Efficient Engine Program. The design effort included a final definition of the engine, major components, internal subsystems, and nacelle. Various analytical representations and results from component technology programs are used to verify aerodynamic and structural design concepts and to predict performance. Specific design goals and specifications, reflecting future commercial aircraft propulsion system requirements for the mid-1980's, are detailed by NASA and used as guidelines during engine definition. Information is also included which details salient results from a separate study to define a turbofan propulsion system, known as the maximum efficiency engine, which reoptimized the advanced fuel saving technologies for improved fuel economy and direct operating costs relative to the flight propulsion system.

  2. Energy Efficient Engine Flight Propulsion System Preliminary Analysis and Design Report

    NASA Technical Reports Server (NTRS)

    Bisset, J. W.; Howe, D. C.

    1983-01-01

    The final design and analysis of the flight propulsion system is presented. This system is the conceptual study engine defined to meet the performance, economic and environmental goals established for the Energy Efficient Engine Program. The design effort included a final definition of the engine, major components, internal subsystems, and nacelle. Various analytical representations and results from component technology programs are used to verify aerodynamic and structural design concepts and to predict performance. Specific design goals and specifications, reflecting future commercial aircraft propulsion system requirements for the mid-1980's, are detailed by NASA and used as guidelines during engine definition. Information is also included which details salient results from a separate study to define a turbofan propulsion system, known as the maximum efficiency engine, which reoptimized the advanced fuel saving technologies for improved fuel economy and direct operating costs relative to the flight propulsion system.

  3. Recovering Residual Xenon Propellant for an Ion Propulsion System

    NASA Technical Reports Server (NTRS)

    Ganapathi, Gani; Skakkottai, P.; wu, Jiunn Jeng

    2006-01-01

    Future nuclear-powered Ion-Propulsion- System-propelled spacecraft such as Jupiter Icy Moon Orbiter (JIMO) will carry more than 10,000 kg of xenon propellant. Typically, a small percentage of this propellant cannot be used towards the end of the mission because of the pressure drop requirements for maintaining flow. For large missions such as JIMO, this could easily translate to over 250 kg of unusable xenon. A proposed system, the Xenon Recovery System (XRS), for recovering almost all of the xenon remaining in the tank, would include a cryopump in the form of a condenser/evaporator that would be alternatively cooled by a radiator, then heated electrically. When the pressure of the xenon in the tank falls below 0.7 MPa (100 psia), the previously isolated XRS will be brought online and the gas from the tank would enter the cryopump that is initially cooled to a temperature below saturation temperature of xenon. This causes xenon liquefaction and further cryopumping from the tank till the cryopump is full of liquid xenon. At this point, the cryopump is heated electrically by small heaters (70 to 80 W) to evaporate the liquid that is collected as high-pressure gas (<7 MPa; 1,000 psia) in an intermediate accumulator. Check valves between the tank and the XRS prevent the reverse flow of xenon during the heating cycle. The accumulator serves as the high-pressure source of xenon gas to the Xenon Feed System (XFS) downstream of the XRS. This cycle is repeated till almost all the xenon is recovered. Currently, this system is being baselined for JIMO.

  4. Mission Capabilities of Ion Engines Using SNAP-8 Power Supplies

    NASA Technical Reports Server (NTRS)

    Edelbaum, T. N.; Fimple, W. R.; Gobetz, F. W.; London, H. S.

    1961-01-01

    Mission performance capabilities of ion engines powered by the 30 kw and 60 kw SNAP-8 power supplies are compared for the following missions: a 24-hr equatorial satellite, a 100 n mi lunar satellite, a 500 n mi Mars satellite, a Mercury probe, and an out-of-the-ecliptic probe. The capabilities of arc- jet engines and chemical engines for the same missions are compared with those of the ion engines. The majority of the comparisons are for 8500-lb spacecraft which are boosted into a 300 n mi orbit by the Atlas-Centaur. Variations in initial orbit altitude, the use of actual launch dates rather than dates based on simplifying assumptions, and the combined use of chemical and electrical propulsion systems were also evaluated in terms of their effect on mission performance.

  5. Environmental Testing of the NEXT PM1R Ion Engine

    NASA Technical Reports Server (NTRS)

    Snyder, John S.; Anderson, John R.; VanNoord, Jonathan L.; Soulas, George C.

    2007-01-01

    The NEXT propulsion system is an advanced ion propulsion system presently under development that is oriented towards robotic exploration of the solar system using solar electric power. The subsystem includes an ion engine, power processing unit, feed system components, and thruster gimbal. The Prototype Model engine PM1 was subjected to qualification-level environmental testing in 2006 to demonstrate compatibility with environments representative of anticipated mission requirements. Although the testing was largely successful, several issues were identified including the fragmentation of potting cement on the discharge and neutralizer cathode heater terminations during vibration which led to abbreviated thermal testing, and generation of particulate contamination from manufacturing processes and engine materials. The engine was reworked to address most of these findings, renamed PM1R, and the environmental test sequence was repeated. Thruster functional testing was performed before and after the vibration and thermal-vacuum tests. Random vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10.0 Grms for 2 min in each of three axes. Thermal-vacuum testing included three thermal cycles from 120 to 215 C with hot engine re-starts. Thruster performance was nominal throughout the test program, with minor variations in a few engine operating parameters likely caused by facility effects. There were no significant changes in engine performance as characterized by engine operating parameters, ion optics performance measurements, and beam current density measurements, indicating no significant changes to the hardware as a result of the environmental testing. The NEXT PM1R engine and the breadboard gimbal were found to be well-designed against environmental requirements based on the results reported herein. The redesigned cathode heater terminations successfully survived the vibration environments. Based on the results of this test

  6. A 50 cm diameter annular ion engine

    NASA Technical Reports Server (NTRS)

    Aston, Graeme; Brophy, John R.

    1989-01-01

    An ion engine design is presented which uses an annular geometry as a means of achieving large engine diameters and hence, high thrust levels. Preliminary results are discussed for discharge-only operation of a 50-cm-diameter annular ion engine. Measured operating parameters presented include discharge current and voltage characteristics, discharge chamber ion current distribution, engine body temperatures, plasma flatness parameter effects and total integrated grid ion current.

  7. Electrical Prototype Power Processor for the 30-cm Mercury electric propulsion engine

    NASA Technical Reports Server (NTRS)

    Biess, J. J.; Frye, R. J.

    1978-01-01

    An Electrical Prototpye Power Processor has been designed to the latest electrical and performance requirements for a flight-type 30-cm ion engine and includes all the necessary power, command, telemetry and control interfaces for a typical electric propulsion subsystem. The power processor was configured into seven separate mechanical modules that would allow subassembly fabrication, test and integration into a complete power processor unit assembly. The conceptual mechanical packaging of the electrical prototype power processor unit demonstrated the relative location of power, high voltage and control electronic components to minimize electrical interactions and to provide adequate thermal control in a vacuum environment. Thermal control was accomplished with a heat pipe simulator attached to the base of the modules.

  8. Electrical Prototype Power Processor for the 30-cm Mercury electric propulsion engine

    NASA Technical Reports Server (NTRS)

    Biess, J. J.; Frye, R. J.

    1978-01-01

    An Electrical Prototpye Power Processor has been designed to the latest electrical and performance requirements for a flight-type 30-cm ion engine and includes all the necessary power, command, telemetry and control interfaces for a typical electric propulsion subsystem. The power processor was configured into seven separate mechanical modules that would allow subassembly fabrication, test and integration into a complete power processor unit assembly. The conceptual mechanical packaging of the electrical prototype power processor unit demonstrated the relative location of power, high voltage and control electronic components to minimize electrical interactions and to provide adequate thermal control in a vacuum environment. Thermal control was accomplished with a heat pipe simulator attached to the base of the modules.

  9. Thermal Environmental Testing of NSTAR Engineering Model Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Rawlin, Vincent K.; Patterson, Michael J.; Becker, Raymond A.

    1999-01-01

    NASA's New Millenium program will fly a xenon ion propulsion system on the Deep Space 1 Mission. Tests were conducted under NASA's Solar Electric Propulsion Technology Applications Readiness (NSTAR) Program with 3 different engineering model ion thrusters to determine thruster thermal characteristics over the NSTAR operating range in a variety of thermal environments. A liquid nitrogen-cooled shroud was used to cold-soak the thruster to -120 C. Initial tests were performed prior to a mature spacecraft design. Those results and the final, severe, requirements mandated by the spacecraft led to several changes to the basic thermal design. These changes were incorporated into a final design and tested over a wide range of environmental conditions.

  10. Structured system engineering methodologies used to develop a nuclear thermal propulsion engine

    NASA Technical Reports Server (NTRS)

    Corban, R.; Wagner, R.

    1993-01-01

    To facilitate the development of a space nuclear thermal propulsion engine for manned flights to Mars, requirements must be established early in the technology development cycle. The long lead times for the acquisition of the engine system and nuclear test facilities demands that the engine system size, performance and safety goals be defined at the earliest possible time. These systems are highly complex and require a large multidisciplinary systems engineering team to develop and track requirements, and to ensure that the as-built system reflects the intent of the mission. A methodology has been devised which uses sophisticated computer tools to effectively develop and interpret functional requirements, and furnish these to the specification level for implementation.

  11. Development of an electrostatic propulsion engine using sub-micron powders as the reaction mass

    NASA Technical Reports Server (NTRS)

    Herbert, F.; Kendall, K. R.

    1991-01-01

    Asteroid sample return missions would benefit from development of an improved rocket engine. Chemical rockets achieve their large thrust with high mass consumption rate (dm/dt) but low exhaust velocity; therefore, a large fraction of their total mass is fuel. Present day ion thrusters are characterized by high exhaust velocity, but low dm/dt; thus, they are inherently low thrust devices. However, their high exhausy velocity is poorly matched to typical mission requirements and therefore, wastes energy. A better match would be intermediate between the two forms of propulsion. This could be achieved by electrostatically accelerating solid powder grains, raising the possibility that interplanetary material could be processed to use as reaction mass. An experiment to study the charging properties of sub-micron sized powder grains is described. If a suitable material can be identified, then it could be used as the reaction mass in an electrostatic propulsion engine. The experiment employs a time of flight measurement to determine the exhaust velocity (v) of various negatively charged powder grains that were charged and accelerated in a simple device. The purpose is to determine the charge to mass ratio that can be sustained for various substances. In order to be competitive with present day ion thrusters, a specific impulse (v/g) of 3000 to 5000 seconds is required. Preliminary results are presented. More speculatively, there are some mission profiles that would benefit from collection of reaction mass at the remote asteroid site. Experiments that examine the generation of sub-micron clusters by electrostatic self-disruption of geologically derived material are planned.

  12. In-Flight Operation of the Dawn Ion Propulsion System - The First Nine Months

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Brophy, John R.; Mikes, Steven C.; Raymond, Marc D.

    2008-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218 kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) which will provide most of the delta-V needed for heliocentric transfer to Vesta, orbit capture at Vesta, transfer to Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer to Ceres science orbits. The Dawn ion engine design is based on the design validated on NASA's Deep Space 1 mission. However, because of the very substantial (11 km/s) delta-V requirements for this mission Dawn requires two engines to complete its mission objectives. The power processor units (PPU), digital control and interface units (DCIU) slice boards and the xenon control assembly (XCA) are also based on the DS1 design. The DCIUs and thrust gimbal assemblies (TGA) were developed at the Jet Propulsion Laboratory. The spacecraft was provided by Orbital Sciences Corporation, Sterling, Virginia, and the mission is managed by and operated from the Jet Propulsion Laboratory. Dawn partnered with Germany, Italy and Los Alamos National Laboratory for the science instruments. The mission is led by the principal investigator, Dr. Christopher Russell, from the University of California, Los Angeles. The first 80 days after launch were dedicated to the initial checkout of the spacecraft prior to the initiation of long-term thrusting for the heliocentric transfer to Vesta. The IPS hardware, consisting of three ion thrusters and TGAs, two PPUs and DCIUs, xenon feed system, and spacecraft control software, was investigated extensively. Thrust measurements, roll torque measurements, pointing capabilities, control

  13. In-Flight Operation of the Dawn Ion Propulsion System - The First Nine Months

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Brophy, John R.; Mikes, Steven C.; Raymond, Marc D.

    2008-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218 kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) which will provide most of the delta-V needed for heliocentric transfer to Vesta, orbit capture at Vesta, transfer to Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer to Ceres science orbits. The Dawn ion engine design is based on the design validated on NASA's Deep Space 1 mission. However, because of the very substantial (11 km/s) delta-V requirements for this mission Dawn requires two engines to complete its mission objectives. The power processor units (PPU), digital control and interface units (DCIU) slice boards and the xenon control assembly (XCA) are also based on the DS1 design. The DCIUs and thrust gimbal assemblies (TGA) were developed at the Jet Propulsion Laboratory. The spacecraft was provided by Orbital Sciences Corporation, Sterling, Virginia, and the mission is managed by and operated from the Jet Propulsion Laboratory. Dawn partnered with Germany, Italy and Los Alamos National Laboratory for the science instruments. The mission is led by the principal investigator, Dr. Christopher Russell, from the University of California, Los Angeles. The first 80 days after launch were dedicated to the initial checkout of the spacecraft prior to the initiation of long-term thrusting for the heliocentric transfer to Vesta. The IPS hardware, consisting of three ion thrusters and TGAs, two PPUs and DCIUs, xenon feed system, and spacecraft control software, was investigated extensively. Thrust measurements, roll torque measurements, pointing capabilities, control

  14. Primary propulsion of electrothermal, ion, and chemical systems for space-based radar orbit transfer

    NASA Technical Reports Server (NTRS)

    Wang, S.-Y.; Staiger, P. J.

    1985-01-01

    An orbit transfer mission concept has been studied for a Space-Based Radar (SBR) where 40 kW required for radar operation is assumed available for orbit transfer propulsion. Arcjet, pulsed electrothermal (PET), ion, and storable chemical systems are considered for the primary propulsion. Transferring two SBR per shuttle flight to 1112 km/60 deg using eiectrical propulsion systems offers an increased payload at the expense of increased trip time, up to 2000 kg each, which may be critical for survivability. Trade offs between payload mass, transfer time, launch site, inclination, and height of parking orbits are presented.

  15. Primary propulsion of electrothermal, ion and chemical systems for space-based radar orbit transfer

    NASA Technical Reports Server (NTRS)

    Wang, S. Y.; Staiger, P. J.

    1985-01-01

    An orbit transfer mission concept has been studied for a Space-Based Radar (SBR) where 40 kW required for radar operation is assumed available for orbit transfer propulsion. Arcjet, pulsed electrothermal (PET), ion, and storable chemical systems are considered for the primary propulsion. Transferring two SBR per shuttle flight to 1112 km/60 deg using electrical propulsion systems offers an increased payload at the expense of increased trip time, up to 2000 kg each, which may be critical for survivability. Trade offs between payload mass, transfer time, launch site, inclination, and height of parking orbits are presented.

  16. Primary propulsion of electrothermal, ion, and chemical systems for space-based radar orbit transfer

    NASA Technical Reports Server (NTRS)

    Wang, S.-Y.; Staiger, P. J.

    1985-01-01

    An orbit transfer mission concept has been studied for a Space-Based Radar (SBR) where 40 kW required for radar operation is assumed available for orbit transfer propulsion. Arcjet, pulsed electrothermal (PET), ion, and storable chemical systems are considered for the primary propulsion. Transferring two SBR per shuttle flight to 1112 km/60 deg using eiectrical propulsion systems offers an increased payload at the expense of increased trip time, up to 2000 kg each, which may be critical for survivability. Trade offs between payload mass, transfer time, launch site, inclination, and height of parking orbits are presented.

  17. Current Density Measurements of an Annular-Geometry Ion Engine

    NASA Technical Reports Server (NTRS)

    Shastry, Rohit; Patterson, Michael J.; Herman, Daniel A.; Foster, John E.

    2012-01-01

    The concept of the annular-geometry ion engine, or AGI-Engine, has been shown to have many potential benefits when scaling electric propulsion technologies to higher power. However, the necessary asymmetric location of the discharge cathode away from thruster centerline could potentially lead to non-uniformities in the discharge not present in conventional geometry ion thrusters. In an effort to characterize the degree of this potential nonuniformity, a number of current density measurements were taken on a breadboard AGI-Engine. Fourteen button probes were used to measure the ion current density of the discharge along a perforated electrode that replaced the ion optics during conditions of simulated beam extraction. Three Faraday probes spaced apart in the vertical direction were also used in a separate test to interrogate the plume of the AGI-Engine during true beam extraction. It was determined that both the discharge and the plume of the AGI-Engine are highly uniform, with variations under most conditions limited to 10% of the average current density in the discharge and 5% of the average current density in the plume. Beam flatness parameter measured 30 mm from the ion optics ranged from 0.85 0.95, and overall uniformity was shown to generally increase with increasing discharge and beam currents. These measurements indicate that the plasma is highly uniform despite the asymmetric location of the discharge cathode.

  18. Current Density Measurements of an Annular-Geometry Ion Engine

    NASA Technical Reports Server (NTRS)

    Shastry, Rohit; Patterson, Michael J.; Herman, Daniel A.; Foster, John E.

    2012-01-01

    The concept of the annular-geometry ion engine, or AGI-Engine, has been shown to have many potential benefits when scaling electric propulsion technologies to higher power. However, the necessary asymmetric location of the discharge cathode away from thruster centerline could potentially lead to non-uniformities in the discharge not present in conventional geometry ion thrusters. In an effort to characterize the degree of this potential non-uniformity, a number of current density measurements were taken on a breadboard AGI-Engine. Fourteen button probes were used to measure the ion current density of the discharge along a perforated electrode that replaced the ion optics during conditions of simulated beam extraction. Three Faraday probes spaced apart in the vertical direction were also used in a separate test to interrogate the plume of the AGI-Engine during true beam extraction. It was determined that both the discharge and the plume of the AGI-Engine are highly uniform, with variations under most conditions limited to +/-10% of the average current density in the discharge and +/-5% of the average current density in the plume. Beam flatness parameter measured 30 mm from the ion optics ranged from 0.85 - 0.95, and overall uniformity was shown to generally increase with increasing discharge and beam currents. These measurements indicate that the plasma is highly uniform despite the asymmetric location of the discharge cathode.

  19. ECR Discharge Ion Engines and Their Space Experiences

    NASA Astrophysics Data System (ADS)

    Kuninaka, Hitoshi; Nakai, Tatsuya; Nishiyama, Kazutaka

    2006-10-01

    Ion engine μ10 has a long life and high reliability because of electrodeless ECR plasma generation in both the ion generator and the neutralizer using 4GHz microwave. Measurements on the electron energy distribution in the ion generator revealed the discharge mechanism to heat gradually a part of the thermal electron along magnetic track. The high-energy electrons generate ions in collision process and return to the thermal electrons. The recycling process of electrons results in the effective plasma generation in comparison with the DC discharge ion generator, in which the high-energy electrons are expendable. Four μ10, each generating a thrust of 8 mN, specific impulse of 3,200 seconds, and consuming 350 W of electric power, propel the ``HAYABUSA'' asteroid explorer launched on May 2003. After vacuum exposure and several runs of baking to reduce residual gas, the ion engine system established continuous acceleration. In 2005, HAYABUSA, using solar electric propulsion, managed to successfully cover the distance between 0.86 AU and 1.7 AU in the solar system, as well as rendezvous with, land on, and lift off from the asteroid. During the 3-year flight, the ion engine system generated a delta-V of 1,400 m/s while consuming 22 kg of xenon propellant and operating for 25,900 hours.

  20. Impact of ion propulsion on performance, design, testing and operation of a geosynchronous spacecraft. Master's thesis

    SciTech Connect

    Lugtu, S.D.

    1990-06-01

    This thesis presents the implementation issues of an ion propulsion subsystem (IPS) on a geosynchronous communications satellite. As an example, Ultra High Frequency (UHF) Follow-On class satellite is selected for this study. The issues include: (1) impact of integration of IPS with other subsystems, such as the electrical power subsystem to take care of the heavy demand of power requirements and location of the subsystem with least impact on attitude control and plume impingement on solar arrays, (2) environmental considerations-particulate contamination, electrostatic discharge (ESD), and electromagnetic interference (EMI), and finally risks and benefits. Ion propulsion offers significant advantages over chemical propulsion due to its high specific impulse and the advent of xenon thruster technology, multikilowatt spacecraft, and nickel-hydrogen (Ni-H2) batteries with demonstrated high cycle life have combined to make the ion thruster attractive for North-South Station Keeping (NSSK).

  1. Deep Space 1 Using its Ion Engine (Artist's Concept)

    NASA Technical Reports Server (NTRS)

    2003-01-01

    NASA's New Millennium Deep Space 1 spacecraft approaching the comet 19P/Borrelly. With its primary mission to serve as a technology demonstrator--testing ion propulsion and 11 other advanced technologies--successfully completed in September 1999, Deep Space 1 is now headed for a risky, exciting rendezvous with Comet Borrelly. NASA extended the mission, taking advantage of the ion propulsion and other systems to target the daring encounter with the comet in September 2001. Once a sci-fi dream, the ion propulsion engine has powered the spacecraft for over 12,000 hours. Another onboard experiment includes software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The first flight in NASA's New Millennium Program, Deep Space 1 was launched October 24, 1998 aboard a Boeing Delta 7326 rocket from Cape Canaveral Air Station, FL. Deep Space 1 successfully completed and exceeded its mission objectives in July 1999 and flew by a near-Earth asteroid, Braille (1992 KD), in September 1999.

  2. Deep Space 1 Using its Ion Engine (Artist's Concept)

    NASA Technical Reports Server (NTRS)

    2003-01-01

    NASA's New Millennium Deep Space 1 spacecraft approaching the comet 19P/Borrelly. With its primary mission to serve as a technology demonstrator--testing ion propulsion and 11 other advanced technologies--successfully completed in September 1999, Deep Space 1 is now headed for a risky, exciting rendezvous with Comet Borrelly. NASA extended the mission, taking advantage of the ion propulsion and other systems to target the daring encounter with the comet in September 2001. Once a sci-fi dream, the ion propulsion engine has powered the spacecraft for over 12,000 hours. Another onboard experiment includes software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The first flight in NASA's New Millennium Program, Deep Space 1 was launched October 24, 1998 aboard a Boeing Delta 7326 rocket from Cape Canaveral Air Station, FL. Deep Space 1 successfully completed and exceeded its mission objectives in July 1999 and flew by a near-Earth asteroid, Braille (1992 KD), in September 1999.

  3. Deep Space 1 Using its Ion Engine Artist Concept

    NASA Image and Video Library

    2003-07-02

    NASA's New Millennium Deep Space 1 spacecraft approaching the comet 19P/Borrelly. With its primary mission to serve as a technology demonstrator--testing ion propulsion and 11 other advanced technologies--successfully completed in September 1999, Deep Space 1 is now headed for a risky, exciting rendezvous with Comet Borrelly. NASA extended the mission, taking advantage of the ion propulsion and other systems to target the daring encounter with the comet in September 2001. Once a sci-fi dream, the ion propulsion engine has powered the spacecraft for over 12,000 hours. Another onboard experiment includes software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The first flight in NASA's New Millennium Program, Deep Space 1 was launched October 24, 1998 aboard a Boeing Delta 7326 rocket from Cape Canaveral Air Station, FL. Deep Space 1 successfully completed and exceeded its mission objectives in July 1999 and flew by a near-Earth asteroid, Braille (1992 KD), in September 1999. http://photojournal.jpl.nasa.gov/catalog/PIA04604

  4. Extending Ion Engine Technology to NEXT and Beyond

    NASA Technical Reports Server (NTRS)

    Domonkos, Matthew T.; Patterson, Michael J.; Foster, John E.; Rawlin, Vince K.; Soulas, George C.; Sovey, James S.; Kovaleski, Scott D.; Roman, Robert F.; Williams, George J., Jr.; Lyons, Valerie J. (Technical Monitor)

    2002-01-01

    Extending ion engine technology beyond the current state-of-the art primary interplanetary electric propulsion system, the 2.3-kW NASA Solar Electric Propulsion Technology and Applications Readiness (NSTAR) system, will require thrusters with improved propellant throughput and total impulse capability. Many of the design choices that culminated in the NSTAR thrusters must be revisited, and their application to next generation ion engine technology must be evaluated. The concept of derating, which was successfully employed in NSTAR, has been applied to the 40 cm NASA Evolutionary Xenon Thruster (NEXT) currently under development at NASA Glenn Research Center (GRC). At 5-kW, NEXT operates with the same average beam current density as NSTAR, and at 10-kW, the peak beam current density is only ten percent greater than NSTAR. The result is that similar Ion optics technology is expected to yield comparable lifetime. Thick-accelerator- grid ion optics are also being tested to realize additional lifetime benefits. A 40-A discharge cathode is being developed for NEXT based on scaling the NSTAR design. Nevertheless, the experiences of the NSTAR ground tests and the thruster on the Deep Space One spacecraft indicate that the discharge cathode wear must be studied experimentally and theoretically to ensure that it meets the lifetime requirements. Although NEXT is in its infancy, investigations have already begun to examine possible modifications to engine design for even higher-power and higher-specific impulse engines. Ion optics using alternate materials such as titanium, graphite, or carbon-carbon composite are currently being investigated due to their low sputter yields at high voltage. To avoid the difficulties encountered using electrodes at high-currents, the use of a microwave-based ion thruster is under investigation for potential high-power ion thruster systems requiring long lifetimes. Additionally, alternative propellants are being considered for applications

  5. Solar Thermal Propulsion Optical Figure Measuring and Rocket Engine Testing

    NASA Technical Reports Server (NTRS)

    Bonometti, Joseph

    1997-01-01

    Solar thermal propulsion has been an important area of study for four years at the Propulsion Research Center. Significant resources have been devoted to the development of the UAH Solar Thermal Laboratory that provides unique, high temperature, test capabilities. The facility is fully operational and has successfully conducted a series of solar thruster shell experiments. Although presently dedicated to solar thermal propulsion, the facility has application to a variety of material processing, power generation, environmental clean-up, and other fundamental research studies. Additionally, the UAH Physics Department has joined the Center in support of an in-depth experimental investigation on Solar Thermal Upper Stage (STUS) concentrators. Laboratory space has been dedicated to the concentrator evaluation in the UAH Optics Building which includes a vertical light tunnel. Two, on-going, research efforts are being sponsored through NASA MSFC (Shooting Star Flight Experiment) and the McDonnell Douglas Corporation (Solar Thermal Upper Stage Technology Ground Demonstrator).

  6. Main propulsion system test requirements for the two-engine Shuttle-C

    NASA Technical Reports Server (NTRS)

    Lynn, E. E.; Platt, G. K.

    1989-01-01

    The Shuttle-C is an unmanned cargo carrying derivative of the space shuttle with optional two or three space shuttle main engines (SSME's), whereas the shuttle has three SSME's. Design and operational differences between the Shuttle-C and shuttle were assessed to determine requirements for additional main propulsion system (MPS) verification testing. Also, reviews were made of the shuttle main propulsion test program objectives and test results and shuttle flight experience. It was concluded that, if significant MPS modifications are not made beyond those currently planned, then main propulsion system verification can be concluded with an on-pad flight readiness firing.

  7. Analysis of the effect of engine characteristics on the external aerodynamics of STOL wing propulsion systems

    NASA Technical Reports Server (NTRS)

    Albers, J. A.

    1972-01-01

    The effects of engine presssure ratio, engine size, and engine location on the pressure distribution, lift coefficient, and flow field of a STOL wing propulsion system are presented. The flow variables of the engines are included in the two-dimensional potential flow analysis by considering the effects of mass flow coefficient at the engine inlet and thrust coefficient at the engine exit. A functional relation between these coefficients and engine pressure ratio is given. The results of this study indicate that the effect of engine pressure ratio on the external aerodynamics is a function of engine location. For engines located on the bottom of the wing, the highest pressure ratio engine resulted in the highest lift coefficient. For engines located on the top of the wing, the lowest pressure ratio engine resulted in the highest lift coefficient.

  8. Environmental Testing of the NEXT PM1 Ion Engine

    NASA Technical Reports Server (NTRS)

    Synder, John S.; Anderson, John R.; VanNoord, Jonathan L.; Soulas, George C.

    2008-01-01

    The NEXT propulsion system is an advanced ion propulsion system presently under development that is oriented towards robotic exploration of the solar system using solar electric power. The Prototype Model engine PM1 was subjected to qualification-level environmental testing to demonstrate compatibility with environments representative of anticipated mission requirements. Random vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10.0 Grms for 2 minutes in each of three axes. Thermal-vacuum testing included a deep cold soak of the engine to temperatures of -168 C and thermal cycling from -120 to 203 C. Although the testing was largely successful, several issues were identified including the fragmentation of potting cement on the discharge and neutralizer cathode heater terminations during vibration which led to abbreviated thermal testing, and generation of particulate contamination from manufacturing processes and engine materials. Thruster performance was nominal throughout the test program, with minor variations in some engine operating parameters likely caused by facility effects. In general, the NEXT PM1 engine and the breadboard gimbal were found to be well-designed against environmental requirements based on the results reported herein. After resolution of the findings from this test program the hardware environmental qualification program can proceed with confidence.

  9. 14 CFR 34.62 - Test procedure (propulsion engines).

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... FUEL VENTING AND EXHAUST EMISSION REQUIREMENTS FOR TURBINE ENGINE POWERED AIRPLANES Test Procedures for Engine Exhaust Gaseous Emissions (Aircraft and Aircraft Gas Turbine Engines) § 34.62 Test...

  10. Thermal and Environmental Barrier Coating Development for Advanced Propulsion Engine Systems

    NASA Technical Reports Server (NTRS)

    Zhu, Dongming; Miller, Robert A.; Fox, Dennis S.

    2008-01-01

    Ceramic thermal and environmental barrier coatings (TEBCs) are used in gas turbine engines to protect engine hot-section components in the harsh combustion environments, and extend component lifetimes. Advanced TEBCs that have significantly lower thermal conductivity, better thermal stability and higher toughness than current coatings will be beneficial for future low emission and high performance propulsion engine systems. In this paper, ceramic coating design and testing considerations will be described for turbine engine high temperature and high-heat-flux applications. Thermal barrier coatings for metallic turbine airfoils and thermal/environmental barrier coatings for SiC/SiC ceramic matrix composite (CMC) components for future supersonic aircraft propulsion engines will be emphasized. Further coating capability and durability improvements for the engine hot-section component applications can be expected by utilizing advanced modeling and design tools.

  11. Performance and optimization of a derated ion thruster for auxiliary propulsion

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Foster, John E.

    1991-01-01

    The characteristics and implications of use of a derated ion thruster for north-south stationkeeping (NSSK) propulsion are discussed. A derated thruster is a 30 cm diameter primary propulsion ion thruster operated at highly throttled conditions appropriate to NSSK functions. The performance characteristics of a 30 cm ion thruster are presented, emphasizing throttled operation at low specific impulse and high thrust-to-power ratio. Performance data and component erosion are compared to other NSSK ion thrusters. Operations benefits derived from the performance advantages of the derated approach are examined assuming an INTELSAt 7-type spacecraft. Minimum ground test facility pumping capabilities required to maintain facility enhanced accelerator grid erosion at acceptable levels in a lifetest are quantified as a function of thruster operating condition. Approaches to reducing the derated thruster mass and volume are also discussed.

  12. Subsonic flight test evaluation of a propulsion system parameter estimation process for the F100 engine

    NASA Technical Reports Server (NTRS)

    Orme, John S.; Gilyard, Glenn B.

    1992-01-01

    Integrated engine-airframe optimal control technology may significantly improve aircraft performance. This technology requires a reliable and accurate parameter estimator to predict unmeasured variables. To develop this technology base, NASA Dryden Flight Research Facility (Edwards, CA), McDonnell Aircraft Company (St. Louis, MO), and Pratt & Whitney (West Palm Beach, FL) have developed and flight-tested an adaptive performance seeking control system which optimizes the quasi-steady-state performance of the F-15 propulsion system. This paper presents flight and ground test evaluations of the propulsion system parameter estimation process used by the performance seeking control system. The estimator consists of a compact propulsion system model and an extended Kalman filter. The extended Laman filter estimates five engine component deviation parameters from measured inputs. The compact model uses measurements and Kalman-filter estimates as inputs to predict unmeasured propulsion parameters such as net propulsive force and fan stall margin. The ability to track trends and estimate absolute values of propulsion system parameters was demonstrated. For example, thrust stand results show a good correlation, especially in trends, between the performance seeking control estimated and measured thrust.

  13. A data acquisition and storage system for the ion auxiliary propulsion system cyclic thruster test

    NASA Technical Reports Server (NTRS)

    Hamley, John A.

    1989-01-01

    A nine-track tape drive interfaced to a standard personal computer was used to transport data from a remote test site to the NASA Lewis mainframe computer for analysis. The Cyclic Ground Test of the Ion Auxiliary Propulsion System (IAPS), which successfully achieved its goal of 2557 cycles and 7057 hr of thrusting beam on time generated several megabytes of test data over many months of continuous testing. A flight-like controller and power supply were used to control the thruster and acquire data. Thruster data was converted to RS232 format and transmitted to a personal computer, which stored the raw digital data on the nine-track tape. The tape format was such that with minor modifications, mainframe flight data analysis software could be used to analyze the Cyclic Ground Test data. The personal computer also converted the digital data to engineering units and displayed real time thruster parameters. Hardcopy data was printed at a rate dependent on thruster operating conditions. The tape drive provided a convenient means to transport the data to the mainframe for analysis, and avoided a development effort for new data analysis software for the Cyclic test. This paper describes the data system, interfacing and software requirements.

  14. An ion thruster module for primary propulsion systems.

    NASA Technical Reports Server (NTRS)

    King, H. J.; Poeschel, R. L.

    1972-01-01

    The development of a 30 cm thruster module having the operational characteristics, weight, and structural integrity consistent with flight hardware is described. Elements of the program discussed in this paper are selection of an ion optical system design, development of the discharge chamber and its control, and the results of extensive performance mapping tests. The thruster system operates at 2750 sec specific impulse at 69% over-all efficiency and can be throttled from 2.0 A to 0.16 A beam current with a control system requiring a single electrical input. The 1 kV ion beam is formed by a high perveance, two grid ion optical system.

  15. System-Engineering Methods and Design Decisions for the Mirror Fusion Propulsion System (MFPS)

    NASA Astrophysics Data System (ADS)

    Deveny, Marc E.; Carpenter, Scott A.

    1994-07-01

    We describe the design trades and rationale supporting development of a continuous-thrusting space-fusion-propulsion system called the Mirror Fusion Propulsion System (MFPS). The MFPS is the result of an earlier design study to adapt and optimize a terrestrial fusion reactor for propulsion in space. In this paper, we focus on the configuration trades that are necessary to make top-level design decisions. Configuration trades include the fusion reactor configuration, fuel combinations (fuel mix and fuel-pellet shelling), plasma temperature, reduced-electron-temperature operating mode, magnetic-field-ripple, electrically-conducting-wall stabilization, superconductor technology and cooling mode (closed-cycle cryocooler or LH2-propellant cooled), and many others. To qualitatively sort through all of these trades and identify directions for further improvement in performance, we developed and applied three distinct design principles useful to adapt, and then optimize, terrestrial fusion reactor configurations for propulsion in space. To quantitatively optimize MFPS, we developed an engineering-design tool that embeds the User in all phases of the design. This Tool is called IDEAs (Integrated Design Environment Algorithms) and it allows the systems engineer to ``see'' several varying results simultaneously. IDEAs converts the top-level systems design into a much easier task. The decision flow results in an advanced space propulsion system with a 500-tonne dry-engine, 4-kWthrust / kgengine specific power, and 4-full-power-year (FPY) design end of life (EOL).

  16. A Lunar-Based Spacecraft Propulsion Concept - The Ion Beam Sail

    NASA Technical Reports Server (NTRS)

    Brown, Ian G.; Lane, John E.; Youngquist, Robert C.

    2006-01-01

    We describe a concept for spacecraft propulsion by means of an energetic ion beam, with the ion source fixed at the spacecraft starting point (e.g., a lunar-based ion beam generator) and not onboard the vessel. This approach avoids the substantial mass penalty associated with the onboard ion source and power supply hardware, and vastly more energetic ion beam systems can be entertained. We estimate the ion beam parameters required for various scenarios, and consider some of the constraints limiting the concept. We find that the "ion beam sail' approach can be viable and attractive for journey distances not too great, for example within the Earth-Moon system, and could potentially provide support for journeys to the inner planets.

  17. A lunar-based spacecraft propulsion concept—The ion beam sail

    NASA Astrophysics Data System (ADS)

    Brown, Ian G.; Lane, John E.; Youngquist, Robert C.

    2007-05-01

    We describe a concept for spacecraft propulsion by means of an energetic ion beam, with the ion source fixed at the spacecraft starting point (e.g., a lunar-based ion beam generator) and not onboard the vessel. This approach avoids the substantial mass penalty associated with the onboard ion source and power supply hardware, and vastly more energetic ion beam systems can be entertained. We estimate the ion beam parameters required for various scenarios and consider some of the constraints limiting the concept. We find that the “ion beam sail” approach can be viable and attractive for journey distances not too great, for example, within the Earth Moon system, and could potentially provide support for journeys to the inner planets.

  18. A Study on Aircraft Engine Control Systems for Integrated Flight and Propulsion Control

    NASA Astrophysics Data System (ADS)

    Yamane, Hideaki; Matsunaga, Yasushi; Kusakawa, Takeshi

    A flyable FADEC system engineering model incorporating Integrated Flight and Propulsion Control (IFPC) concept is developed for a highly maneuverable aircraft and a fighter-class engine. An overview of the FADEC system and functional assignments for its components such as the Engine Control Unit (ECU) and the Integrated Control Unit (ICU) are described. Overall system reliability analysis, convex analysis and multivariable controller design for the engine, fault detection/redundancy management, and response characteristics of a fuel system are addressed. The engine control performance of the FADEC is demonstrated by hardware-in-the-loop simulation for fast acceleration and thrust transient characteristics.

  19. The 8-CM ion thruster characterization. [mercury ion engine

    NASA Technical Reports Server (NTRS)

    Wessel, F. J.; Williamson, W. S.

    1983-01-01

    The performance capabilities of the 8 cm diameter mercury ion thruster were increased by modifying the thruster operating parameters and component hardware. The initial performance levels, representative of the Hughes/NASA Lewis Research Center Ion Auxiliary Propulsion Subsystem (IAPS) thruster, were raised from the baseline values of thrust, T = 5 mN, and specific impulse, I sub sp = 2,900s, to thrust, T = 25 mN and specific impulse, I sub sp = 4,300 s. Performance characteristics including estmates of the erosion rates of various component surfaces are presented.

  20. Cryogenic upper stage propulsion: RL10 and derivative engines

    NASA Technical Reports Server (NTRS)

    Brown, James R.

    1991-01-01

    The capabilities and characteristics of the RL10 rocket engine are examined. The engine model history is presented. The RL10 derivatives are also outlined. The presentation is represented by viewgraphs.

  1. 14 CFR 34.62 - Test procedure (propulsion engines).

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... Section 34.62 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT FUEL VENTING AND EXHAUST EMISSION REQUIREMENTS FOR TURBINE ENGINE POWERED AIRPLANES Test Procedures for Engine Exhaust Gaseous Emissions (Aircraft and Aircraft Gas Turbine Engines) § 34.62 Test procedure...

  2. Numerical Study of the Propulsive Performance of the Hollow Rotating Detonation Engine with a Laval Nozzle

    NASA Astrophysics Data System (ADS)

    Yao, Songbai; Tang, Xinmeng; Wang, Jianping

    2017-04-01

    The aim of the present paper is to investigate the propulsive performance of the hollow rotating detonation engine (RDE) with a Laval nozzle. Three-dimensional simulations are carried out with a one-step Arrhenius chemistry model. The Laval nozzle is found to improve the propulsive performance of hollow RDE in all respects. The thrust and fuel-based specific impulse are increased up to 12.60 kN and 7484.40 s, respectively, from 6.46 kN and 6720.48 s. Meanwhile, the total mass flow rate increases from 3.63 kg/s to 6.68 kg/s. Overall, the Laval nozzle significantly improves the propulsive performance of the hollow RDE and makes it a promising model among detonation engines.

  3. Thermal and Environmental Barrier Coatings for Advanced Propulsion Engine Systems

    NASA Technical Reports Server (NTRS)

    Zhu, Dong-Ming; Miller, Robert A.

    2004-01-01

    Ceramic thermal and environmental barrier coatings (TEBCs) are used in gas turbine engines to protect engine hot-section components in the harsh combustion environments, and extend component lifetimes. For future high performance engines, the development of advanced ceramic barrier coating systems will allow these coatings to be used to simultaneously increase engine operating temperature and reduce cooling requirements, thereby leading to significant improvements in engine power density and efficiency. In order to meet future engine performance and reliability requirements, the coating systems must be designed with increased high temperature stability, lower thermal conductivity, and improved thermal stress and erosion resistance. In this paper, ceramic coating design and testing considerations will be described for high temperature and high-heat-flux engine applications in hot corrosion and oxidation, erosion, and combustion water vapor environments. Further coating performance and life improvements will be expected by utilizing advanced coating architecture design, composition optimization, and improved processing techniques, in conjunction with modeling and design tools.

  4. Advanced supersonic propulsion study, phases 3 and 4. [variable cycle engines

    NASA Technical Reports Server (NTRS)

    Allan, R. D.; Joy, W.

    1977-01-01

    An evaluation of various advanced propulsion concepts for supersonic cruise aircraft resulted in the identification of the double-bypass variable cycle engine as the most promising concept. This engine design utilizes special variable geometry components and an annular exhaust nozzle to provide high take-off thrust and low jet noise. The engine also provides good performance at both supersonic cruise and subsonic cruise. Emission characteristics are excellent. The advanced technology double-bypass variable cycle engine offers an improvement in aircraft range performance relative to earlier supersonic jet engine designs and yet at a lower level of engine noise. Research and technology programs required in certain design areas for this engine concept to realize its potential benefits include refined parametric analysis of selected variable cycle engines, screening of additional unconventional concepts, and engine preliminary design studies. Required critical technology programs are summarized.

  5. FJ44 Turbofan Engine Test at NASA Glenn Research Center's Aero-Acoustic Propulsion Laboratory

    NASA Technical Reports Server (NTRS)

    Lauer, Joel T.; McAllister, Joseph; Loew, Raymond A.; Sutliff, Daniel L.; Harley, Thomas C.

    2009-01-01

    A Williams International FJ44-3A 3000-lb thrust class turbofan engine was tested in the NASA Glenn Research Center s Aero-Acoustic Propulsion Laboratory. This report presents the test set-up and documents the test conditions. Farfield directivity, in-duct unsteady pressures, duct mode data, and phased-array data were taken and are reported separately.

  6. A method to estimate weight and dimensions of small aircraft propulsion gas turbine engines: User's guide

    NASA Technical Reports Server (NTRS)

    Hale, P. L.

    1982-01-01

    The weight and major envelope dimensions of small aircraft propulsion gas turbine engines are estimated. The computerized method, called WATE-S (Weight Analysis of Turbine Engines-Small) is a derivative of the WATE-2 computer code. WATE-S determines the weight of each major component in the engine including compressors, burners, turbines, heat exchangers, nozzles, propellers, and accessories. A preliminary design approach is used where the stress levels, maximum pressures and temperatures, material properties, geometry, stage loading, hub/tip radius ratio, and mechanical overspeed are used to determine the component weights and dimensions. The accuracy of the method is generally better than + or - 10 percent as verified by analysis of four small aircraft propulsion gas turbine engines.

  7. Propulsion System and Orbit Maneuver Integration in CubeSats: Trajectory Control Strategies Using Micro Ion Propulsion

    NASA Technical Reports Server (NTRS)

    Hudson, Jennifer; Martinez, Andres; Petro, Andrew

    2015-01-01

    The Propulsion System and Orbit Maneuver Integration in CubeSats project aims to solve the challenges of integrating a micro electric propulsion system on a CubeSat in order to perform orbital maneuvers and control attitude. This represents a fundamentally new capability for CubeSats, which typically do not contain propulsion systems and cannot maneuver far beyond their initial orbits.

  8. Propulsion technology needs for advanced space transportation systems. [orbit maneuvering engine (space shuttle), space shuttle boosters

    NASA Technical Reports Server (NTRS)

    Gregory, J. W.

    1975-01-01

    Plans are formulated for chemical propulsion technology programs to meet the needs of advanced space transportation systems from 1980 to the year 2000. The many possible vehicle applications are reviewed and cataloged to isolate the common threads of primary propulsion technology that satisfies near term requirements in the first decade and at the same time establish the technology groundwork for various potential far term applications in the second decade. Thrust classes of primary propulsion engines that are apparent include: (1) 5,000 to 30,000 pounds thrust for upper stages and space maneuvering; and (2) large booster engines of over 250,000 pounds thrust. Major classes of propulsion systems and the important subdivisions of each class are identified. The relative importance of each class is discussed in terms of the number of potential applications, the likelihood of that application materializing, and the criticality of the technology needed. Specific technology programs are described and scheduled to fulfill the anticipated primary propulsion technology requirements.

  9. Four-cycle engine for marine propulsion device

    SciTech Connect

    Suzuki, T.

    1986-01-07

    This patent describes a lubricating system for an internal combustion engine including a camshaft rotated by the engine. The lubricating system includes a lubricant pump for delivering lubricant under pressure to at least some components of the engine. The improvement in such a system consists of porous means adapted to engage a surface of the camshaft to be lubricated and means for delivering lubricant from the lubricant pump to the porous means.

  10. Additional Mission Applications for NASA's 13.3-kW Ion Propulsion System

    NASA Technical Reports Server (NTRS)

    Snyder, John Steven; Manzella, David; Lisman, Doug; Lock, Robert E.; Nicholas, Austin; Woolley, Ryan

    2016-01-01

    NASA's Space Technology Mission Directorate has been recently developing critical technologies for high-power solar electric propulsion (SEP), including large deployable solar array structures and high-power electric propulsion components. An ion propulsion system based on these developments has been considered for many SEP technology demonstration missions, including the Asteroid Redirect Robotic Mission (ARRM) concept. These studies and the highpower SEP technology developments have generated excitement within NASA about the use of the ARRM ion propulsion system design for other types of potential missions. One application of interest is for Mars missions, especially with the types of orbiters now under consideration for flights in the early 2020's to replace the aging Mars Reconnaissance Orbiter. High-power SEP can deliver large payloads to Mars with many additional capabilities, including large orbital plane changes and roundtrip missions, compared to chemically-propelled spacecraft. Another application for high-power SEP is for exo-planet observation missions, where a large starshade spacecraft would need to be repositioned with respect to its companion telescope relatively frequently and rapidly. SEP is an enabling technology for the ambitious science goals of these types of missions. This paper will discuss the benefits of high-power SEP for these concepts based on the STMD technologies now under development.

  11. Quiet Clean Short-Haul Experimental Engine (QCSEE) Over-The-Wing (OTW) propulsion system test report. Volume 2: Aerodynamics and performance. [engine performance tests to define propulsion system performance on turbofan engines

    NASA Technical Reports Server (NTRS)

    1978-01-01

    The design and testing of the over the wing engine, a high bypass, geared turbofan engine, are discussed. The propulsion system performance is examined for uninstalled performance and installed performance. The fan aerodynamic performance and the D nozzle and reverser thrust performance are evaluated.

  12. The convertible engine: A dual-mode propulsion system

    NASA Technical Reports Server (NTRS)

    Mcardle, Jack G.

    1988-01-01

    A variable inlet guide vane (VIGV) convertible engine that could be used to power future high-speed rotorcraft was tested on an outdoor stand. The engine ran stably and smoothly in the turbofan, turboshaft, and dual (combined fan and shaft) power modes. In the turbofan mode with the VIGV open, fuel consumption was comparable to that of a conventional turbofan engine. In the turboshaft mode with the VIGV closed, fuel consumption was higher than that of present turboshaft engines because power was wasted in churning fan-tip air flow. In dynamic performance tests with a specially built digital engine control and using a waterbrake dynamometer for shaft load, the engine responded effectively to large steps in thrust command and shaft torque. Previous mission analyses of a conceptual X-wing rotorcraft capable of 400-knot cruise speed were revised to account for more fan-tip churning power loss that was originally estimated. The calculations confirm that using convertible engines rather than separate life and cruise engines would result in a smaller, lighter craft with lower fuel use and direct operating cost.

  13. Operationally Efficient Propulsion System Study (OEPSS) Data Book. Volume 8; Integrated Booster Propulsion Module (BPM) Engine Start Dynamics

    NASA Technical Reports Server (NTRS)

    Kemp, Victoria R.

    1992-01-01

    A fluid-dynamic, digital-transient computer model of an integrated, parallel propulsion system was developed for the CDC mainframe and the SUN workstation computers. Since all STME component designs were used for the integrated system, computer subroutines were written characterizing the performance and geometry of all the components used in the system, including the manifolds. Three transient analysis reports were completed. The first report evaluated the feasibility of integrated engine systems in regards to the start and cutoff transient behavior. The second report evaluated turbopump out and combined thrust chamber/turbopump out conditions. The third report presented sensitivity study results in staggered gas generator spin start and in pump performance characteristics.

  14. The systems engineering upgrade intiative at NASA's Jet Propulsion Laboratory

    NASA Technical Reports Server (NTRS)

    Jones, Ross M.

    2005-01-01

    JPL is implementing an initiative to significantly upgrade our systems engineering capabilities. This Systems Engineering Upgrade Initiative [SUI] has been authorized by the highest level technical management body of JPL and is sponsored with internal funds. The SUI objective is to upgrade system engineering at JPL to a level that is world class, professional and efficient compared to the FY04/05 baseline. JPL system engineering, along with the other engineering disciplines, is intended to support optimum designs; controlled and efficient implementations; and high quality, reliable, cost effective products. SUI technical activities are categorized into those dealing with people, process and tools. The purpose of this paper is to describe the rationale, objectives/plans and current status of the JPL SUI.

  15. The systems engineering upgrade intiative at NASA's Jet Propulsion Laboratory

    NASA Technical Reports Server (NTRS)

    Jones, Ross M.

    2005-01-01

    JPL is implementing an initiative to significantly upgrade our systems engineering capabilities. This Systems Engineering Upgrade Initiative [SUI] has been authorized by the highest level technical management body of JPL and is sponsored with internal funds. The SUI objective is to upgrade system engineering at JPL to a level that is world class, professional and efficient compared to the FY04/05 baseline. JPL system engineering, along with the other engineering disciplines, is intended to support optimum designs; controlled and efficient implementations; and high quality, reliable, cost effective products. SUI technical activities are categorized into those dealing with people, process and tools. The purpose of this paper is to describe the rationale, objectives/plans and current status of the JPL SUI.

  16. Ion-gap sensing for engine control

    SciTech Connect

    1995-09-01

    This article reports that in addition to detecting misfire to conform with California onboard diagnostic (OBD II) regulations, Delco Electronics and Mecel AB engineers are looking at ion-gap sensing to control knock, A/F ratio, and other possible engine control parameters. The combustion of fuel in an engine cylinder produces ions. Detection of those ions by the spark plug (ion-gap sensing), and use of the resulting ion currents, has been employed in engine management systems since 1988. Saab introduced the first application, for cam-phase sensing. The main driving force for ion-gap sensing is OBD II requirements for 100% misfire detection at all speeds and loads. The technique has been expanded in subsequent applications to include misfire, knock, and pre-ignition detection and control, and more recently in combustion-ion detection using a capacitance-type, ion-current measurement method. Use of the ion current`s wave shape to control knock allows elimination of the separate piezoelectric type (PZT) sensor. Future applications could provide additional engine-control features including air/fuel ratio measurement and control.

  17. Assessments of Hollow Cathode Wear in the Xenon Ion Propulsion System (XIPs(c)) by Numerical Analyses and Wear Tests

    NASA Technical Reports Server (NTRS)

    Mikellides, Ioannis G.; Katz, Ira; Goebel, Dan M.; Polk, James E.

    2008-01-01

    The standard approach presently followed by NASA to qualify electric propulsion for the required mission throughput has been based largely on life tests, which can be costly and time consuming. Revised electric propulsion lifequalification approaches are being formulated that combine analytical and/or computational methods with (shorter-duration) wear tests. As a model case, a wear test is being performed at JPL to assess the lifetime of the discharge hollow cathode in the Xenon Ion Propulsion System (XIPS(c)), a 25-cm ion engine developed by L-3 Communications Electron Technologies, Inc. for commercial applications. Wear and plasma data accumulated throughout this life-assessment program are being used to validate the existing 2-D hollow cathode code OrCa2D. We find that the OrCa2D steady-state solution predicts very well the time-averaged plasma data and the keeper voltage after 5500 hrs of operation in high-power mode. When the wave motion that occurs naturally in these devices is accounted for, based on an estimate of the maximum wave amplitude, the molybdenum-keeper erosion profile observed in the XIPS(c) discharge cathode is also reproduced within a factor of two of the observation. When the same model is applied to predict the erosion of a tantalum keeper we find that erosion is reduced by more than two orders of magnitude compared to the molybdenum keeper due the significantly lower sputtering yield of tantalum. A tantalum keeper would therefore allow keeper lifetimes that greatly exceed the present requirements for deep-space robotic missions considered by NASA. Moreover, such large reduction of the erosion renders the largest uncertainties in the models, which are associated with the wave amplitude estimates and the electron transport model, negligible.

  18. Assessments of Hollow Cathode Wear in the Xenon Ion Propulsion System (XIPs(c)) by Numerical Analyses and Wear Tests

    NASA Technical Reports Server (NTRS)

    Mikellides, Ioannis G.; Katz, Ira; Goebel, Dan M.; Polk, James E.

    2008-01-01

    The standard approach presently followed by NASA to qualify electric propulsion for the required mission throughput has been based largely on life tests, which can be costly and time consuming. Revised electric propulsion lifequalification approaches are being formulated that combine analytical and/or computational methods with (shorter-duration) wear tests. As a model case, a wear test is being performed at JPL to assess the lifetime of the discharge hollow cathode in the Xenon Ion Propulsion System (XIPS(c)), a 25-cm ion engine developed by L-3 Communications Electron Technologies, Inc. for commercial applications. Wear and plasma data accumulated throughout this life-assessment program are being used to validate the existing 2-D hollow cathode code OrCa2D. We find that the OrCa2D steady-state solution predicts very well the time-averaged plasma data and the keeper voltage after 5500 hrs of operation in high-power mode. When the wave motion that occurs naturally in these devices is accounted for, based on an estimate of the maximum wave amplitude, the molybdenum-keeper erosion profile observed in the XIPS(c) discharge cathode is also reproduced within a factor of two of the observation. When the same model is applied to predict the erosion of a tantalum keeper we find that erosion is reduced by more than two orders of magnitude compared to the molybdenum keeper due the significantly lower sputtering yield of tantalum. A tantalum keeper would therefore allow keeper lifetimes that greatly exceed the present requirements for deep-space robotic missions considered by NASA. Moreover, such large reduction of the erosion renders the largest uncertainties in the models, which are associated with the wave amplitude estimates and the electron transport model, negligible.

  19. Advanced Supersonic Technology Study: Engine Program Summary. Supersonic Propulsion: 1971 to 1976

    NASA Technical Reports Server (NTRS)

    Krebs, J. N.

    1976-01-01

    Sustained supersonic cruise propulsion systems for military applications are studied. The J79-5 in the Mach 2 B-58; YJ93 in the Mach 3.0 B-70 and the current F101 in the B-1, are all examples of military propulsion systems and airplanes operated at sustained supersonic cruise speeds. The Mach 2.7 B2707 transport powered by GE4 turbojet engines was the only non-military, sustained supersonic cruise vehicle intended for commercial passenger service.

  20. Validation of an Integrated Airframe and Turbofan Engine Simulation for Evaluation of Propulsion Control Modes

    NASA Technical Reports Server (NTRS)

    Litt, Jonathan S.; Sowers, T Shane; Liu, Yuan; Owen, A. Karl; Guo, Ten-Huei

    2015-01-01

    The National Aeronautics and Space Administration (NASA) has developed independent airframe and engine models that have been integrated into a single real-time aircraft simulation for piloted evaluation of propulsion control algorithms. In order to have confidence in the results of these evaluations, the integrated simulation must be validated to demonstrate that its behavior is realistic and that it meets the appropriate Federal Aviation Administration (FAA) certification requirements for aircraft. The paper describes the test procedures and results, demonstrating that the integrated simulation generally meets the FAA requirements and is thus a valid testbed for evaluation of propulsion control modes.

  1. Ground Test Facility for Propulsion and Power Modes of Nuclear Engine Operation

    SciTech Connect

    Michael, WILLIAMS

    2004-11-22

    Existing DOE Ground Test Facilities have not been used to support nuclear propulsion testing since the Rover/NERVA programs of the 1960's. Unlike the Rover/NERVA programs, DOE Ground Test facilities for space exploration enabling nuclear technologies can no longer be vented to the open atmosphere. The optimal selection of DOE facilities and accompanying modifications for confinement and treatment of exhaust gases will permit the safe testing of NASA Nuclear Propulsion and Power devices involving variable size and source nuclear engines for NASA Jupiter Icy Moon Orbiter (JIMO) and Commercial Space Exploration Missions with minimal cost, schedule and environmental impact. NASA site selection criteria and testing requirements are presented.

  2. Design and Development of the MITEE-B Bi-Modal Nuclear Propulsion Engine

    NASA Astrophysics Data System (ADS)

    Paniagua, John C.; Powell, James R.; Maise, George

    2003-01-01

    Previous studies of compact, ultra-lightweight high performance nuclear thermal propulsion engines have concentrated on systems that only deliver high thrust. However, many potential missions also require substantial amounts of electric power. Studies of a new, very compact and lightweight bi-modal nuclear engine that provides both high propulsive thrust and high electric power for planetary science missions are described. The design is a modification of the MITEE nuclear thermal engine concept that provided only high propulsive thrust. In the new design, MITEE-B, separate closed cooling circuits are incorporated into the reactor, which transfers useful amounts of thermal energy to a small power conversion system that generates continuous electric power over the full life of the mission, even when the engine is not delivering propulsive thrust. Two versions of the MITEE-B design are described and analyzed. Version 1 generates 1 kW(e) of continuous power for control of the spacecraft, sensors, data transmission, etc. This power level eliminates the need for RTG's on missions to the outer planets, and allowing considerably greater operational capability for the spacecraft. This, plus its high thrust and high specific impulse propulsive capabilities, makes MITEE-B very attractive for such missions. In Version 2, of MITEE-B, a total of 20 kW(e) is generated, enabling the use of electric propulsion. The combination of high open cycle propulsion thrust (20,000 Newtons) with a specific impulse of ~1000 seconds for short impulse burns, and long term (months to years), electric propulsion greatly increases MITEE's ΔV capability. Version 2 of MITEE-B also enables the production and replenishment of H2 propellant using in-situ resources, such as electrolysis of water from the ice sheet on Europa and other Jovian moons. This capability would greatly increase the ΔV available for certain planetary science missions. The modifications to the MITEE multiple pressure tube

  3. High Pressure Regenerative Turbine Engine: 21st Century Propulsion

    NASA Technical Reports Server (NTRS)

    Lear, W. E.; Laganelli, A. L.; Senick, Paul (Technical Monitor)

    2001-01-01

    A novel semi-closed cycle gas turbine engine was demonstrated and was found to meet the program goals. The proof-of-principle test of the High Pressure Regenerative Turbine Engine produced data that agreed well with models, enabling more confidence in designing future prototypes based on this concept. Emission levels were significantly reduced as predicted as a natural attribute of this power cycle. Engine testing over a portion of the operating range allowed verification of predicted power increases compared to the baseline.

  4. Impact of Ion Propulsion on Performance, Design, Testing and Operation of a Geosynchronous Spacecraft

    DTIC Science & Technology

    1990-06-01

    11 Title (Include Security Classification) IMPACT OF ION PROPULSION ON PERFORMANCE, DESIGN, TESTING AND OPERATION OF A GEOSYNCHRONOUS SATELLITE 12...June 1990 I 11 16 Supplementary Notation The views expressed in this thesis are those of the author and do not reflect the official policy or position...Abstract 21 Abstract Security Classification N unclassified/unlimited 11 same as report IJ DTIC users Unclassified 22a Name of Responsible Individual

  5. A Survey of Xenon Ion Sputter Yield Data and Fits Relevant to Electric Propulsion Spacecraft Integration

    NASA Technical Reports Server (NTRS)

    Yim, John T.

    2017-01-01

    A survey of low energy xenon ion impact sputter yields was conducted to provide a more coherent baseline set of sputter yield data and accompanying fits for electric propulsion integration. Data uncertainties are discussed and different available curve fit formulas are assessed for their general suitability. A Bayesian parameter fitting approach is used with a Markov chain Monte Carlo method to provide estimates for the fitting parameters while characterizing the uncertainties for the resulting yield curves.

  6. Advanced Space Propulsion Based on Vacuum (Spacetime Metirc) Engineering

    NASA Astrophysics Data System (ADS)

    Puthoff, H.

    A theme that has come to the fore in advanced planning for long-range space exploration is the concept that empty space itself (the quantum vacuum, or spacetime metric) might be engineered so as to provide energy/thrust for future space vehicles. Although far-reaching, such a proposal is solidly grounded in modern physical theory, and therefore the possibility that matter/ vacuum interactions might be engineered for space-flight applications is not a priori ruled out [1]. As examples, the current development of theoretical physics addresses such topics as warp drives, traversable wormholes and time machines that provide for such vacuum engineering possibilities [2-6]. We provide here from a broad perspective the physics and correlates/ consequences of the engineering of the spacetime metric.

  7. Test of a CMC liquid propulsion rocket engine turbine stator

    NASA Astrophysics Data System (ADS)

    Berque, J.; Georges, J. M.

    The succesful testing of two Ceramic Matrix Composite (CMC) stators in conditions fully representative of a cryogenic rocket engine turbine is reported. Both stators possessed the same overall geometry as the actual metallic component used in the turbine second stage of the HM7 engine. They sustained several mid duration tests devoted to combustion parameter tuning and composite material behavior control, before long duration runs, representative of the engine duty cycle. The main objectives were to increase the gas temperature above 1600 K, and to simulate the thermal shock that occurs during the chilldown at the end of the combustion phase, in order to provide a good insight of the benefits in terms of performance, mass, durability and reliability associated with CMC application to a large set of aerospace engine turbines.

  8. Test of a CMC liquid propulsion rocket engine turbine stator

    NASA Astrophysics Data System (ADS)

    Berque, J.; Georges, J. M.

    1992-07-01

    During the first semester of 1991, SEP successfully tested two Ceramic Matrix Composite stators in conditions fully representative of a cryogenic rocket engine turbine. Both stators possessed the same overall geometry as the actual metallic component used in the turbine second stage of the HM7 engine. They sustained several mid-duration tests devoted to combustion parameter tuning and composite material behavior control, before long duration runs, representative of the engine duty cycle. The main objectives were to increase the gas temperature above 1600 K, and to simulate the thermal shock that occurs during the chilldown at the end of the combustion phase, in order to provide a good insight of the benefits in terms of performance, mass, durability and reliability associated with CMC application to a large set of aerospace engine turbines.

  9. Performance Evaluation of the T6 Ion Engine

    NASA Technical Reports Server (NTRS)

    Snyder, John Steven; Goebel, Dan M.; Hofer, Richard R.; Polk, James E.; Wallace, Neil C.; Simpson, Huw

    2010-01-01

    The T6 ion engine is a 22-cm diameter, 4.5-kW Kaufman-type ion thruster produced by QinetiQ, Ltd., and is baselined for the European Space Agency BepiColombo mission to Mercury and is being qualified under ESA sponsorship for the extended range AlphaBus communications satellite platform. The heritage of the T6 includes the T5 ion thruster now successfully operating on the ESA GOCE spacecraft. As a part of the T6 development program, an engineering model thruster was subjected to a suite of performance tests and plume diagnostics at the Jet Propulsion Laboratory. The engine was mounted on a thrust stand and operated over its nominal throttle range of 2.5 to 4.5 kW. In addition to the typical electrical and flow measurements, an E x B mass analyzer, scanning Faraday probe, thrust vector probe, and several near-field probes were utilized. Thrust, beam divergence, double ion content, and thrust vector movement were all measured at four separate throttle points. The engine performance agreed well with published data on this thruster. At full power the T6 produced 143 mN of thrust at a specific impulse of 4120 seconds and an efficiency of 64%; optimization of the neutralizer for lower flow rates increased the specific impulse to 4300 seconds and the efficiency to nearly 66%. Measured beam divergence was less than, and double ion content was greater than, the ring-cusp-design NSTAR thruster that has flown on NASA missions. The measured thrust vector offset depended slightly on throttle level and was found to increase with time as the thruster approached thermal equilibrium.

  10. Performance Evaluation of the T6 Ion Engine

    NASA Technical Reports Server (NTRS)

    Snyder, John Steven; Goebel, Dan M.; Hofer, Richard R.; Polk, James E.; Wallace, Neil C.; Simpson, Huw

    2010-01-01

    The T6 ion engine is a 22-cm diameter, 4.5-kW Kaufman-type ion thruster produced by QinetiQ, Ltd., and is baselined for the European Space Agency BepiColombo mission to Mercury and is being qualified under ESA sponsorship for the extended range AlphaBus communications satellite platform. The heritage of the T6 includes the T5 ion thruster now successfully operating on the ESA GOCE spacecraft. As a part of the T6 development program, an engineering model thruster was subjected to a suite of performance tests and plume diagnostics at the Jet Propulsion Laboratory. The engine was mounted on a thrust stand and operated over its nominal throttle range of 2.5 to 4.5 kW. In addition to the typical electrical and flow measurements, an E x B mass analyzer, scanning Faraday probe, thrust vector probe, and several near-field probes were utilized. Thrust, beam divergence, double ion content, and thrust vector movement were all measured at four separate throttle points. The engine performance agreed well with published data on this thruster. At full power the T6 produced 143 mN of thrust at a specific impulse of 4120 seconds and an efficiency of 64%; optimization of the neutralizer for lower flow rates increased the specific impulse to 4300 seconds and the efficiency to nearly 66%. Measured beam divergence was less than, and double ion content was greater than, the ring-cusp-design NSTAR thruster that has flown on NASA missions. The measured thrust vector offset depended slightly on throttle level and was found to increase with time as the thruster approached thermal equilibrium.

  11. Design of a multivariable integrated control for a supersonic propulsion system. [variable stream control engine

    NASA Technical Reports Server (NTRS)

    Beattie, E. C.

    1980-01-01

    An inlet/engine/nozzle integrated control mode for the propulsion system of an advanced supersonic commercial aircraft was studied. Results show that integration of these control functions can result in both operational and performance benefits for the propulsion system. For example, this integrated control mode may make it possible to minimize the use of inlet bypass doors for shock position control. This may be of benefit to the aircraft as a result of minimizing: (1) bypass bleed drag effects; (2) perturbations to the aircraft resulting from the side thrust effect of the bypass bleeds; and (3) potential unstarts of the inlet. A conceptual integrated control mode was developed which makes use of many cross coupling paths between inlet and engine control variables and inlet and engine sensed variables. A multivariable control design technique based upon linear quadratic regulator theory was applied to designing the feedback gains for this control to allow a simulation evaluation of the benefits of the integrated control mode.

  12. Ice Crystal Icing Engine Testing in the NASA Glenn Research Center's Propulsion Systems Laboratory: Altitude Investigation

    NASA Technical Reports Server (NTRS)

    Oliver, Michael J.

    2014-01-01

    The National Aeronautics and Space Administration (NASA) conducted a full scale ice crystal icing turbofan engine test using an obsolete Allied Signal ALF502-R5 engine in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The test article used was the exact engine that experienced a loss of power event after the ingestion of ice crystals while operating at high altitude during a 1997 Honeywell flight test campaign investigating the turbofan engine ice crystal icing phenomena. The test plan included test points conducted at the known flight test campaign field event pressure altitude and at various pressure altitudes ranging from low to high throughout the engine operating envelope. The test article experienced a loss of power event at each of the altitudes tested. For each pressure altitude test point conducted the ambient static temperature was predicted using a NASA engine icing risk computer model for the given ambient static pressure while maintaining the engine speed.

  13. E-Alerts: Combustion, engines, and propellants. E-mail newsletter

    SciTech Connect

    1999-04-01

    Contents: Combustion and ignition; electric and ion propulsion; fuel and propellant tanks; jet and gas turbine engines; rocket engines and motors; rocket propellants; nuclear propulsion; reciprocation and rotating combustion engines.

  14. A Brief Review of the Need for Robust Smart Wireless Sensor Systems for Future Propulsion Systems, Distributed Engine Controls, and Propulsion Health Management

    NASA Technical Reports Server (NTRS)

    Hunter, Gary W.; Behbahani, Alireza

    2012-01-01

    Smart Sensor Systems with wireless capability operational in high temperature, harsh environments are a significant component in enabling future propulsion systems to meet a range of increasingly demanding requirements. These propulsion systems must incorporate technology that will monitor engine component conditions, analyze the incoming data, and modify operating parameters to optimize propulsion system operations. This paper discusses the motivation towards the development of high temperature, smart wireless sensor systems that include sensors, electronics, wireless communication, and power. The challenges associated with the use of traditional wired sensor systems will be reviewed and potential advantages of Smart Sensor Systems will be discussed. A brief review of potential applications for wireless smart sensor networks and their potential impact on propulsion system operation, with emphasis on Distributed Engine Control and Propulsion Health Management, will be given. A specific example related to the development of high temperature Smart Sensor Systems based on silicon carbide electronics will be discussed. It is concluded that the development of a range of robust smart wireless sensor systems are a foundation for future development of intelligent propulsion systems with enhanced capabilities.

  15. A Brief Review of the Need for Robust Smart Wireless Sensor Systems for Future Propulsion Systems, Distributed Engine Controls, and Propulsion Health Management

    NASA Technical Reports Server (NTRS)

    Hunter, Gary W.; Behbahani, Alireza

    2012-01-01

    Smart Sensor Systems with wireless capability operational in high temperature, harsh environments are a significant component in enabling future propulsion systems to meet a range of increasingly demanding requirements. These propulsion systems must incorporate technology that will monitor engine component conditions, analyze the incoming data, and modify operating parameters to optimize propulsion system operations. This paper discusses the motivation towards the development of high temperature, smart wireless sensor systems that include sensors, electronics, wireless communication, and power. The challenges associated with the use of traditional wired sensor systems will be reviewed and potential advantages of Smart Sensor Systems will be discussed. A brief review of potential applications for wireless smart sensor networks and their potential impact on propulsion system operation, with emphasis on Distributed Engine Control and Propulsion Health Management, will be given. A specific example related to the development of high temperature Smart Sensor Systems based on silicon carbide electronics will be discussed. It is concluded that the development of a range of robust smart wireless sensor systems are a foundation for future development of intelligent propulsion systems with enhanced capabilities.

  16. Mathematical model of marine diesel engine simulator for a new methodology of self propulsion tests

    SciTech Connect

    Izzuddin, Nur; Sunarsih,; Priyanto, Agoes

    2015-05-15

    As a vessel operates in the open seas, a marine diesel engine simulator whose engine rotation is controlled to transmit through propeller shaft is a new methodology for the self propulsion tests to track the fuel saving in a real time. Considering the circumstance, this paper presents the real time of marine diesel engine simulator system to track the real performance of a ship through a computer-simulated model. A mathematical model of marine diesel engine and the propeller are used in the simulation to estimate fuel rate, engine rotating speed, thrust and torque of the propeller thus achieve the target vessel’s speed. The input and output are a real time control system of fuel saving rate and propeller rotating speed representing the marine diesel engine characteristics. The self-propulsion tests in calm waters were conducted using a vessel model to validate the marine diesel engine simulator. The simulator then was used to evaluate the fuel saving by employing a new mathematical model of turbochargers for the marine diesel engine simulator. The control system developed will be beneficial for users as to analyze different condition of vessel’s speed to obtain better characteristics and hence optimize the fuel saving rate.

  17. Mathematical model of marine diesel engine simulator for a new methodology of self propulsion tests

    NASA Astrophysics Data System (ADS)

    Izzuddin, Nur; Sunarsih, Priyanto, Agoes

    2015-05-01

    As a vessel operates in the open seas, a marine diesel engine simulator whose engine rotation is controlled to transmit through propeller shaft is a new methodology for the self propulsion tests to track the fuel saving in a real time. Considering the circumstance, this paper presents the real time of marine diesel engine simulator system to track the real performance of a ship through a computer-simulated model. A mathematical model of marine diesel engine and the propeller are used in the simulation to estimate fuel rate, engine rotating speed, thrust and torque of the propeller thus achieve the target vessel's speed. The input and output are a real time control system of fuel saving rate and propeller rotating speed representing the marine diesel engine characteristics. The self-propulsion tests in calm waters were conducted using a vessel model to validate the marine diesel engine simulator. The simulator then was used to evaluate the fuel saving by employing a new mathematical model of turbochargers for the marine diesel engine simulator. The control system developed will be beneficial for users as to analyze different condition of vessel's speed to obtain better characteristics and hence optimize the fuel saving rate.

  18. An airline study of advanced technology requirements for advanced high speed commercial engines. 3: Propulsion system requirements

    NASA Technical Reports Server (NTRS)

    Sallee, G. P.

    1973-01-01

    The advanced technology requirements for an advanced high speed commercial transport engine are presented. The results of the phase 3 effort cover the requirements and objectives for future aircraft propulsion systems. These requirements reflect the results of the Task 1 and 2 efforts and serve as a baseline for future evaluations, specification development efforts, contract/purchase agreements, and operational plans for future subsonic commercial engines. This report is divided into five major sections: (1) management objectives for commercial propulsion systems, (2) performance requirements for commercial transport propulsion systems, (3) design criteria for future transport engines, (4) design requirements for powerplant packages, and (5) testing.

  19. Energy efficient engine flight propulsion system preliminary analysis and design report

    NASA Technical Reports Server (NTRS)

    Gardner, W. B.

    1979-01-01

    A flight propulsion system preliminary design was established that meets the program goals of at least a 12 percent reduction in thrust specific fuel consumption, at least a five percent reduction in direct operating cost, and one-half the performance deterioration rate of the most efficient current commercial engines. The engine provides a high probability of meeting the 1978 noise rule goal. Smoke and gaseous emissions defined by the EPA proposed standards for engines newly certified after 1 January 1981 are met with the exception of NOx, despite incorporation of all known NOx reduction technology.

  20. The use of gaseous fuels mixtures for SI engines propulsion

    NASA Astrophysics Data System (ADS)

    Flekiewicz, M.; Kubica, G.

    2016-09-01

    Paper presents results of SI engine tests, carried on for different gaseous fuels. Carried out analysis made it possible to define correlation between fuel composition and engine operating parameters. Tests covered various gaseous mixtures: of methane and hydrogen and LPG with DME featuring different shares. The first group, considered as low carbon content fuels can be characterized by low CO2 emissions. Flammability of hydrogen added in those mixtures realizes the function of combustion process activator. That is why hydrogen addition improves the energy conversion by about 3%. The second group of fuels is constituted by LPG and DME mixtures. DME mixes perfectly with LPG, and differently than in case of other hydrocarbon fuels consists also of oxygen makes the stoichiometric mixture less oxygen demanding. In case of this fuel an improvement in engine volumetric and overall engine efficiency has been noticed, when compared to LPG. For the 11% DME share in the mixture an improvement of 2% in the efficiency has been noticed. During the tests standard CNG/LPG feeding systems have been used, what underlines utility value of the research. The stand tests results have been followed by combustion process simulation including exhaust forming and charge exchange.

  1. In-Flight Operation of the Dawn Ion Propulsion System: Status at One Year from the Vesta Rendezvous

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.

    2010-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218 kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide most of the delta V needed for heliocentric transfer to Vesta, orbit capture at Vesta, transfer among Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer among Ceres science orbits. The Dawn ion thruster [I thought we only called it a thruster. Both terms are used in the paper, but I think a replacement of every occurrence of "engine" with "thruster" would be clearer.] design is based on the design validated on NASA's Deep Space 1 (DS1) mission. However, because of the very substantial (11 km/s) delta V requirements for this mission Dawn requires two engines to complete its mission objectives. The power processor units (PPU), digital control and interface units (DCIU) slice boards and the xenon control assembly (XCA) are derivatives of the components used on DS1. The DCIUs and thrust gimbal assemblies (TGA) were developed at the Jet Propulsion Laboratory. The spacecraft was provided by Orbital Sciences Corporation, Sterling, Virginia, and the mission is managed by and operated from the Jet Propulsion Laboratory. Dawn partnered with Germany, Italy and Los Alamos National Laboratory for the science instruments. The mission is led by the principal investigator, Dr. Christopher Russell, from the University of California, Los Angeles. The first 80 days after launch were dedicated to the initial checkout of the spacecraft followed by cruise to Mars. Cruise thrusting leading to a Mars gravity assist began on December 17

  2. In-Flight Operation of the Dawn Ion Propulsion System: Status at One Year from the Vesta Rendezvous

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.

    2010-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218 kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide most of the delta V needed for heliocentric transfer to Vesta, orbit capture at Vesta, transfer among Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer among Ceres science orbits. The Dawn ion thruster [I thought we only called it a thruster. Both terms are used in the paper, but I think a replacement of every occurrence of "engine" with "thruster" would be clearer.] design is based on the design validated on NASA's Deep Space 1 (DS1) mission. However, because of the very substantial (11 km/s) delta V requirements for this mission Dawn requires two engines to complete its mission objectives. The power processor units (PPU), digital control and interface units (DCIU) slice boards and the xenon control assembly (XCA) are derivatives of the components used on DS1. The DCIUs and thrust gimbal assemblies (TGA) were developed at the Jet Propulsion Laboratory. The spacecraft was provided by Orbital Sciences Corporation, Sterling, Virginia, and the mission is managed by and operated from the Jet Propulsion Laboratory. Dawn partnered with Germany, Italy and Los Alamos National Laboratory for the science instruments. The mission is led by the principal investigator, Dr. Christopher Russell, from the University of California, Los Angeles. The first 80 days after launch were dedicated to the initial checkout of the spacecraft followed by cruise to Mars. Cruise thrusting leading to a Mars gravity assist began on December 17

  3. Energy efficient engine: Flight propulsion system final design and analysis. Report, November 1978-August 1983

    SciTech Connect

    Davis, D.Y.; Stearns, E.M.

    1985-08-01

    The Energy Efficient Engine (E3) is a NASA program to create fuel saving technology for future transport engines. The Flight Propulsion System (FPS) is the engine designed to achieve E3 goals. Achieving these goals required aerodynamic, mechanical and system technologies advanced beyond that of current production engines. These technologies were successfully demonstrated in component rigs, a core engine and a turbofan ground test engine. The design and benefits of the FPS are presented. All goals for efficiency, environmental considerations, and economic payoff were met. The FPS has, at maximum cruise, 10.67 km (35,000 ft), M0.8, standard day, a 16.9 percent lower installed specific fuel consumption than a CF6-50C. It provides an 8.6 percent reduction in direct operating cost for a short haul domestic transport and a 16.2 percent reduction for an international long distance transport.

  4. An update of engine system research at the Army Propulsion Directorate

    NASA Technical Reports Server (NTRS)

    Bobula, George A.

    1990-01-01

    The Small Turboshaft Engine Research (STER) program provides a vehicle for evaluating the application of emerging technologies to Army turboshaft engine systems and to investigate related phenomena. Capitalizing on the resources at hand, in the form of both the NASA facilities and the Army personnel, the program goal of developing a physical understanding of engine system dynamics and/or system interactions is being realized. STER entries investigate concepts and components developed both in-house and out-of-house. Emphasis is placed upon evaluations which have evolved from on-going basic research and advanced development programs. Army aviation program managers are also encouraged to make use of STER resources, both people and facilities. The STER personnel have established their reputations as experts in the fields of engine system experimental evaluations and engine system related phenomena. The STER facility has demonstrated its utility in both research and development programs. The STER program provides the Army aviation community the opportunity to perform system level investigations, and then to offer the findings to the entire engine community for their consideration in next generation propulsion systems. In this way results of the fundamental research being conducted to meet small turboshaft engine technology challenges expeditiously find their way into that next generation of propulsion systems.

  5. An update of engine system research at the Army Propulsion Directorate

    NASA Technical Reports Server (NTRS)

    Bobula, George A.

    1990-01-01

    The Small Turboshaft Engine Research (STER) program provides a vehicle for evaluating the application of emerging technologies to Army turboshaft engine systems and to investigate related phenomena. Capitalizing on the resources at hand, in the form of both the NASA facilities and the Army personnel, the program goal of developing a physical understanding of engine system dynamics and/or system interactions is being realized. STER entries investigate concepts and components developed both in-house and out-of-house. Emphasis is placed upon evaluations which evolved from on-going basic research and advanced development programs. Army aviation program managers are also encouraged to make use of STER resources, both people and facilities. The STER personnel have established their reputations as experts in the fields of engine system experimental evaluations and engine system related phenomena. The STER facility has STER program provides the Army aviation community the opportunity to perform system level investigations, and then to offer the findings to the entire engine community for their consideration in next generation propulsion systems. In this way results of the fundamental research being conducted to meet small turboshaft engine technology challenges expeditiously find their way into that next generation of propulsion systems.

  6. Efficient space propulsion engines based on laser ablation

    SciTech Connect

    Phipps, C.R.

    1993-08-01

    Recent results have shown laser momentum transfer coefficients C{sub m} as large as 700 dynes/J from visible and near-infrared laser pulses with heterogeneous targets. Using inexpensive target materials, it is now possible to deliver a 1-tonne satellite from LEO to GEO in 21 days using a 10-kW onboard laser ablation engine, or to maintain several 1-tonne GEO satellites on station from Earth indefinitely using a laser with 100-W average power.

  7. Performance Evaluation of the NEXT Ion Engine

    NASA Technical Reports Server (NTRS)

    Soulas, George C.; Domonkos, Matthew T.; Patterson, Michael J.

    2003-01-01

    The performance test results of three NEXT ion engines are presented. These ion engines exhibited peak specific impulse and thrust efficiency ranges of 4060 4090 s and 0.68 0.69, respectively, at the full power point of the NEXT throttle table. The performance of the ion engines satisfied all project requirements. Beam flatness parameters were significantly improved over the NSTAR ion engine, which is expected to improve accelerator grid service life. The results of engine inlet pressure and temperature measurements are also presented. Maximum main plenum, cathode, and neutralizer pressures were 12,000 Pa, 3110 Pa, and 8540 Pa, respectively, at the full power point of the NEXT throttle table. Main plenum and cathode inlet pressures required about 6 hours to increase to steady-state, while the neutralizer required only about 0.5 hour. Steady-state engine operating temperature ranges throughout the power throttling range examined were 179 303 C for the discharge chamber magnet rings and 132 213 C for the ion optics mounting ring.

  8. Nuclear electric propulsion mission engineering study development program and costs estimates, Phase 2 review

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The results are presented of the second six-month performance period of the Nuclear Electric Propulsion Mission Engineering Study. A brief overview of the program, identifying the study objectives and approach, and a discussion of the program status and schedule are presented. The program results are reviewed and key conclusions to date are summarized. Planned effort for the remainder of the program is reviewed.

  9. Liquid-metal-ion source development for space propulsion at ARC.

    PubMed

    Tajmar, M; Scharlemann, C; Genovese, A; Buldrini, N; Steiger, W; Vasiljevich, I

    2009-04-01

    The Austrian Research Centers have a long history of developing indium Liquid-Metal-Ion Source (LMIS) for space applications including spacecraft charging compensators, SIMS and propulsion. Specifically the application as a thruster requires long-term operation as well as high-current operation which is very challenging. Recently, we demonstrated the operation of a cluster of single LMIS at an average current of 100muA each for more than 4800h and developed models for tip erosion and droplet deposition suggesting that such a LMIS can operate up to 20,000h or more. In order to drastically increase the current, a porous multi-tip source that allows operation up to several mA was developed. Our paper will highlight the problem areas and challenges from our LMIS development focusing on space propulsion applications.

  10. Photographic documentation of the High Power Engine Propulsion HiPEP after a duration test. Also ph

    NASA Technical Reports Server (NTRS)

    2005-01-01

    Photographic documentation of the High Power Engine Propulsion HiPEP after a duration test. Also photographed are the instrumentation and installation articles to reveal post test conditions such as corrosion and pitting.

  11. Coupled multidisciplinary simulation of composite engine structures in propulsion environment

    SciTech Connect

    Chamis, C.C. ); Singhal, S.N. )

    1993-04-01

    A computational simulation procedure is described for the coupled response of multilayered multimaterial composite engine structural components that are subjected to simultaneous multidisciplinary thermal, structural, vibration, and acoustic loading including the effect of hostile environments. The simulation is based on a three-dimensional finite element analysis technique in conjunction with structural mechanics codes and with the acoustic analysis methods. The composite material behavior is assessed at the various composite scales, i.e., the laminate/ply/fiber and matrix constituents, via a nonlinear material characterization model. Sample cases exhibiting nonlinear geometric, material, loading, and environmental behavior of aircraft engine fan blades are presented. Results for deformed shape, vibration frequencies, mode shapes, and acoustic noise emitted from the fan blade are discussed for their coupled effect in hot and humid environments. Results such as acoustic noise for coupled composite-mechanics/heat transfer/structural/vibration/acoustic analyses demonstrate the effectiveness of coupled multidisciplinary computational simulation and the various advantages of composite materials compared to metals.

  12. Antenna Design Method and Performance Improvement of a Micro Ion Engine Using Microwave Discharge

    NASA Astrophysics Data System (ADS)

    Koizumi, Hiroyuki; Kuninaka, Hitoshi

    In this study, we are proposing a novel miniaturized ion engine system µ1. Recently microspacecraft and propulsion system to be installed there have attracted a lot of attentions. To accomplish the miniaturization of spacecraft component, multifunctionalization of devices are key technologies. The ion engine we are proposing here is distributed on microspacecraft and give a number of functions and strong redundancy to the spacecraft. To realize this concept, we introduced a novel idea for an ion engine system. That is to use single plasma source as both ion beam source and neutralizing electron source only by electrical connection. This ion engine system is released from the necessity of a number of neutralizers. Our concept requires a plasma source driven by very low power microwave. Here we proposed an antenna design method for a small plasma source using microwave discharge, and developed a miniaturized ion engine. As a result, the performance of the miniaturized ion engine was improved up to the ion production cost of 240 V and propellant utilization efficiency of 40 % at the input microwave power of 1.0 W and mass flow rate of 0.15 sccm.

  13. A Modular Aero-Propulsion System Simulation of a Large Commercial Aircraft Engine

    NASA Technical Reports Server (NTRS)

    DeCastro, Jonathan A.; Litt, Jonathan S.; Frederick, Dean K.

    2008-01-01

    A simulation of a commercial engine has been developed in a graphical environment to meet the increasing need across the controls and health management community for a common research and development platform. This paper describes the Commercial Modular Aero Propulsion System Simulation (C-MAPSS), which is representative of a 90,000-lb thrust class two spool, high bypass ratio commercial turbofan engine. A control law resembling the state-of-the-art on board modern aircraft engines is included, consisting of a fan-speed control loop supplemented by relevant engine limit protection regulator loops. The objective of this paper is to provide a top-down overview of the complete engine simulation package.

  14. The supersonic fan engine - An advanced concept in supersonic cruise propulsion

    NASA Technical Reports Server (NTRS)

    Franciscus, L. C.

    1981-01-01

    Engine performance and mission studies were conducted for a novel turbofan engine concept incorporating a supersonic through-flow fan, and comparisons were made with two supersonic transport (SST) engine concepts of equivalent thrust and technological sophistication. It was found that in the case of an SST with a cruise speed of Mach 2.32, the through-flow fan engine may yield ranges 10 to 20% greater than the two alternatives considered. The engine has a conventional core, with the supersonic fan being driven by a concentric low-pressure turbine that is uncoupled with the single, high pressure turbine/compressor core spool. Among the topics discussed are the methods of analysis employed and perturbation studies concerning supersonic fan adiabatic efficiency, fan discharge characteristics and propulsion system weight.

  15. The Power Control Unit for the Propulsion Engine of GOCE Program

    NASA Astrophysics Data System (ADS)

    Tato, C.; Palencia, J.; de La Cruz, F.

    2004-10-01

    The IPCU Unit is in charge of controlling and monitoring the Ion Propulsion Assembly (Proportional Xenon Flow Assembly, PXFA, and Ion Thruster Assembly, ITA) belonging the DFACS (Drag Free Attitude Control System). Being the controlling function of the Propulsion Assembly, this unit involves the driving of the power supplies powering the propulsion system as well as the SW controlling all these power supplies (commanding the thruster and PXFA functions and monitoring the relevant parameters). The IPCU architecture involves two separate main functions. First one is the controlling electronics including an ERC32 uprocessor and a Mil-Bus I/F (based on the 1553B standard). The second main function, being the most challenging one, is the power controlling and supplying function. The PCU (Power Control Unit) faces three main aspects: eleven power converters driving the IPA (Ion Propulsion Assembly), the high voltage (1200 V) grounding concept used to refer five of these power converters, and an AC Bus distributing and powering all these power converters. This architecture concept makes the IPCU a challenging Unit in two main aspects: - the high voltage (1200 V) grounding system used for some important part of the electronics inside the Unit - the achieved performances. The IPCU measured functionality for the initial breadboard test bench showed the high performances expected. For a thrust range going from 0.6 to 20 mN the achieved performances in noise, linearity (0.9 %), and maximum error (< 2 %) with respect the commanded thrust meet the expectances for this innovative and demanding unit.

  16. Emissions from main propulsion engine on container ship at sea

    NASA Astrophysics Data System (ADS)

    Agrawal, Harshit; Welch, William A.; Henningsen, Svend; Miller, J. Wayne; Cocker, David R.

    2010-12-01

    Emission measurements were made for major gases and PM2.5 mass for a post PanaMax Class container vessel operating on heavy fuel oil at sea. Additional measurements were made for PM composition, elemental and organic carbon, select hydrocarbons, including PAHs, carbonyls, and n-alkanes. The testing followed the International Standard Organization protocols for emission measurements and operating test cycle. Results showed the weighted emission factor for NOx and PM2.5 were 19.77 ± 0.28 and 2.40 ± 0.05 g/kWh, respectively. The study provided a rare opportunity to repeat measurements made three years earlier on the same vessel. Emission factors of CO2 and NOx closely matched the earlier values, suggesting a low deterioration factor. Results showed the black carbon emission factor was 0.007 ± 0.001 g/kWh, an important metric for determining the radiative forcing contribution of marine engines.

  17. Advanced propulsion engine assessment based on a cermet reactor

    NASA Technical Reports Server (NTRS)

    Parsley, Randy C.

    1993-01-01

    A preferred Pratt & Whitney conceptual Nuclear Thermal Rocket Engine (NTRE) has been designed based on the fundamental NASA priorities of safety, reliability, cost, and performance. The basic philosophy underlying the design of the XNR2000 is the utilization of the most reliable form of ultrahigh temperature nuclear fuel and development of a core configuration which is optimized for uniform power distribution, operational flexibility, power maneuverability, weight, and robustness. The P&W NTRE system employs a fast spectrum, cermet fueled reactor configured in an expander cycle to ensure maximum operational safety. The cermet fuel form provides retention of fuel and fission products as well as high strength. A high level of confidence is provided by benchmark analysis and independent evaluations.

  18. Electrode erosion in steady-state electric propulsion engines

    NASA Technical Reports Server (NTRS)

    Pivirotto, Thomas J.; Deininger, William D.

    1988-01-01

    The anode and cathode of a 30 kW class arcjet engine were sectioned and analyzed. This arcjet was operated for a total time of 573 hr at power levels between 25 and 30 kW with ammonia at flow rates of 0.25 and 0.27 gm/s. The accumulated run time was sufficient to clearly establish erosion patterns and their causes. The type of electron emission from various parts of the cathode surface was made clear by scanning electron microscope analysis. A scanning electron microscope was used to study recrystallization on the hot anode surface. These electrodes were made of 2 percent thoriated tungsten and the surface thorium content and gradient perpendicular to the surfaces was determined by quantitative microprobe analysis. The results of this material analysis on the electrodes and recommendations for improving electrode operational life time are presented.

  19. Z-Pinch Magneto-Inertial Fusion Propulsion Engine Design Concept

    NASA Technical Reports Server (NTRS)

    Miernik, Janie H.; Statham, Geoffrey; Adams, Robert B.; Polsgrove, Tara; Fincher, Sharon; Fabisinski, Leo; Maples, C. Dauphne; Percy, Thomas K.; Cortez, Ross J.; Cassibry, Jason

    2011-01-01

    Fusion-based nuclear propulsion has the potential to enable fast interplanetary transportation. Due to the great distances between the planets of our solar system and the harmful radiation environment of interplanetary space, high specific impulse (Isp) propulsion in vehicles with high payload mass fractions must be developed to provide practical and safe vehicles for human spaceflight missions. Magneto-Inertial Fusion (MIF) is an approach which has been shown to potentially lead to a low cost, small fusion reactor/engine assembly (1). The Z-Pinch dense plasma focus method is an MIF concept in which a column of gas is compressed to thermonuclear conditions by an estimated axial current of approximately 100 MA. Recent advancements in experiments and the theoretical understanding of this concept suggest favorable scaling of fusion power output yield as I(sup 4) (2). The magnetic field resulting from the large current compresses the plasma to fusion conditions, and this is repeated over short timescales (10(exp -6) sec). This plasma formation is widely used in the field of Nuclear Weapons Effects (NWE) testing in the defense industry, as well as in fusion energy research. There is a wealth of literature characterizing Z-Pinch physics and existing models (3-5). In order to be useful in engineering analysis, a simplified Z-Pinch fusion thermodynamic model was developed to determine the quantity of plasma, plasma temperature, rate of expansion, energy production, etc. to calculate the parameters that characterize a propulsion system. The amount of nuclear fuel per pulse, mixture ratio of the D-T and nozzle liner propellant, and assumptions about the efficiency of the engine, enabled the sizing of the propulsion system and resulted in an estimate of the thrust and Isp of a Z-Pinch fusion propulsion system for the concept vehicle. MIF requires a magnetic nozzle to contain and direct the nuclear pulses, as well as a robust structure and radiation shielding. The structure

  20. Viewgraph description of Penn State's Propulsion Engineering Research Center: Activity highlights and future plans

    NASA Technical Reports Server (NTRS)

    Merkle, Charles L.

    1991-01-01

    Viewgraphs are presented that describe the progress and status of Penn State's Propulsion Engineering Research Center. The Center was established in Jul. 1988 by a grant from NASA's University Space Engineering Research Centers Program. After two and one-half years of operation, some 16 faculty are participating, and the Center is supporting 39 graduate students plus 18 undergraduates. In reviewing the Center's status, long-term plans and goals are reviewed and then the present status of the Center and the highlights and accomplishments of the past year are summarized. An overview of plans for the upcoming year are presented.

  1. Energy efficient engine: Flight propulsion system, preliminary analysis and design update

    NASA Technical Reports Server (NTRS)

    Stearns, E. M.

    1982-01-01

    The preliminary design of General Electric's Energy Efficient Engine (E3) was reported in detail in 1980. Since then, the design has been refined and the components have been rig-tested. The changes which have occurred in the engine and a reassessment of the economic payoff are presented in this report. All goals for efficiency, environmental considerations, and economic payoff are being met. The E3 Flight Propulsion System has 14.9% lower sfc than a CF6-50C. It provides a 7.1% reduction in direct operating cost for a short haul domestic transport and 14.5% reduction for an international long distance transport.

  2. Structural integrity and durability for Space Shuttle main engine and future reusable space propulsion systems

    NASA Technical Reports Server (NTRS)

    Marsik, S. J.; Gawrylowicz, H. T.

    1986-01-01

    NASA is conducting a program which will establish a technology base for the orderly evolution of reusable space propulsion systems. As part of that program, NASA initiated a Structural Integrity and Durability effort for advanced high-pressure oxygen-hydrogen rocket engine technology. That effort focuses on the development of: (1) accurate analytical models to describe flow fields; aerothermodynamic loads; structural responses; and fatigue/fracture, from which life prediction codes can be evolved; and (2) advanced instrumentation with capabilities to verify the codes in an SSME-like environment as well as the potential for future use as diagnostic sensors for real-time condition monitoring of critical engine components.

  3. Iroquois Engine for the Avro Arrow in the Propulsion Systems Laboratory

    NASA Image and Video Library

    1957-08-21

    A researcher examines the Orenda Iroquois PS.13 turbojet in a Propulsion Systems Laboratory test chamber at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Iroquois was being developed to power the CF-105 Arrow fighter designed by the Avro Canada Company. Avro began design work on the Arrow jet fighter in 1952. The company’s Orenda branch suggested building a titanium-based PS.13 Iroquois engine after development problems arose with the British engines that Avro had originally intended to use. The 10-stage, 20,000-pound-thrust Iroquois would prove to be more powerful than any contemporary US or British turbojet. It was also significantly lighter and more fuel efficient. An Iroquois was sent to Cleveland in April 1957 so that Lewis researchers could study the engine’s basic performance for the air force in the Propulsion Systems Laboratory. The tests were run over a wide range of speeds and altitudes with variations in exhaust-nozzle area. Initial studies determined the Iroquois’s windmilling and ignition characteristics at high altitude. After operating for 64 minutes, the engine was reignited at altitudes up to the 63,000-foot limit of the facility. Various modifications were attempted to reduce the occurrence of stall but did not totally eradicate the problem. The Arrow jet fighter made its initial flight in March 1958 powered by a substitute engine. In February 1959, however, both the engine and the aircraft programs were cancelled. The world’s superpowers had quickly transitioned from bombers to ballistic missiles which rendered the Avro Arrow prematurely obsolete.

  4. Marine propulsion device with engine heat recovery system and streamlining hull closures

    SciTech Connect

    Haynes, H. W.

    1985-11-12

    A Marine Jet Propulsion System for use as an inboard engine for boats is herein described. An engine or motor means is attached in a driving relationship to a pump and thrust output apparatus. Heat generated by and rejected by the engine or motor is passed into the pump base for dissipation into the outputted jet thrust stream. Air and/or exhaust gas from the engine is ejected around the jet output stream to reduce against-the-hull turbulence and jet stream or thrust energy losses. Streamlining hull closures for the jet pump intake and output ports are provided to reduce system hull drag when not in use and to limit marine organism growth inside the pump.

  5. Analysis of airframe/engine interactions in integrated flight and propulsion control

    NASA Technical Reports Server (NTRS)

    Schierman, John D.; Schmidt, David K.

    1991-01-01

    An analysis framework for the assessment of dynamic cross-coupling between airframe and engine systems from the perspective of integrated flight/propulsion control is presented. This analysis involves to determining the significance of the interactions with respect to deterioration in stability robustness and performance, as well as critical frequency ranges where problems may occur due to these interactions. The analysis illustrated here investigates both the airframe's effects on the engine control loops and the engine's effects on the airframe control loops in two case studies. The second case study involves a multi-input/multi-output analysis of the airframe. Sensitivity studies are performed on critical interactions to examine the degradations in the system's stability robustness and performance. Magnitudes of the interactions required to cause instabilities, as well as the frequencies at which the instabilities occur are recorded. Finally, the analysis framework is expanded to include control laws which contain cross-feeds between the airframe and engine systems.

  6. Graphene engineering by neon ion beams

    SciTech Connect

    Iberi, Vighter; Ievlev, Anton V.; Vlassiouk, Ivan; Jesse, Stephen; Kalinin, Sergei V.; Joy, David C.; Rondinone, Adam J.; Belianinov, Alex; Ovchinnikova, Olga S.

    2016-02-18

    Achieving the ultimate limits of materials and device performance necessitates the engineering of matter with atomic, molecular, and mesoscale fidelity. While common for organic and macromolecular chemistry, these capabilities are virtually absent for 2D materials. In contrast to the undesired effect of ion implantation from focused ion beam (FIB) lithography with gallium ions, and proximity effects in standard e-beam lithography techniques, the shorter mean free path and interaction volumes of helium and neon ions offer a new route for clean, resist free nanofabrication. Furthermore, with the advent of scanning helium ion microscopy, maskless He+ and Ne+ beam lithography of graphene based nanoelectronics is coming to the forefront. Here, we will discuss the use of energetic Ne ions in engineering graphene devices and explore the mechanical, electromechanical and chemical properties of the ion-milled devices using scanning probe microscopy (SPM). By using SPM-based techniques such as band excitation (BE) force modulation microscopy, Kelvin probe force microscopy (KPFM) and Raman spectroscopy, we demonstrate that the mechanical, electrical and optical properties of the exact same devices can be quantitatively extracted. Additionally, the effect of defects inherent in ion beam direct-write lithography, on the overall performance of the fabricated devices is elucidated.

  7. Graphene engineering by neon ion beams

    DOE PAGES

    Iberi, Vighter; Ievlev, Anton V.; Vlassiouk, Ivan; ...

    2016-02-18

    Achieving the ultimate limits of materials and device performance necessitates the engineering of matter with atomic, molecular, and mesoscale fidelity. While common for organic and macromolecular chemistry, these capabilities are virtually absent for 2D materials. In contrast to the undesired effect of ion implantation from focused ion beam (FIB) lithography with gallium ions, and proximity effects in standard e-beam lithography techniques, the shorter mean free path and interaction volumes of helium and neon ions offer a new route for clean, resist free nanofabrication. Furthermore, with the advent of scanning helium ion microscopy, maskless He+ and Ne+ beam lithography of graphenemore » based nanoelectronics is coming to the forefront. Here, we will discuss the use of energetic Ne ions in engineering graphene devices and explore the mechanical, electromechanical and chemical properties of the ion-milled devices using scanning probe microscopy (SPM). By using SPM-based techniques such as band excitation (BE) force modulation microscopy, Kelvin probe force microscopy (KPFM) and Raman spectroscopy, we demonstrate that the mechanical, electrical and optical properties of the exact same devices can be quantitatively extracted. Additionally, the effect of defects inherent in ion beam direct-write lithography, on the overall performance of the fabricated devices is elucidated.« less

  8. Flyer Acceleration by Pulsed Ion Beam Ablation and Application for Space Propulsion

    SciTech Connect

    Harada, Nobuhiro; Buttapeng, Chainarong; Yazawa, Masaru; Kashine, Kenji; Jiang Weihua; Yatsui, Kiyoshi

    2004-02-04

    Flyer acceleration by ablation plasma pressure produced by irradiation of intense pulsed ion beam has been studied. Acceleration process including expansion of ablation plasma was simulated based on fluid model. And interaction between incident pulsed ion beam and a flyer target was considered as accounting stopping power of it. In experiments, we used ETIGO-II intense pulsed ion beam generator with two kinds of diodes; 1) Magnetically Insulated Diode (MID, power densities of <100 J/cm2) and 2) Spherical-focused Plasma Focus Diode (SPFD, power densities of up to 4.3 kJ/cm2). Numerical results of accelerated flyer velocity agreed well with measured one over wide range of incident ion beam energy density. Flyer velocity of 5.6 km/s and ablation plasma pressure of 15 GPa was demonstrated by the present experiments. Acceleration of double-layer target consists of gold/aluminum was studied. For adequate layer thickness, such a flyer target could be much more accelerated than a single layer. Effect of waveform of ion beam was also examined. Parabolic waveform could accelerate more efficiently than rectangular waveform. Applicability of ablation propulsion was discussed. Specific impulse of 7000{approx}8000 seconds and time averaged thrust of up to 5000{approx}6000N can be expected. Their values can be controllable by changing power density of incident ion beam and pulse duration.

  9. Drag and Propulsive Characteristics of Air-Cooled Engine-Nacelle Installations for Large Airplane

    NASA Technical Reports Server (NTRS)

    Silverstein, Abe; Wilson, Herbert A , Jr

    1942-01-01

    An investigation was conducted in the NACA full-scale wind tunnel to determine the drag and the propulsive efficiency of nacelle-propeller arrangements for a large range of nacelle sizes. In contrast with usual tests with a single nacelle, these tests were conducted with nacelle-propeller installations on a large model of a four-engine airplane. Data are presented on the first part of the investigation, covering seven nacelle arrangements with nacelle diameters from 0.53 to 1.5 times the wing thickness. These ratios are similar to those occurring on airplanes weighing from about 20 to 100 tons. The results show the drag, the propulsive efficiency, and the over-all efficiency of the various nacelle arrangements as functions of the nacelle size, the propeller position, and the airplane lift coefficient. The effect of the nacelles on the aerodynamic characteristics of the model is shown for both propeller-removed and propeller-operating conditions.

  10. Upper-stage space shuttle propulsion by means of separate scramjet and rocket engines

    NASA Technical Reports Server (NTRS)

    Franciscus, L. C.; Allen, J. L.

    1972-01-01

    A preliminary mission study of a reusable vehicle from staging to orbit indicates payload advantages for a dual-propulsion system consisting of separate scramjet and rocket engines. In the analysis the scramjet operated continuously and the initiation of rocket operation was varied. For a stage weight of 500,000 lb the payload was 10.4 percent of stage weight or 70 percent greater than that of a comparable all-rocket-powered stage. When compared with a reusable two-state rocket vehicle having 50,000 lb payload, the use of the dual propulsion system for the second stage resulted in significant decreases in lift-off weight and empty weight, indicating possible lower hardware costs.

  11. Discussion on Performance History and Operations of Hayabusa Ion Engines

    NASA Astrophysics Data System (ADS)

    Nishiyama, Kazutaka; Kuninaka, Hitoshi

    The μ10 cathode-less electron cyclotron resonance ion engines, have propelled the Hayabusa asteroid explorer for seven years since its launch in May 2003. The spacecraft was focused on demonstrating the technology needed for a sample return from an asteroid, using electric propulsion, optical navigation, material sampling in a zero gravity field, and direct re-entry from a heliocentric orbit. The final stage of the return cruise and the subsequent trajectory correction maneuvers have been accomplished by using a special combined operation of neutralizer A and ion source B after the exhaustion of the other neutralizers' lives by the autumn of 2009. The total duration of the powered spaceflight was 25,590 h, which provided a delta-V of approximately 2.2 km/s and a total impulse of 1 MN·s. The degradation trends of the thruster performances have been investigated. It seems that the main cause of the degradation was the decrease in effective microwave power input to the discharge plasma induced by the increase in the transmission loss of the microwave feed system, and not due to the increase in the gas leakage through the accelerator grid apertures enlarged by erosion. Unintentional engine stop events have been summarized and analyzed. Most of them occurred due to the limit check errors of the backward microwave powers. Such errors can be decreased by carefully monitoring the trend change in microwave backward power as a function of xenon flow rate in future missions.

  12. The QED engine system: Direct-electric fusion-powered rocket propulsion systems

    NASA Astrophysics Data System (ADS)

    Bussard, Robert W.

    1993-01-01

    Practical ground-to-orbit and inter-orbital space flights both require propulsion systems of large flight-path-averaged specific impulse (Isp) and engine system thrust-to-mass-ratio (F/me=[F]) for useful payload and structure fractions in single-stage vehicles (Hunter 1966). Current rocket and air-breathing engine technologies lead to enormous vehicles and small payloads; a natural result of the limited specific energy available from chemical reactions. While nuclear energy far exceeds these specific energy limits (Bussard and DeLauer 1958), the inherent high-Isp advantages of fission propulsion concepts for space and air-breathing flight (Bussard and DeLauer 1965) are negated for manned systems by the massive radiation shielding required by their high radiation output (Bussard 1971). However, there are well-known radiation-free nuclear fusion reactions (Gross 1984) between isotopes of selected light elements (such as H+11B, D+3He) that yield only energetic charged particles, whose energy can be converted directly into electricity by confining electric fields (Moir and Barr 1973,1983). New confinement concepts using magnetic-electric-potentials (Bussard 1989a) or inertial-collisional-compression (ICC) (Bussard 1990) have been found that offer the prospect of clean, compact fusion systems with very high output and low mass. Their radiation-free d.c. electrical output can power unique new electron-beam-driven thrust systems of extremely high performance. Parametric design studies show that such charged-particle electric-discharge engines (``QED'' engines) might yield rocket propulsion systems with performance in the ranges of 2<[F]<6 and 1500

  13. In-Flight Operation of the Dawn Ion Propulsion System Through Start of the Vesta Cruise Phase

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.; Brophy, John R.

    2009-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218 kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) which will provide most of the delta V needed for heliocentric transfer to Vesta, orbit capture at Vesta, transfer to Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer to Ceres science orbits. The Dawn ion design is based on the design validated on NASA's Deep Space 1 (DS1) mission. However, because of the very substantial (11 km/s) delta V requirements for this mission Dawn requires two engines to complete its mission objectives. The power processor units (PPU), digital control and interface units (DCIU) slice boards and the xenon control assembly (XCA) are derivatives of the components used on DS1. The DCIUs and thrust gimbal assemblies (TGA) were developed at the Jet Propulsion Laboratory. The spacecraft was provided by Orbital Sciences Corporation, Sterling, Virginia, and the mission is managed by and operated from the Jet Propulsion Laboratory. Dawn partnered with Germany, Italy and Los Alamos National Laboratory for the science instruments. The mission is led by the principal investigator, Dr. Christopher Russell, from the University of California, Los Angeles. The first 80 days after launch were dedicated to the initial checkout of the spacecraft followed by cruise to Mars. Cruise thrusting leading to a Mars gravity assist began on December 17, 2007 and was successfully concluded as planned on October 31, 2008. During this time period the Dawn IPS was operated mostly at full power for approximately 6500 hours, consumed 71.7 kg of xenon and delivered approximately 1.8 km

  14. In-Flight Operation of the Dawn Ion Propulsion System Through Start of the Vesta Cruise Phase

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.; Brophy, John R.

    2009-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218 kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) which will provide most of the delta V needed for heliocentric transfer to Vesta, orbit capture at Vesta, transfer to Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer to Ceres science orbits. The Dawn ion design is based on the design validated on NASA's Deep Space 1 (DS1) mission. However, because of the very substantial (11 km/s) delta V requirements for this mission Dawn requires two engines to complete its mission objectives. The power processor units (PPU), digital control and interface units (DCIU) slice boards and the xenon control assembly (XCA) are derivatives of the components used on DS1. The DCIUs and thrust gimbal assemblies (TGA) were developed at the Jet Propulsion Laboratory. The spacecraft was provided by Orbital Sciences Corporation, Sterling, Virginia, and the mission is managed by and operated from the Jet Propulsion Laboratory. Dawn partnered with Germany, Italy and Los Alamos National Laboratory for the science instruments. The mission is led by the principal investigator, Dr. Christopher Russell, from the University of California, Los Angeles. The first 80 days after launch were dedicated to the initial checkout of the spacecraft followed by cruise to Mars. Cruise thrusting leading to a Mars gravity assist began on December 17, 2007 and was successfully concluded as planned on October 31, 2008. During this time period the Dawn IPS was operated mostly at full power for approximately 6500 hours, consumed 71.7 kg of xenon and delivered approximately 1.8 km

  15. Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid-Wing-Body Aircraft

    NASA Technical Reports Server (NTRS)

    Felder, James L.; Kim, Hyun Dae; Brown, Gerald V.

    2009-01-01

    Meeting NASA's N+3 goals requires a fundamental shift in approach to aircraft and engine design. Material and design improvements allow higher pressure and higher temperature core engines which improve the thermal efficiency. Propulsive efficiency, the other half of the overall efficiency equation, however, is largely determined by the fan pressure ratio (FPR). Lower FPR increases propulsive efficiency, but also dramatically reduces fan shaft speed through the combination of larger diameter fans and reduced fan tip speed limits. The result is that below an FPR of 1.5 the maximum fan shaft speed makes direct drive turbines problematic. However, it is the low pressure ratio fans that allow the improvement in propulsive efficiency which, along with improvements in thermal efficiency in the core, contributes strongly to meeting the N+3 goals for fuel burn reduction. The lower fan exhaust velocities resulting from lower FPRs are also key to meeting the aircraft noise goals. Adding a gear box to the standard turbofan engine allows acceptable turbine speeds to be maintained. However, development of a 50,000+ hp gearbox required by fans in a large twin engine transport aircraft presents an extreme technical challenge, therefore another approach is needed. This paper presents a propulsion system which transmits power from the turbine to the fan electrically rather than mechanically. Recent and anticipated advances in high temperature superconducting generators, motors, and power lines offer the possibility that such devices can be used to transmit turbine power in aircraft without an excessive weight penalty. Moving to such a power transmission system does more than provide better matching between fan and turbine shaft speeds. The relative ease with which electrical power can be distributed throughout the aircraft opens up numerous other possibilities for new aircraft and propulsion configurations and modes of operation. This paper discusses a number of these new

  16. Autonomous Propulsion System Technology Being Developed to Optimize Engine Performance Throughout the Lifecycle

    NASA Technical Reports Server (NTRS)

    Litt, Jonathan S.

    2004-01-01

    The goal of the Autonomous Propulsion System Technology (APST) project is to reduce pilot workload under both normal and anomalous conditions. Ongoing work under APST develops and leverages technologies that provide autonomous engine monitoring, diagnosing, and controller adaptation functions, resulting in an integrated suite of algorithms that maintain the propulsion system's performance and safety throughout its life. Engine-to-engine performance variation occurs among new engines because of manufacturing tolerances and assembly practices. As an engine wears, the performance changes as operability limits are reached. In addition to these normal phenomena, other unanticipated events such as sensor failures, bird ingestion, or component faults may occur, affecting pilot workload as well as compromising safety. APST will adapt the controller as necessary to achieve optimal performance for a normal aging engine, and the safety net of APST algorithms will examine and interpret data from a variety of onboard sources to detect, isolate, and if possible, accommodate faults. Situations that cannot be accommodated within the faulted engine itself will be referred to a higher level vehicle management system. This system will have the authority to redistribute the faulted engine's functionality among other engines, or to replan the mission based on this new engine health information. Work is currently underway in the areas of adaptive control to compensate for engine degradation due to aging, data fusion for diagnostics and prognostics of specific sensor and component faults, and foreign object ingestion detection. In addition, a framework is being defined for integrating all the components of APST into a unified system. A multivariable, adaptive, multimode control algorithm has been developed that accommodates degradation-induced thrust disturbances during throttle transients. The baseline controller of the engine model currently being investigated has multiple control

  17. Autonomous Propulsion System Technology Being Developed to Optimize Engine Performance Throughout the Lifecycle

    NASA Technical Reports Server (NTRS)

    Litt, Jonathan S.

    2004-01-01

    The goal of the Autonomous Propulsion System Technology (APST) project is to reduce pilot workload under both normal and anomalous conditions. Ongoing work under APST develops and leverages technologies that provide autonomous engine monitoring, diagnosing, and controller adaptation functions, resulting in an integrated suite of algorithms that maintain the propulsion system's performance and safety throughout its life. Engine-to-engine performance variation occurs among new engines because of manufacturing tolerances and assembly practices. As an engine wears, the performance changes as operability limits are reached. In addition to these normal phenomena, other unanticipated events such as sensor failures, bird ingestion, or component faults may occur, affecting pilot workload as well as compromising safety. APST will adapt the controller as necessary to achieve optimal performance for a normal aging engine, and the safety net of APST algorithms will examine and interpret data from a variety of onboard sources to detect, isolate, and if possible, accommodate faults. Situations that cannot be accommodated within the faulted engine itself will be referred to a higher level vehicle management system. This system will have the authority to redistribute the faulted engine's functionality among other engines, or to replan the mission based on this new engine health information. Work is currently underway in the areas of adaptive control to compensate for engine degradation due to aging, data fusion for diagnostics and prognostics of specific sensor and component faults, and foreign object ingestion detection. In addition, a framework is being defined for integrating all the components of APST into a unified system. A multivariable, adaptive, multimode control algorithm has been developed that accommodates degradation-induced thrust disturbances during throttle transients. The baseline controller of the engine model currently being investigated has multiple control

  18. Convert Ten Foot Environmental Test Chamber into an Ion Engine Test Chamber

    NASA Technical Reports Server (NTRS)

    VanVelzer, Paul

    2006-01-01

    The 10 Foot Space Simulator at the Jet Propulsion Laboratory has been used for the last 40 years to test numerous spacecraft, including the Ranger series, several Mariner class, among many others and finally, the Spirit and Opportunity Mars Rovers. The request was made to convert this facility to an Ion Engine test facility, with a possible long term life test. The Ion engine was to propel the Prometheus spacecraft to Jupiter's moons. This paper discusses the challenges that were met, both from a procedural and physical standpoint. The converted facility must operate unattended, support a 30 Kw Ion Engine, operate economically, and be easily converted back to former operation as a spacecraft test facility.

  19. Convert Ten Foot Environmental Test Chamber into an Ion Engine Test Chamber

    NASA Technical Reports Server (NTRS)

    VanVelzer, Paul

    2006-01-01

    The 10 Foot Space Simulator at the Jet Propulsion Laboratory has been used for the last 40 years to test numerous spacecraft, including the Ranger series, several Mariner class, among many others and finally, the Spirit and Opportunity Mars Rovers. The request was made to convert this facility to an Ion Engine test facility, with a possible long term life test. The Ion engine was to propel the Prometheus spacecraft to Jupiter's moons. This paper discusses the challenges that were met, both from a procedural and physical standpoint. The converted facility must operate unattended, support a 30 Kw Ion Engine, operate economically, and be easily converted back to former operation as a spacecraft test facility.

  20. General Electric I-40 Engine at the Lewis Flight Propulsion Laboratory

    NASA Image and Video Library

    1946-08-21

    A mechanic works on a General Electric I-40 turbojet at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The military selected General Electric’s West Lynn facility in 1941 to secretly replicate the centrifugal turbojet engine designed by British engineer Frank Whittle. General Electric’s first attempt, the I-A, was fraught with problems. The design was improved somewhat with the subsequent I-16 engine. It was not until the engine's next reincarnation as the I-40 in 1943 that General Electric’s efforts paid off. The 4000-pound thrust I-40 was incorporated into the Lockheed Shooting Star airframe and successfully flown in June 1944. The Shooting Star became the US’s first successful jet aircraft and the first US aircraft to reach 500 miles per hour. The NACA’s Lewis Flight Propulsion Laboratory studied all of General Electric’s centrifugal turbojets both during World War II and afterwards. The entire Shooting Star aircraft was investigated in the Altitude Wind Tunnel during 1945. The researchers studied the engine compressor performance, thrust augmentation using a water injection, and compared different fuel blends in a single combustor. The mechanic in this photograph is inserting a combustion liner into one of the 14 combustor cans. The compressor, which is not yet installed in this photograph, pushed high pressure air into these combustors. There the air mixed with the fuel and was heated. The hot air was then forced through a rotating turbine that powered the engine before being expelled out the nozzle to produce thrust.

  1. Quiet Clean Short-haul Experimental Engine (QCSEE) preliminary under the wing flight propulsion system analysis report

    NASA Technical Reports Server (NTRS)

    Howard, D. F.

    1976-01-01

    The preliminary design and installation of high bypass, geared turbofan engine with a composite nacelle forming the propulsion system for a short haul passenger aircraft are described. The technology required for externally blown flap aircraft with under the wing (UTW) propulsion system installations for introduction into passenger service in the mid 1980's is included. The design, fabrication, and testing of this UTW experimental engine containing the required technology items for low noise, fuel economy, with composite structure for reduced weight and digital engine control are provided.

  2. Integrated Design and Engineering Analysis (IDEA) Environment - Propulsion Related Module Development and Vehicle Integration

    NASA Technical Reports Server (NTRS)

    Kamhawi, Hilmi N.

    2013-01-01

    This report documents the work performed during the period from May 2011 - October 2012 on the Integrated Design and Engineering Analysis (IDEA) environment. IDEA is a collaborative environment based on an object-oriented, multidisciplinary, distributed framework using the Adaptive Modeling Language (AML). This report will focus on describing the work done in the areas of: (1) Integrating propulsion data (turbines, rockets, and scramjets) in the system, and using the data to perform trajectory analysis; (2) Developing a parametric packaging strategy for a hypersonic air breathing vehicles allowing for tank resizing when multiple fuels and/or oxidizer are part of the configuration; and (3) Vehicle scaling and closure strategies.

  3. Workstation technology for engineering mission operations at the Jet Propulsion Laboratory

    NASA Astrophysics Data System (ADS)

    Miller, Kevin J.; Murphy, Susan C.

    1990-10-01

    The Operations Engineering Laboratory (OEL) at the Jet Propulsion Laboratory has been developing graphics tools to automate document preparation in support of space flight mission operations. One such tool, which generates a daily Space Flight Operations Schedule (SFOS), a timeline display of the schedule of spacecraft activities for the Voyager mission is described. The tool consists of two parts: a series of programs that preprocess various command files and a graphics editor. The code of the graphics editor was developed with reusability as a primary objective and has since served as the basis for the generation of other automation tools.

  4. Use of Soft Computing Technologies for a Qualitative and Reliable Engine Control System for Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Trevino, Luis; Brown, Terry; Crumbley, R. T. (Technical Monitor)

    2001-01-01

    The problem to be addressed in this paper is to explore how the use of Soft Computing Technologies (SCT) could be employed to improve overall vehicle system safety, reliability, and rocket engine performance by development of a qualitative and reliable engine control system (QRECS). Specifically, this will be addressed by enhancing rocket engine control using SCT, innovative data mining tools, and sound software engineering practices used in Marshall's Flight Software Group (FSG). The principle goals for addressing the issue of quality are to improve software management, software development time, software maintenance, processor execution, fault tolerance and mitigation, and nonlinear control in power level transitions. The intent is not to discuss any shortcomings of existing engine control methodologies, but to provide alternative design choices for control, implementation, performance, and sustaining engineering, all relative to addressing the issue of reliability. The approaches outlined in this paper will require knowledge in the fields of rocket engine propulsion (system level), software engineering for embedded flight software systems, and soft computing technologies (i.e., neural networks, fuzzy logic, data mining, and Bayesian belief networks); some of which are briefed in this paper. For this effort, the targeted demonstration rocket engine testbed is the MC-1 engine (formerly FASTRAC) which is simulated with hardware and software in the Marshall Avionics & Software Testbed (MAST) laboratory that currently resides at NASA's Marshall Space Flight Center, building 4476, and is managed by the Avionics Department. A brief plan of action for design, development, implementation, and testing a Phase One effort for QRECS is given, along with expected results. Phase One will focus on development of a Smart Start Engine Module and a Mainstage Engine Module for proper engine start and mainstage engine operations. The overall intent is to demonstrate that by

  5. Use of Soft Computing Technologies for a Qualitative and Reliable Engine Control System for Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Trevino, Luis; Brown, Terry; Crumbley, R. T. (Technical Monitor)

    2001-01-01

    The problem to be addressed in this paper is to explore how the use of Soft Computing Technologies (SCT) could be employed to improve overall vehicle system safety, reliability, and rocket engine performance by development of a qualitative and reliable engine control system (QRECS). Specifically, this will be addressed by enhancing rocket engine control using SCT, innovative data mining tools, and sound software engineering practices used in Marshall's Flight Software Group (FSG). The principle goals for addressing the issue of quality are to improve software management, software development time, software maintenance, processor execution, fault tolerance and mitigation, and nonlinear control in power level transitions. The intent is not to discuss any shortcomings of existing engine control methodologies, but to provide alternative design choices for control, implementation, performance, and sustaining engineering, all relative to addressing the issue of reliability. The approaches outlined in this paper will require knowledge in the fields of rocket engine propulsion (system level), software engineering for embedded flight software systems, and soft computing technologies (i.e., neural networks, fuzzy logic, data mining, and Bayesian belief networks); some of which are briefed in this paper. For this effort, the targeted demonstration rocket engine testbed is the MC-1 engine (formerly FASTRAC) which is simulated with hardware and software in the Marshall Avionics & Software Testbed (MAST) laboratory that currently resides at NASA's Marshall Space Flight Center, building 4476, and is managed by the Avionics Department. A brief plan of action for design, development, implementation, and testing a Phase One effort for QRECS is given, along with expected results. Phase One will focus on development of a Smart Start Engine Module and a Mainstage Engine Module for proper engine start and mainstage engine operations. The overall intent is to demonstrate that by

  6. Quiet Clean Short-haul Experimental Engine (QCSEE) preliminary over-the-wing flight propulsion system analysis report

    NASA Technical Reports Server (NTRS)

    Howard, D. F.

    1977-01-01

    The preliminary design of the over-the-wing flight propulsion system installation and nacelle component and systems design features of a short-haul, powered lift aircraft are presented. Economic studies are also presented and show that high bypass, low pressure ratio turbofan engines have the potential of providing an economical propulsion system for achieving the very quiet aircraft noise level of 95 EPNdB on a 152.4 m sideline.

  7. Development of a miniature microwave electron cyclotron resonance plasma ion thruster for exospheric micro-propulsion

    SciTech Connect

    Dey, Indranuj; Toyoda, Yuji; Yamamoto, Naoji; Nakashima, Hideki

    2015-12-15

    A miniature microwave electron cyclotron resonance plasma source [(discharge diameter)/(microwave cutoff diameter) < 0.3] has been developed at Kyushu University to be used as an ion thruster in micro-propulsion applications in the exosphere. The discharge source uses both radial and axial magnetostatic field confinement to facilitate electron cyclotron resonance and increase the electron dwell time in the volume, thereby enhancing plasma production efficiency. Performance of the ion thruster is studied at 3 microwave frequencies (1.2 GHz, 1.6 GHz, and 2.45 GHz), for low input powers (<15 W) and small xenon mass flow rates (<40 μg/s), by experimentally measuring the extracted ion beam current through a potential difference of ≅1200 V. The discharge geometry is found to operate most efficiently at an input microwave frequency of 1.6 GHz. At this frequency, for an input power of 8 W, and propellant (xenon) mass flow rate of 21 μg/s, 13.7 mA of ion beam current is obtained, equivalent to an calculated thrust of 0.74 mN.

  8. Development of a miniature microwave electron cyclotron resonance plasma ion thruster for exospheric micro-propulsion.

    PubMed

    Dey, Indranuj; Toyoda, Yuji; Yamamoto, Naoji; Nakashima, Hideki

    2015-12-01

    A miniature microwave electron cyclotron resonance plasma source [(discharge diameter)/(microwave cutoff diameter) < 0.3] has been developed at Kyushu University to be used as an ion thruster in micro-propulsion applications in the exosphere. The discharge source uses both radial and axial magnetostatic field confinement to facilitate electron cyclotron resonance and increase the electron dwell time in the volume, thereby enhancing plasma production efficiency. Performance of the ion thruster is studied at 3 microwave frequencies (1.2 GHz, 1.6 GHz, and 2.45 GHz), for low input powers (<15 W) and small xenon mass flow rates (<40 μg/s), by experimentally measuring the extracted ion beam current through a potential difference of ≅1200 V. The discharge geometry is found to operate most efficiently at an input microwave frequency of 1.6 GHz. At this frequency, for an input power of 8 W, and propellant (xenon) mass flow rate of 21 μg/s, 13.7 mA of ion beam current is obtained, equivalent to an calculated thrust of 0.74 mN.

  9. Development of a miniature microwave electron cyclotron resonance plasma ion thruster for exospheric micro-propulsion

    NASA Astrophysics Data System (ADS)

    Dey, Indranuj; Toyoda, Yuji; Yamamoto, Naoji; Nakashima, Hideki

    2015-12-01

    A miniature microwave electron cyclotron resonance plasma source [(discharge diameter)/(microwave cutoff diameter) < 0.3] has been developed at Kyushu University to be used as an ion thruster in micro-propulsion applications in the exosphere. The discharge source uses both radial and axial magnetostatic field confinement to facilitate electron cyclotron resonance and increase the electron dwell time in the volume, thereby enhancing plasma production efficiency. Performance of the ion thruster is studied at 3 microwave frequencies (1.2 GHz, 1.6 GHz, and 2.45 GHz), for low input powers (<15 W) and small xenon mass flow rates (<40 μg/s), by experimentally measuring the extracted ion beam current through a potential difference of ≅1200 V. The discharge geometry is found to operate most efficiently at an input microwave frequency of 1.6 GHz. At this frequency, for an input power of 8 W, and propellant (xenon) mass flow rate of 21 μg/s, 13.7 mA of ion beam current is obtained, equivalent to an calculated thrust of 0.74 mN.

  10. High Altitude Small Engine Test Techniques at the NASA Glenn Propulsion Systems Lab

    NASA Technical Reports Server (NTRS)

    Castner, Raymond; Wyzykowski, John; Chiapetta, Santo

    2001-01-01

    A High Altitude Test was performed in the Propulsion Systems Lab (PSL) at the NASA Glenn Research Center using a Pratt and Whitney Canada PW545 jet engine. This engine was tested to develop a highaltitude database on small, high-bypass ratio, engine performance and operability. Industry is interested in the use of high-bypass engines for Uninhabited Aerial Vehicles (UAV's) to perform high altitude surveillance. The tests were a combined effort between Pratt & Whitney Canada (PWC) and NASA Glenn Research Center. A large portion of this test activity was to collect performance data with a highly instrumented low-pressure turbine. Low-pressure turbine aerodynamic performance at low Reynolds numbers was collected and compared to analytical models developed by NASA and PWC. This report describes the test techniques implemented to obtain high accuracy turbine performance data in an altitude test facility, including high accuracy airflow at high altitudes, very low mass flow, and low air temperatures. Major accomplishments from this test activity were to collect accurate and repeatable turbine performance data at high altitudes to within 1 percent. Data were collected at 19,800m, 16,750m, and 13,700m providing documentation of diminishing LPT performance with reductions in Reynolds number in an actual engine flight environment. The test provided a unique database for the development of engine analysis codes to be used for future LPT performance improvements.

  11. A One Dimensional, Time Dependent Inlet/Engine Numerical Simulation for Aircraft Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Garrard, Doug; Davis, Milt, Jr.; Cole, Gary

    1999-01-01

    The NASA Lewis Research Center (LeRC) and the Arnold Engineering Development Center (AEDC) have developed a closely coupled computer simulation system that provides a one dimensional, high frequency inlet/engine numerical simulation for aircraft propulsion systems. The simulation system, operating under the LeRC-developed Application Portable Parallel Library (APPL), closely coupled a supersonic inlet with a gas turbine engine. The supersonic inlet was modeled using the Large Perturbation Inlet (LAPIN) computer code, and the gas turbine engine was modeled using the Aerodynamic Turbine Engine Code (ATEC). Both LAPIN and ATEC provide a one dimensional, compressible, time dependent flow solution by solving the one dimensional Euler equations for the conservation of mass, momentum, and energy. Source terms are used to model features such as bleed flows, turbomachinery component characteristics, and inlet subsonic spillage while unstarted. High frequency events, such as compressor surge and inlet unstart, can be simulated with a high degree of fidelity. The simulation system was exercised using a supersonic inlet with sixty percent of the supersonic area contraction occurring internally, and a GE J85-13 turbojet engine.

  12. Ion Propulsion Thruster Including a Plurality of Ion Optic Electrode Pairs

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J. (Inventor)

    2016-01-01

    Ion optics for use in a conventional or annular or other shaped ion thruster are disclosed including a plurality of planar, spaced apart ion optic electrode pairs sized to include a diameter smaller than the diameter of thruster exhaust and retained in, on or otherwise associated with a frame across the thruster exhaust. An electrical connection may be provided for establishing electrical connectivity among a set of first upstream electrodes and an electrical connection may be provided for establishing electrical connectivity among the second downstream electrodes.

  13. Planetary Science Enabled by High Power Ion Propulsion Systems from NASA's Prometheus Program

    NASA Astrophysics Data System (ADS)

    Cooper, John

    2004-11-01

    NASA's Prometheus program seeks to develop new generations of spacecraft nuclear-power and ion propulsion systems for applications to future planetary missions. The Science Definition Team for the first mission in the Prometheus series, the Jupiter Icy Moons Orbiter (JIMO), has defined science objectives for in-situ orbital exploration of the icy Galilean moons (Europa, Ganymede, Callisto) and the Jovian magnetosphere along with remote observations of Jupiter's atmosphere and aurorae, the volcanic moon Io, and other elements of the Jovian system. Important to this forum is that JIMO power and propulsion systems will need to be designed to minimize magnetic, radio, neutral gas, and plasma backgrounds that might otherwise interfere with achievement of mission science objectives. Another potential Prometheus mission of high science interest would be an extended tour of primitive bodies in the solar system, including asteroids, Jupiter family comets, Centaurs, and Kuiper Belt Objects (KBO). The final landed phase of this mission might include an active keplerian experiment for detectable (via downlink radio doppler shift) acceleration of a small kilometer-size Centaur or KBO object, likely the satellite of a larger object observable from Earth. This would have obvious application to testing of mitigation techniques for Earth impact hazards.

  14. Dynamic Temperature and Pressure Measurements in the Core of a Propulsion Engine

    NASA Technical Reports Server (NTRS)

    Schuster, Bill; Gordon, Grant; Hultgren, Lennart S.

    2015-01-01

    Dynamic temperature and pressure measurements were made in the core of a TECH977 propulsion engine as part of a NASA funded investigation into indirect combustion noise. Dynamic temperature measurements were made in the combustor, the inter-turbine duct, and the mixer using ten two-wire thermocouple probes. Internal dynamic pressure measurements were made at the same locations using piezoresistive transducers installed in semi-infinite coils. Measurements were acquired at four steady state operating conditions covering the range of aircraft approach power settings. Fluctuating gas temperature spectra were computed from the thermocouple probe voltage measurements using a compensation procedure that was developed under previous NASA test programs. A database of simultaneously acquired dynamic temperature and dynamic pressure measurements was produced. Spectral and cross-spectral analyses were conducted to explore the characteristics of the temperature and pressure fluctuations inside the engine, with a particular focus on attempting to identify the presence of indirect combustion noise.

  15. Middle atmosphere NO/x/ production due to ion propulsion induced radiation belt proton precipitation

    NASA Technical Reports Server (NTRS)

    Aikin, A. C.; Jackman, C. H.

    1980-01-01

    The suggestion that keV Ar(+) resulting from ion propulsion operations during solar power satellite construction could cause energetic proton precipitation from the inner radiation belt is examined to determine if such precipitation could cause significant increases in middle atmosphere nitric oxide concentrations thereby adversely affecting stratospheric ozone. It is found that the initial production rate of NO (mole/cu cm-sec) at 50 km is 130 times that due to nitrous oxide reacting with excited oxygen. However, since the time required to empty the inner belt of protons is about 1 sec and short compared to the replenishment time due to neutron decay, precipitation of inner radiation belt protons will have no adverse atmospheric environmental effect.

  16. Ice Crystal Icing Engine Testing in the NASA Glenn Research Center's Propulsion Systems Laboratory (PSL): Altitude Investigation

    NASA Technical Reports Server (NTRS)

    Oliver, Michael J.

    2015-01-01

    The National Aeronautics and Space Administration conducted a full scale ice crystal icing turbofan engine test in the NASA Glenn Research Centers Propulsion Systems Laboratory (PSL) Facility in February 2013. Honeywell Engines supplied the test article, an obsolete, unmodified Lycoming ALF502-R5 turbofan engine serial number LF01 that experienced an un-commanded loss of thrust event while operating at certain high altitude ice crystal icing conditions. These known conditions were duplicated in the PSL for this testing.

  17. Wear Mechanisms in Electron Sources for Ion Propulsion, 2: Discharge Hollow Cathode

    NASA Technical Reports Server (NTRS)

    Mikellides, Ioannis G.; Katz, Ira; Goebel, Dan M.; Jameson, Kristina K.; Polk, James E.

    2008-01-01

    The wear of the keeper electrode in discharge hollow cathodes is a major impediment to the implementation of ion propulsion onboard long-duration space science missions. The development of a predictive theoretical model for hollow cathode keeper life has long been sought, but its realization has been hindered by the complexities associated with the physics of the partially ionized gas and the associated erosion mechanisms in these devices. Thus, although several wear mechanisms have been hypothesized, a quantitative explanation of life test erosion profiles has remained incomplete. A two-dimensional model of the partially ionized gas in a discharge cathode has been developed and applied to understand the mechanisms that drove the erosion of the keeper in two long-duration life tests of a 30-cm ion thruster. An extensive set of comparisons between predictions by the numerical simulations and measurements of the plasma properties and of the erosion patterns is presented. It is found that the near-plume plasma oscillations, predicted by theory and observed by experiment, effectively enhance the resistivity of the plasma as well as the energy of ions striking the keeper.

  18. Wear Mechanisms in Electron Sources for Ion Propulsion, 2: Discharge Hollow Cathode

    NASA Technical Reports Server (NTRS)

    Mikellides, Ioannis G.; Katz, Ira; Goebel, Dan M.; Jameson, Kristina K.; Polk, James E.

    2008-01-01

    The wear of the keeper electrode in discharge hollow cathodes is a major impediment to the implementation of ion propulsion onboard long-duration space science missions. The development of a predictive theoretical model for hollow cathode keeper life has long been sought, but its realization has been hindered by the complexities associated with the physics of the partially ionized gas and the associated erosion mechanisms in these devices. Thus, although several wear mechanisms have been hypothesized, a quantitative explanation of life test erosion profiles has remained incomplete. A two-dimensional model of the partially ionized gas in a discharge cathode has been developed and applied to understand the mechanisms that drove the erosion of the keeper in two long-duration life tests of a 30-cm ion thruster. An extensive set of comparisons between predictions by the numerical simulations and measurements of the plasma properties and of the erosion patterns is presented. It is found that the near-plume plasma oscillations, predicted by theory and observed by experiment, effectively enhance the resistivity of the plasma as well as the energy of ions striking the keeper.

  19. Starter circuit for an ion engine

    NASA Technical Reports Server (NTRS)

    Cardwell, Jr., Gilbert I. (Inventor); Phelps, Thomas K. (Inventor)

    2002-01-01

    A starter circuit particularly suitable for a plasma of an ion engine for a spacecraft includes a power supply having an output inductor with a tap. A switch is coupled to the tap. The switch has a control input. A pulse control logic circuit is coupled to said control input, said pulse control logic circuit controlling said switch to an off state to generate a high voltage discharge.

  20. Starter circuit for an ion engine

    NASA Technical Reports Server (NTRS)

    Cardwell, Jr., Gilbert I. (Inventor); Phelps, Thomas K. (Inventor)

    2001-01-01

    A starter circuit particularly suitable for a plasma of an ion engine for a spacecraft includes a power supply having an output inductor with a tap. A switch is coupled to the tap. The switch has a control input. A pulse control logic circuit is coupled to said control input, said pulse control logic circuit controlling said switch to an off state to generate a high voltage discharge.

  1. Graphene engineering by neon ion beams

    NASA Astrophysics Data System (ADS)

    Iberi, Vighter; Ievlev, Anton V.; Vlassiouk, Ivan; Jesse, Stephen; Kalinin, Sergei V.; Joy, David C.; Rondinone, Adam J.; Belianinov, Alex; Ovchinnikova, Olga S.

    2016-03-01

    Achieving the ultimate limits of lithographic resolution and material performance necessitates engineering of matter with atomic, molecular, and mesoscale fidelity. With the advent of scanning helium ion microscopy, maskless He+ and Ne+ beam lithography of 2D materials, such as graphene-based nanoelectronics, is coming to the forefront as a tool for fabrication and surface manipulation. However, the effects of using a Ne focused-ion-beam on the fidelity of structures created out of 2D materials have yet to be explored. Here, we will discuss the use of energetic Ne ions in engineering graphene nanostructures and explore their mechanical, electromechanical and chemical properties using scanning probe microscopy (SPM). By using SPM-based techniques such as band excitation (BE) force modulation microscopy, Kelvin probe force microscopy (KPFM) and Raman spectroscopy, we are able to ascertain changes in the mechanical, electrical and optical properties of Ne+ beam milled graphene nanostructures and surrounding regions. Additionally, we are able to link localized defects around the milled graphene to ion milling parameters such as dwell time and number of beam passes in order to characterize the induced changes in mechanical and electromechanical properties of the graphene surface.

  2. Advanced Space Propulsion

    NASA Technical Reports Server (NTRS)

    Frisbee, Robert H.

    1996-01-01

    system with a low initial development and infrastructure cost and a high operating cost. Note however that this has resulted in a 'Catch 22' standoff between the need for large initial investment that is amortized over many launches to reduce costs, and the limited number of launches possible at today's launch costs. Some examples of missions enabled (either in cost or capability) by advanced propulsion include long-life station-keeping or micro-spacecraft applications using electric propulsion or BMDO-derived micro-thrusters, low-cost orbit raising (LEO to GEO or Lunar orbit) using electric propulsion, robotic planetary missions using aerobraking or electric propulsion, piloted Mars missions using aerobraking and/or propellant production from Martian resources, very fast (100-day round-trip) piloted Mars missions using fission or fusion propulsion, and, finally, interstellar missions using fusion, antimatter, or beamed energy. The NASA Advanced Propulsion Technology program at the Jet Propulsion Laboratory (JPL) is aimed at assessing the feasibility of a range of near-term to far term advanced propulsion technologies that have the potential to reduce costs and/or enable future space activities. The program includes cooperative modeling and research activities between JPL and various universities and industry; and directly supported independent research at universities and industry. The cooperative program consists of mission studies, research and development of ion engine technology using C60 (Buckminsterfullerene) propellant, and research and development of lithium-propellant Lorentz-force accelerator (LFA) engine technology. The university/industry-supported research includes modeling and proof-of-concept experiments in advanced, high-lsp, long-life electric propulsion, and in fusion propulsion.

  3. High performance auxiliary-propulsion ion thruster with ion-machined accelerator grid

    NASA Technical Reports Server (NTRS)

    Hudson, W. R.; Banks, B. A.

    1975-01-01

    An improvement in thruster performance was achieved by reducing the diameter of the accelerator grid holes. The smaller accelerator grid holes resulted in a reduction in neutral mercury atoms escaping the discharge chamber, which in turn enhanced the discharge propellant utilization from approximately 68 percent to 92 percent. The accelerator grids were fabricated by ion machining with an 8-centimeter-diameter thruster, and the screen grid holes individually focused ion beamlets onto the blank accelerator grid. The resulting accelerator grid holes are less than 1.12 millimeters in diameter, while previously used accelerator grids had hole diameters of 1.69 millimeters. The thruster could be operated with the small-hole accelerator grid at neutralizer potential.

  4. High Speed Propulsion: Engine Design - Integration and Thermal Management (Propulsion a vitesse elevee : Conception du moteur - integration et gestion thermique)

    DTIC Science & Technology

    2010-09-01

    hours). Although hypersonic technology has significantly matured over the last 40 years technical challenges remain: intake design and optimization...Lecture Series is to provide clear guidelines regarding the design of the propulsion unit and integration into the airframe. First the intake physics... Hypersonic systems will provide a revolution in commercial transport, space access (lower power density), and future NATO missions (global reach in 2

  5. MITEE: A Compact Ultralight Nuclear Thermal Propulsion Engine for Planetary Science Missions

    NASA Astrophysics Data System (ADS)

    Powell, J.; Maise, G.; Paniagua, J.

    2001-01-01

    A new approach for a near-term compact, ultralight nuclear thermal propulsion engine, termed MITEE (Miniature Reactor Engine) is described. MITEE enables a wide range of new and unique planetary science missions that are not possible with chemical rockets. With U-235 nuclear fuel and hydrogen propellant the baseline MITEE engine achieves a specific impulse of approximately 1000 seconds, a thrust of 28,000 newtons, and a total mass of only 140 kilograms, including reactor, controls, and turbo-pump. Using higher performance nuclear fuels like U-233, engine mass can be reduced to as little as 80 kg. Using MITEE, V additions of 20 km/s for missions to outer planets are possible compared to only 10 km/s for H2/O2 engines. The much greater V with MITEE enables much faster trips to the outer planets, e.g., two years to Jupiter, three years to Saturn, and five years to Pluto, without needing multiple planetary gravity assists. Moreover, MITEE can utilize in-situ resources to further extend mission V. One example of a very attractive, unique mission enabled by MITEE is the exploration of a possible subsurface ocean on Europa and the return of samples to Earth. Using MITEE, a spacecraft would land on Europa after a two-year trip from Earth orbit and deploy a small nuclear heated probe that would melt down through its ice sheet. The probe would then convert to a submersible and travel through the ocean collecting samples. After a few months, the probe would melt its way back up to the MITEE lander, which would have replenished its hydrogen propellant by melting and electrolyzing Europa surface ice. The spacecraft would then return to Earth. Total mission time is only five years, starting from departure from Earth orbit. Other unique missions include Neptune and Pluto orbiter, and even a Pluto sample return. MITEE uses the cermet Tungsten-UO2 fuel developed in the 1960's for the 710 reactor program. The W-UO2 fuel has demonstrated capability to operate in 3000 K hydrogen for

  6. Aero-Propulsion Technology (APT) Task V Low Noise ADP Engine Definition Study

    NASA Technical Reports Server (NTRS)

    Holcombe, V.

    2003-01-01

    A study was conducted to identify and evaluate noise reduction technologies for advanced ducted prop propulsion systems that would allow increased capacity operation and result in an economically competitive commercial transport. The study investigated the aero/acoustic/structural advancements in fan and nacelle technology required to match or exceed the fuel burned and economic benefits of a constrained diameter large Advanced Ducted Propeller (ADP) compared to an unconstrained ADP propulsion system with a noise goal of 5 to 10 EPNDB reduction relative to FAR 36 Stage 3 at each of the three measuring stations namely, takeoff (cutback), approach and sideline. A second generation ADP was selected to operate within the maximum nacelle diameter constrain of 160 deg to allow installation under the wing. The impact of fan and nacelle technologies of the second generation ADP on fuel burn and direct operating costs for a typical 3000 nm mission was evaluated through use of a large, twin engine commercial airplane simulation model. The major emphasis of this study focused on fan blade aero/acoustic and structural technology evaluations and advanced nacelle designs. Results of this study have identified the testing required to verify the interactive performance of these components, along with noise characteristics, by wind tunnel testing utilizing and advanced interaction rig.

  7. NEXT Ion Propulsion System Configurations and Performance for Saturn System Exploration

    NASA Technical Reports Server (NTRS)

    Benson, Scott W.; Riehl, John P.; Oleson, Steven R.

    2007-01-01

    The successes of the Cassini/Huygens mission have heightened interest to return to the Saturn system with focused robotic missions. The desire for a sustained presence at Titan, through a dedicated orbiter and in-situ vehicle, either a lander or aerobot, has resulted in definition of a Titan Explorer flagship mission as a high priority in the Solar System Exploration Roadmap. The discovery of active water vapor plumes erupting from the tiger stripes on the moon Enceladus has drawn the attention of the space science community. The NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system is well suited to future missions to the Saturn system. NEXT is used within the inner solar system, in combination with a Venus or Earth gravity assist, to establish a fast transfer to the Saturn system. The NEXT system elements are accommodated in a separable Solar Electric Propulsion (SEP) module, or are integrated into the main spacecraft bus, depending on the mission architecture and performance requirements. This paper defines a range of NEXT system configurations, from two to four thrusters, and the Saturn system performance capability provided. Delivered mass is assessed parametrically over total trip time to Saturn. Launch vehicle options, gravity assist options, and input power level are addressed to determine performance sensitivities. A simple two-thruster NEXT system, launched on an Atlas 551, can deliver a spacecraft mass of over 2400 kg on a transfer to Saturn. Similarly, a four-thruster system, launched on a Delta 4050 Heavy, delivers more than 4000 kg spacecraft mass. A SEP module conceptual design, for a two thruster string, 17 kW solar array, configuration is characterized.

  8. Engineering method for aero-propulsive characteristics at hypersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Goradia, Suresh; Torres, Abel O.; Stack, Sharon H.; Everhart, Joel L.

    1991-01-01

    An engineering method has been developed for the rapid analysis of external aerodynamics and propulsive performance characteristics of airbreathing vehicles at hypersonic Mach numbers. This method, based on the theory of characteristics, has been developed to analyze fuselage-wing body combinations and body flaps with blunt or sharp leading/trailing edges. Arbitrary ratio of specific heat for the flowing medium can be specified in the program. Furthermore, the capability exists in the code to compute the inviscid inlet mass capture and momentum flux. The method is under development for computations of pressure distribution, and flow characteristics in the inlet, along with the effect of viscosity. Correlative studies have been performed for representative hypersonic configurations using the current method. The results of these correlations for various aerodynamics parameters are encouraging.

  9. Advances in Engine Test Capabilities at the NASA Glenn Research Center's Propulsion Systems Laboratory

    NASA Technical Reports Server (NTRS)

    Pachlhofer, Peter M.; Panek, Joseph W.; Dicki, Dennis J.; Piendl, Barry R.; Lizanich, Paul J.; Klann, Gary A.

    2006-01-01

    The Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Glenn Research Center is one of the premier U.S. facilities for research on advanced aeropropulsion systems. The facility can simulate a wide range of altitude and Mach number conditions while supplying the aeropropulsion system with all the support services necessary to operate at those conditions. Test data are recorded on a combination of steady-state and highspeed data-acquisition systems. Recently a number of upgrades were made to the facility to meet demanding new requirements for the latest aeropropulsion concepts and to improve operational efficiency. Improvements were made to data-acquisition systems, facility and engine-control systems, test-condition simulation systems, video capture and display capabilities, and personnel training procedures. This paper discusses the facility s capabilities, recent upgrades, and planned future improvements.

  10. An Introduction to Transient Engine Applications Using the Numerical Propulsion System Simulation (NPSS) and MATLAB

    NASA Technical Reports Server (NTRS)

    Chin, Jeffrey C.; Csank, Jeffrey T.; Haller, William J.; Seidel, Jonathan A.

    2016-01-01

    This document outlines methodologies designed to improve the interface between the Numerical Propulsion System Simulation framework and various control and dynamic analyses developed in the Matlab and Simulink environment. Although NPSS is most commonly used for steady-state modeling, this paper is intended to supplement the relatively sparse documentation on it's transient analysis functionality. Matlab has become an extremely popular engineering environment, and better methodologies are necessary to develop tools that leverage the benefits of these disparate frameworks. Transient analysis is not a new feature of the Numerical Propulsion System Simulation (NPSS), but transient considerations are becoming more pertinent as multidisciplinary trade-offs begin to play a larger role in advanced engine designs. This paper serves to supplement the relatively sparse documentation on transient modeling and cover the budding convergence between NPSS and Matlab based modeling toolsets. The following sections explore various design patterns to rapidly develop transient models. Each approach starts with a base model built with NPSS, and assumes the reader already has a basic understanding of how to construct a steady-state model. The second half of the paper focuses on further enhancements required to subsequently interface NPSS with Matlab codes. The first method being the simplest and most straightforward but performance constrained, and the last being the most abstract. These methods aren't mutually exclusive and the specific implementation details could vary greatly based on the designer's discretion. Basic recommendations are provided to organize model logic in a format most easily amenable to integration with existing Matlab control toolsets.

  11. Flight Investigation of the Effects of Ice on an I-16 Jet-Propulsion Engine

    NASA Technical Reports Server (NTRS)

    Pragliola, Philip C.; Werner, Milton

    1947-01-01

    A flight investigation of an I-16 jet propulsion engine installed in the waist compartment of a B-24M airplane was made to determine the effect of induction-system icing on the performance of the engine. Flights were made at inlet-air temperatures of 15 deg, 20 deg., and 25 F, an indicated airspeed of 180 miles per hour, jet-engine speeds of 13,000 and 15,000 rpm, liquid-water contents of approximately 0.3 to 0.5 gram per cubic meter, and an average water droplet size of approximately 50 microns. Under the most severe icing conditions obtained, ice formed on the screen over the front inlet to the compressor and obstructed about 70 percent of the front-inlet area. The thrust was thereby reduced 13.5 percent, the specific fuel consumption increased 17 percent, and the tail-pipe temperature increased 82 F. No icing of the rear compressor-inlet screen was encountered.

  12. Characteristics of a Hot Jet Discharged from a Jet-Propulsion Engine

    NASA Technical Reports Server (NTRS)

    Fleming, William A.

    1946-01-01

    An investigation of a heated jet was conducted in conjunction with tests of an axial-flow jet-propulsion engine in the Cleveland altitude wind tunnel. Pressure and temperature surveys were made across the jet 10 and 15 feet behind the jet-nozzle outlet of the engine. Surveys were obtained at pressure altitudes of 10,000, 20,000, 30,000, and 40,000 feet with test-section velocities from 30 to 110 feet per second and test-section temperatures from 60 F to -50 F. From measurements taken throughout the operable range of engine speeds, tail-pipe outlet temperatures from 500 F to 1250 F and jet velocities from 400 to 2200 feet per second were obtained. The jet-survey data presented extend the work previously done with low-velocity and low-temperature jets to the region of high velocities and high temperatures. The results obtained agree with previously determined experimental data and with predicted theoretical expressions for the dimensionless transverse velocity and temperature profiles across a jet. The spread of both the temperature and the velocity profiles was very nearly linear. Dimensionless plots of temperature and velocity along the axis of a heated jet agree with experimental results of tests with a cold jet.

  13. Attitude Control Flight Experience: Coping with Solar Radiation and Ion Engines Leak Thrust in Hayabusa (MUSES-C)

    NASA Technical Reports Server (NTRS)

    Kawaguchi, Jun'ichiro; Kominato, Takashi; Shirakawa, Ken'ichi

    2007-01-01

    The paper presents the attitude reorientation taking the advantage of solar radiation pressure without use of any fuel aboard. The strategy had been adopted to make Hayabusa spacecraft keep pointed toward the Sun for several months, while spinning. The paper adds the above mentioned results reported in Sedona this February showing another challenge of combining ion engines propulsion tactically balanced with the solar radiation torque with no spin motion. The operation has been performed since this March for a half year successfully. The flight results are presented with the estimated solar array panel diffusion coefficient and the ion engine's swirl torque.

  14. Design and evaluation of an integrated Quiet, Clean General Aviation Turbofan (QCGAT) engine and aircraft propulsion system

    NASA Technical Reports Server (NTRS)

    German, J.; Fogel, P.; Wilson, C.

    1980-01-01

    The design was based on the LTS-101 engine family for the core engine. A high bypass fan design (BPR=9.4) was incorporated to provide reduced fuel consumption for the design mission. All acoustic and pollutant emissions goals were achieved. A discussion of the preliminary design of a business jet suitable for the developed propulsion system is included. It is concluded that large engine technology can be successfully applied to small turbofans, and noise or pollutant levels need not be constraints for the design of future small general aviation turbofan engines.

  15. In-Flight Operation of the Dawn Ion Propulsion System Through Survey Science Orbit at Ceres

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.

    2015-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total delta V of 11 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional delta V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in September 2012 and concluded with the start of the approach to Ceres phase on December 26, 2015, leading to orbit capture on March 6, 2015. Deterministic thrusting continued during approach to place the spacecraft in its first science orbit, called RC3, which was achieved on April 23, 2015. Following science operations at RC3 ion thrusting was resumed for twenty-five days leading to arrival to the next science orbit, called survey orbit, on June 3, 2015. The IPS will be used for all subsequent orbit transfers and trajectory correction maneuvers until completion of the primary mission in approximately June 2016. To date the IPS has been operated for over 46,774 hours, consumed approximately 393 kg of xenon, and provided

  16. Space propulsion systems. Present performance limits and application and development trends

    NASA Technical Reports Server (NTRS)

    Buehler, R. D.; Lo, R. E.

    1981-01-01

    Typical spaceflight programs and their propulsion requirements as a comparison for possible propulsion systems are summarized. Chemical propulsion systems, solar, nuclear, or even laser propelled rockets with electrical or direct thermal fuel acceleration, nonrockets with air breathing devices and solar cells are considered. The chemical launch vehicles have similar technical characteristics and transportation costs. A possible improvement of payload by using air breathing lower stages is discussed. The electrical energy supply installations which give performance limits of electrical propulsion and the electrostatic ion propulsion systems are described. The development possibilities of thermal, magnetic, and electrostatic rocket engines and the state of development of the nuclear thermal rocket and propulsion concepts are addressed.

  17. Space Propulsion Technology Program Overview

    NASA Technical Reports Server (NTRS)

    Escher, William J. D.

    1991-01-01

    The topics presented are covered in viewgraph form. Focused program elements are: (1) transportation systems, which include earth-to-orbit propulsion, commercial vehicle propulsion, auxiliary propulsion, advanced cryogenic engines, cryogenic fluid systems, nuclear thermal propulsion, and nuclear electric propulsion; (2) space platforms, which include spacecraft on-board propulsion, and station keeping propulsion; and (3) technology flight experiments, which include cryogenic orbital N2 experiment (CONE), SEPS flight experiment, and cryogenic orbital H2 experiment (COHE).

  18. Experimental validation of the dual positive and negative ion beam acceleration in the plasma propulsion with electronegative gases thruster

    SciTech Connect

    Rafalskyi, Dmytro Popelier, Lara; Aanesland, Ane

    2014-02-07

    The PEGASES (Plasma Propulsion with Electronegative Gases) thruster is a gridded ion thruster, where both positive and negative ions are accelerated to generate thrust. In this way, additional downstream neutralization by electrons is redundant. To achieve this, the thruster accelerates alternately positive and negative ions from an ion-ion plasma where the electron density is three orders of magnitude lower than the ion densities. This paper presents a first experimental study of the alternate acceleration in PEGASES, where SF{sub 6} is used as the working gas. Various electrostatic probes are used to investigate the source plasma potential and the energy, composition, and current of the extracted beams. We show here that the plasma potential control in such system is key parameter defining success of ion extraction and is sensitive to both parasitic electron current paths in the source region and deposition of sulphur containing dielectric films on the grids. In addition, large oscillations in the ion-ion plasma potential are found in the negative ion extraction phase. The oscillation occurs when the primary plasma approaches the grounded parts of the main core via sub-millimetres technological inputs. By controlling and suppressing the various undesired effects, we achieve perfect ion-ion plasma potential control with stable oscillation-free operation in the range of the available acceleration voltages (±350 V). The measured positive and negative ion currents in the beam are about 10 mA for each component at RF power of 100 W and non-optimized extraction system. Two different energy analyzers with and without magnetic electron suppression system are used to measure and compare the negative and positive ion and electron fluxes formed by the thruster. It is found that at alternate ion-ion extraction the positive and negative ion energy peaks are similar in areas and symmetrical in position with +/− ion energy corresponding to the amplitude of the applied

  19. Nuclear propulsion for orbital transfer

    SciTech Connect

    Beale, G.A.; Lawrence, T.J. )

    1989-06-01

    The state of the art in nuclear propulsion for orbital transfer is discussed. Cryogenic propulsion, electric propulsion, solar-thermal propulsion and direct nuclear propulsion are examined in this context. New technologies with exceptional promise are addressed, emphasizing the particle test bed nuclear engine.

  20. Changes in the terrestrial atmosphere-ionosphere-magnetosphere system due to ion propulsion for solar power satellite placement

    NASA Technical Reports Server (NTRS)

    Curtis, S. A.; Grebowsky, J. M.

    1980-01-01

    In order to construct solar power satellites using earth-based materials, sections of a satellite must be lifted from low earth to geosynchronous orbit. The most plausible method of accomplishing this task is by means of ion propulsion based on the relatively abundant terrestrial atmospheric component, Ar. The proposed propulsion system will release a dense beam of about 5 keV Ar(+). The total amount of Ar(+) injected in transporting the components for each solar power satellite is comparable to the total ion content of the ionosphere-plasmasphere system while the total energy injected is larger than that of this system. Preliminary estimates are given of the effects massive Ar(+) injections have on the ionosphere-plasmasphere system with specific emphasis on potential communications disruptions. The effects stem from direct Ar(+) precipitation into the atmosphere and from Ar(+) beam induced precipitation of MeV radiation belt protons.

  1. The History of Ion Chromatography: The Engineering Perspective

    ERIC Educational Resources Information Center

    Evans, Barton

    2004-01-01

    The development of ion chromatography from an engineering perspective is presented. As ion chromatography became more widely accepted, researchers developed dozens of standard applications that enabled the creation of many low-end instruments.

  2. The History of Ion Chromatography: The Engineering Perspective

    ERIC Educational Resources Information Center

    Evans, Barton

    2004-01-01

    The development of ion chromatography from an engineering perspective is presented. As ion chromatography became more widely accepted, researchers developed dozens of standard applications that enabled the creation of many low-end instruments.

  3. Measurements of Forward Flight Effects on the Advanced Ducted Propulsion Demonstrator Engine

    NASA Technical Reports Server (NTRS)

    Horne, W. C.; Soderman, P. T.; Larkin, M.; Bock, L.; Olson, Lawrence (Technical Monitor)

    1994-01-01

    The performance of the Pratt & Whitney Advanced Ducted Propulsion (ADP) UHB concept has been recently evaluated with studies of a 17 in. diameter fan simulator. Following the model scale tests, a 118 in. diameter demonstrator was tested at the NASA Ames 40- by 80-Foot Wind Tunnel. The 18 blade fan was driven by the low compressor shaft of a PW2037 core through a reduction gear system fabricated by Fiat with approximately 1:3.7 reduction ratio. ne variable pitch fan was hydraulically actuated with settings for take-off, cruise, feather, and reverse thrust. The low-pressure turbine was built by MTU to provide higher shaft power in comparison with the standard PW2037. The demonstrator was provided with 45 vanes located 2.6 fan chords downstream of the rotor, and 10 case struts approximately 1 fan chord downstream of the vanes. The inlet, mid-duct, and exhaust linings were acoustically treated. Acoustic surveys were taken in the for-ward thrust mode for fan speeds of 898, 1120, 1205, and 1302 R.P.M., and at tunnel speeds of 25, 50, 100, and 140 kts. The lowest speed was achieved with the wind tunnel fans at flat pitch, but with the engine pumping the test section Microphone signals were recorded for 30 seconds at 5 deg. increments. These measurements will be used to assess the effects of forward speed on UHB engines, to compare these effects with the corresponding characteristics of conventional bypass ratio engines, and to discuss the various aspects of testing large engines in the wind tunnel.

  4. MITEE-B: A Compact Ultra Lightweight Bi-Modal Nuclear Propulsion Engine for Robotic Planetary Science Missions

    NASA Astrophysics Data System (ADS)

    Powell, James; Maise, George; Paniagua, John; Borowski, Stanley

    2003-01-01

    Nuclear thermal propulsion (NTP) enables unique new robotic planetary science missions that are impossible with chemical or nuclear electric propulsion systems. A compact and ultra lightweight bi-modal nuclear engine, termed MITEE-B (MInature ReacTor EnginE - Bi-Modal) can deliver 1000's of kilograms of propulsive thrust when it operates in the NTP mode, and many kilowatts of continuous electric power when it operates in the electric generation mode. The high propulsive thrust NTP mode enables spacecraft to land and takeoff from the surface of a planet or moon, to hop to multiple widely separated sites on the surface, and virtually unlimited flight in planetary atmospheres. The continuous electric generation mode enables a spacecraft to replenish its propellant by processing in-situ resources, provide power for controls, instruments, and communications while in space and on the surface, and operate electric propulsion units. Six examples of unique and important missions enabled by the MITEE-B engine are described, including: (1) Pluto lander and sample return; (2) Europa lander and ocean explorer; (3) Mars Hopper; (4) Jupiter atmospheric flyer; (5) SunBurn hypervelocity spacecraft; and (6) He3 mining from Uranus. Many additional important missions are enabled by MITEE-B. A strong technology base for MITEE-B already exists. With a vigorous development program, it could be ready for initial robotic science and exploration missions by 2010 AD. Potential mission benefits include much shorter in-space times, reduced IMLEO requirements, and replenishment of supplies from in-situ resources.

  5. Experimental Investigation of Airbreathing Laser Propulsion Engines: CO2TEA vs. EDL

    NASA Astrophysics Data System (ADS)

    Mori, Koichi; Sasoh, Akihiro; Myrabo, Leik N.

    2005-04-01

    Single pulse laboratory experiments were carried out with a high-power CO2 Transversely-Exited Atmospheric (TEA) laser using parabolic laser propulsion (LP) engines of historic interest: 1) an original Pirri/ AERL bell engine, and 2) a scaled-up 11-cm diameter version with identical geometry. The objective was to quantify the effects of pulse duration upon the impulse coupling coefficient performance — with pulse energy as the parametric variable. Performance data from the TEA laser are contrasted with former results derived from AVCO Everett Research Laboratory and PLVTS CO2 electron discharge lasers (EDL). The `short-pulse' 2-microsecond TEA laser tests generated results that were distinctively different from that of the `long-pulse' EDL sources. The TC-300 TEA laser employed an unstable resonator to deliver up to 380 joules, and the square output beam measured 15-cm on a side, with a hollow 8-cm center. A standard ballistic pendulum was employed to measure the performance.

  6. Modeling of a Turbofan Engine with Ice Crystal Ingestion in the NASA Propulsion System Laboratory

    NASA Technical Reports Server (NTRS)

    Veres, Joseph P.; Jorgenson, Philip C. E.; Jones, Scott M.; Nili, Samaun

    2017-01-01

    The main focus of this study is to apply a computational tool for the flow analysis of the turbine engine that has been tested with ice crystal ingestion in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The PSL has been used to test a highly instrumented Honeywell ALF502R-5A (LF11) turbofan engine at simulated altitude operating conditions. Test data analysis with an engine cycle code and a compressor flow code was conducted to determine the values of key icing parameters, that can indicate the risk of ice accretion, which can lead to engine rollback (un-commanded loss of engine thrust). The full engine aerothermodynamic performance was modeled with the Honeywell Customer Deck specifically created for the ALF502R-5A engine. The mean-line compressor flow analysis code, which includes a code that models the state of the ice crystal, was used to model the air flow through the fan-core and low pressure compressor. The results of the compressor flow analyses included calculations of the ice-water flow rate to air flow rate ratio (IWAR), the local static wet bulb temperature, and the particle melt ratio throughout the flow field. It was found that the assumed particle size had a large effect on the particle melt ratio, and on the local wet bulb temperature. In this study the particle size was varied parametrically to produce a non-zero calculated melt ratio in the exit guide vane (EGV) region of the low pressure compressor (LPC) for the data points that experienced a growth of blockage there, and a subsequent engine called rollback (CRB). At data points where the engine experienced a CRB having the lowest wet bulb temperature of 492 degrees Rankine at the EGV trailing edge, the smallest particle size that produced a non-zero melt ratio (between 3 percent - 4 percent) was on the order of 1 micron. This value of melt ratio was utilized as the target for all other subsequent data points analyzed, while the particle size was varied from 1 micron - 9

  7. Quantum Control Engineering with Trapped Ions

    NASA Astrophysics Data System (ADS)

    Biercuk, Michael

    2015-03-01

    Technologies fundamentally enabled by quantum mechanics are poised to transform a broad range of applications from computation to precision metrology over the coming decades. This talk will introduce a new field of research which is seeing concepts from control engineering translated to the domain of quantum mechanics in an effort to realize the full potential of engineered quantum technologies. We focus on understanding the physics underlying controlled quantum dynamics in the presence of rapidly fluctuating time-dependent Hamiltonians, leveraging the unique capabilities provided by trapped ions as a model quantum system. Our results introduce and experimentally validate generalized filter-transfer functions which cast arbitrary quantum control operations on qubits as noise spectral filters. We demonstrate the utility of these constructs for directly predicting the evolution of a quantum state in a realistic noisy environment, for developing novel robust control and sensing protocols, and for improving the stability of atomic clocks. This work demonstrates how quantum control can be leveraged to overcome some of the most challenging problems in quantum engineering, and even provide totally new functionality to quantum systems.

  8. The NASA-JPL advanced propulsion program

    NASA Technical Reports Server (NTRS)

    Frisbee, Robert H.

    1994-01-01

    The NASA Advanced Propulsion Concepts (APC) program at the Jet Propulsion Laboratory (JPL) consists of two main areas: The first involves cooperative modeling and research activities between JPL and various universities and industry; the second involves research at universities and industry that is directly supported by JPL. The cooperative research program consists of mission studies, research and development of ion engine technology using C-60 (Buckminsterfullerene) propellant, and research and development of lithium-propellant Lorentz-force accelerator (LFA) engine technology. The university/industry- supported research includes research (modeling and proof-of-concept experiments) in advanced, long-life electric propulsion, and in fusion propulsion. These propulsion concepts were selected primarily to cover a range of applications from near-term to far-term missions. For example, the long-lived pulsed-xenon thruster research that JPL is supporting at Princeton University addresses the near-term need for efficient, long-life attitude control and station-keeping propulsion for Earth-orbiting spacecraft. The C-60-propellant ion engine has the potential for good efficiency in a relatively low specific impulse (Isp) range (10,000 - 30,000 m/s) that is optimum for relatively fast (less than 100 day) cis-lunar (LEO/GEO/Lunar) missions employing near-term, high-specific mass electric propulsion vehicles. Research and modeling on the C-60-ion engine are currently being performed by JPL (engine demonstration), Caltech (C-60 properties), MIT (plume modeling), and USC (diagnostics). The Li-propellant LFA engine also has good efficiency in the modest Isp range (40,000 - 50,000 m/s) that is optimum for near-to-mid-term megawatt-class solar- and nuclear-electric propulsion vehicles used for Mars missions transporting cargo (in support of a piloted mission). Research and modeling on the Li-LFA engine are currently being performed by JPL (cathode development), Moscow Aviation

  9. In-Flight Operation of the Dawn Ion Propulsion System Through Orbit Capture at Vesta

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.; Brophy, John R.; Mikes, Steven C.

    2011-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from Cape Canaveral Air Force Station on September 27, 2007 on a Delta -II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218 kg spacecraft into an Earth-escape trajectory. Onboard the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide most of the ?V needed for heliocentric transfer to Vesta, orbit capture at Vesta, transfer among Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer among Ceres science orbits. The first 80 days after launch were dedicated to the initial checkout of the spacecraft which was followed by about ten months of full-power thrusting leading to a Mars gravity assist in February 2009 that provided 1 km/s of heliocentric energy increase and is the only part of the mission following launch in which a needed velocity change is not accomplished by the IPS. Deterministic thrusting for heliocentric transfer to Vesta resumed in June 2009 and was concluded with orbit capture at Vesta in July 2011. IPS was operated for approximately 23,400 hours, consumed approximately 250 kg of xenon, and provided a delta-V of approximately 6.7 km/s to achieve orbit capture at Vesta. IPS performance characteristics are very close to the expected performance characteristics based on analysis performed pre-launch. The only significant problem to have occurred over the almost four years of IPS operations in flight was the temporary failure of a valve driver board in DCIU-1, resulting in a loss of thrust of approximately 29 hours. Thrusting operations resumed after switching to DCIU-2, and power cycling conducted after orbit capture indicates DCIU-1 is completely operational. After about three weeks of survey operations IPS will be used to maneuver the

  10. In-Flight Operation of the Dawn Ion Propulsion System Through Orbit Capture at Vesta

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.; Brophy, John R.; Mikes, Steven C.

    2011-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from Cape Canaveral Air Force Station on September 27, 2007 on a Delta -II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218 kg spacecraft into an Earth-escape trajectory. Onboard the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide most of the ?V needed for heliocentric transfer to Vesta, orbit capture at Vesta, transfer among Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer among Ceres science orbits. The first 80 days after launch were dedicated to the initial checkout of the spacecraft which was followed by about ten months of full-power thrusting leading to a Mars gravity assist in February 2009 that provided 1 km/s of heliocentric energy increase and is the only part of the mission following launch in which a needed velocity change is not accomplished by the IPS. Deterministic thrusting for heliocentric transfer to Vesta resumed in June 2009 and was concluded with orbit capture at Vesta in July 2011. IPS was operated for approximately 23,400 hours, consumed approximately 250 kg of xenon, and provided a delta-V of approximately 6.7 km/s to achieve orbit capture at Vesta. IPS performance characteristics are very close to the expected performance characteristics based on analysis performed pre-launch. The only significant problem to have occurred over the almost four years of IPS operations in flight was the temporary failure of a valve driver board in DCIU-1, resulting in a loss of thrust of approximately 29 hours. Thrusting operations resumed after switching to DCIU-2, and power cycling conducted after orbit capture indicates DCIU-1 is completely operational. After about three weeks of survey operations IPS will be used to maneuver the

  11. Wear Mechanisms in Electron Sources for Ion Propulsion, 1: Neutralizer Hollow Cathode

    NASA Technical Reports Server (NTRS)

    Mikellides, Ioannis G.; Katz, Ira

    2008-01-01

    Upon the completion of two long-duration life tests of a 30-cm ion engine, the orifice channel of the neutralizer hollow cathode was eroded away to as much as twice its original diameter. Whereas the neutralizer cathode orifice opened significantly, no noticeable erosion of the discharge cathode orifice was observed. Noquantitative explanation of these erosion trends has been established since the completion of the two life tests. A two-dimensional model of the partially ionized gas inside these devices has been developed and applied to the neutralizer hollow cathode. The numerical simulations show that the main mechanism responsible for the channel erosion is sputtering by Xe+. These ions are accelerated by the sheath along the channel and bombard the surface with kinetic energy/charge of about 17 V at the beginning of cathode life. The density of the ions inside the neutralizer orifice is computed to be as high as 2.1 x 10(sup 22) m(sup -3). Because of the 3.5-times larger diameter of the discharge cathode orifice, the ion density inside the orifice is more than 40 times lower and the sheath drop 7 V lower compared with the values in the neutralizer. At these conditions, Xe+ can cause no significant sputtering of the surface.

  12. Wear Mechanisms in Electron Sources for Ion Propulsion, 1: Neutralizer Hollow Cathode

    NASA Technical Reports Server (NTRS)

    Mikellides, Ioannis G.; Katz, Ira

    2008-01-01

    Upon the completion of two long-duration life tests of a 30-cm ion engine, the orifice channel of the neutralizer hollow cathode was eroded away to as much as twice its original diameter. Whereas the neutralizer cathode orifice opened significantly, no noticeable erosion of the discharge cathode orifice was observed. Noquantitative explanation of these erosion trends has been established since the completion of the two life tests. A two-dimensional model of the partially ionized gas inside these devices has been developed and applied to the neutralizer hollow cathode. The numerical simulations show that the main mechanism responsible for the channel erosion is sputtering by Xe+. These ions are accelerated by the sheath along the channel and bombard the surface with kinetic energy/charge of about 17 V at the beginning of cathode life. The density of the ions inside the neutralizer orifice is computed to be as high as 2.1 x 10(sup 22) m(sup -3). Because of the 3.5-times larger diameter of the discharge cathode orifice, the ion density inside the orifice is more than 40 times lower and the sheath drop 7 V lower compared with the values in the neutralizer. At these conditions, Xe+ can cause no significant sputtering of the surface.

  13. Solar Thermal Propulsion

    NASA Technical Reports Server (NTRS)

    Gerrish, Harold P., Jr.

    2003-01-01

    This paper presents viewgraphs on Solar Thermal Propulsion (STP). Some of the topics include: 1) Ways to use Solar Energy for Propulsion; 2) Solar (fusion) Energy; 3) Operation in Orbit; 4) Propulsion Concepts; 5) Critical Equations; 6) Power Efficiency; 7) Major STP Projects; 8) Types of STP Engines; 9) Solar Thermal Propulsion Direct Gain Assembly; 10) Specific Impulse; 11) Thrust; 12) Temperature Distribution; 13) Pressure Loss; 14) Transient Startup; 15) Axial Heat Input; 16) Direct Gain Engine Design; 17) Direct Gain Engine Fabrication; 18) Solar Thermal Propulsion Direct Gain Components; 19) Solar Thermal Test Facility; and 20) Checkout Results.

  14. Mission design study of an RTG powered, ion engine equipped interstellar spacecraft

    NASA Astrophysics Data System (ADS)

    Fogel, Joshua A.

    This research explores a variety of mission and system architectures for an unmanned Interstellar Precursor Mission (IPM) spacecraft with a Radioisotope Thermoelectric Generator (RTG) powered Ion Engine using Xenon propellant, traveling on a (direct) ballistic escape trajectory to the undisturbed Interstellar Medium (˜200 AU). The main goal of this work was to determine the relationship between the propulsion system design parameters and the ensuing escape trajectory. To do this, an orbit simulator was created in Matlab using a fourth order Runge-Kutta numerical integration method to propagate the thrusting spacecraft's trajectory through time. The accelerations due to the Sun's gravity and the Ion Engine thrust were modeled separately and then combined into a single total acceleration vector at each time step, with the thrust direction assumed to be in the direction of the spacecraft's instantaneous velocity vector. The propellant of the thruster was also designed to be completely consumed by the time of engine cut-off (ECO), meaning a constant propellant mass flow rate. Simulations were run for burn times of 5, 10 & 15 years, with heliocentric launch velocities of 0, 5, 7, 10 & 12 km/sec from a circular 1 AU Earth orbit, and with RTG supplied engine input powers of 1000, 1500 & 2000 W. A total of 45 simulations were run for the circular 1 AU case, as well as additional comparison simulations for launches from an elliptical Earth orbit at perihelion and aphelion. The results of these simulations yielded many interesting results on the total fly-out times to 200 AU, which ranged dramatically from ˜35 to ˜140 years depending on the propulsion system settings and orbital initial conditions, as well as descriptions of the ECO distances from the Sun for each mission. The simulations also revealed the inherent gravitational maneuver inefficiency felt by all low thrust spacecraft, which becomes more apparent under certain conditions. Relations between launch velocity

  15. In-Space Propulsion Engine Architecture Based on Sublimation of Planetary Resources: From Exploration Robots to NED Mitigation

    NASA Technical Reports Server (NTRS)

    Sibille, Laurent; Mantovani, James G.

    2011-01-01

    Volatile solids occur naturally on most planetary bodies including the Moon, Mars, asteroids and comets. Examples of recent discoveries include water ice, frozen carbon dioxide and hydrocarbons. The ability to utilize readily available resources for in-space propulsion and for powering surface systems during a planetary mission will help minimize the overall cost and extend the op.erational life of a mission. The utilization of volatile solids to achieve these goals is attractive for its simplicity. We have investigated the potential of subliming in situ volatiles and silicate minerals to power propulsion engines for a wide range of in-space applications where environmental conditions are favorable. This paper addresses the' practicality of using planetary solid volatiles as a power source for propulsion and surface systems by presenting results of modeling involving thermodynamic and physical mechanics calculations, and laboratory testing to measure the thrust obtained from ,a volatile solid engine (VSE). Applications of a VSE for planetary exploration are discussed as a means for propulsion and for mechanical actuators and surface mobility platforms.

  16. Thermal Development Test of the NEXT PM1 Ion Engine

    NASA Technical Reports Server (NTRS)

    Anderson, John R.; Snyder, John S.; VanNoord, Jonathan L.; Soulas, George C.

    2010-01-01

    NASA's Evolutionary Xenon Thruster (NEXT) is a next-generation high-power ion propulsion system under development by NASA as a part of the In-Space Propulsion Technology Program. NEXT is designed for use on robotic exploration missions of the solar system using solar electric power. Potential mission destinations that could benefit from a NEXT Solar Electric Propulsion (SEP) system include inner planets, small bodies, and outer planets and their moons. This range of robotic exploration missions generally calls for ion propulsion systems with deep throttling capability and system input power ranging from 0.6 to 25 kW, as referenced to solar array output at 1 Astronomical Unit (AU). Thermal development testing of the NEXT prototype model 1 (PM1) was conducted at JPL to assist in developing and validating a thruster thermal model and assessing the thermal design margins. NEXT PM1 performance prior to, during and subsequent to thermal testing are presented. Test results are compared to the predicted hot and cold environments expected missions and the functionality of the thruster for these missions is discussed.

  17. Electric Propulsion Apparatus

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J. (Inventor)

    2013-01-01

    An electric propulsion machine includes an ion thruster having an annular discharge chamber housing an anode having a large surface area. The ion thruster includes flat annular ion optics with a small span to gap ratio. Optionally, a second electric propulsion thruster may be disposed in a cylindrical space disposed within an interior of the annulus.

  18. Derated ion thruster development status

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Haag, Thomas W.; Williams, George J., Jr.

    1993-01-01

    A 30 cm diameter xenon ion thruster is under development at NASA to provide an ion propulsion option for auxiliary and primary propulsion on missions of national interest. Specific efforts include thruster design optimizations, component life testing and validation, vibration testing, and performance characterizations. Under this program, the ion thruster will be brought to engineering model development status. The activities and preliminary test results to develop a 30 cm engineering model thruster are discussed.

  19. Validation Ice Crystal Icing Engine Test in the Propulsion Systems Laboratory at NASA Glenn Research Center

    NASA Technical Reports Server (NTRS)

    Oliver, Michael J.

    2014-01-01

    The Propulsion Systems Laboratory (PSL) is an existing altitude simulation jet engine test facility located at NASA Glenn Research Center in Clevleand, OH. It was modified in 2012 with the integration of an ice crystal cloud generation system. This paper documents the inaugural ice crystal cloud test in PSLthe first ever full scale, high altitude ice crystal cloud turbofan engine test to be conducted in a ground based facility. The test article was a Lycoming ALF502-R5 high bypass turbofan engine, serial number LF01. The objectives of the test were to validate the PSL ice crystal cloud calibration and engine testing methodologies by demonstrating the capability to calibrate and duplicate known flight test events that occurred on the same LF01 engine and to generate engine data to support fundamental and computational research to investigate and better understand the physics of ice crystal icing in a turbofan engine environment while duplicating known revenue service events and conducting test points while varying facility and engine parameters. During PSL calibration testing it was discovered than heated probes installed through tunnel sidewalls experienced ice buildup aft of their location due to ice crystals impinging upon them, melting and running back. Filtered city water was used in the cloud generation nozzle system to provide ice crystal nucleation sites. This resulted in mineralization forming on flow path hardware that led to a chronic degradation of performance during the month long test. Lacking internal flow path cameras, the response of thermocouples along the flow path was interpreted as ice building up. Using this interpretation, a strong correlation between total water content (TWC) and a weaker correlation between median volumetric diameter (MVD) of the ice crystal cloud and the rate of ice buildup along the instrumented flow path was identified. For this test article the engine anti-ice system was required to be turned on before ice crystal icing

  20. Validation Ice Crystal Icing Engine Test in the Propulsion Systems Laboratory at NASA Glenn Research Center

    NASA Technical Reports Server (NTRS)

    Oliver, Michael J.

    2014-01-01

    The Propulsion Systems Laboratory (PSL) is an existing altitude simulation jet engine test facility located at NASA Glenn Research Center in Cleveland, OH. It was modified in 2012 with the integration of an ice crystal cloud generation system. This paper documents the inaugural ice crystal cloud test in PSL--the first ever full scale, high altitude ice crystal cloud turbofan engine test to be conducted in a ground based facility. The test article was a Lycoming ALF502-R5 high bypass turbofan engine, serial number LF01. The objectives of the test were to validate the PSL ice crystal cloud calibration and engine testing methodologies by demonstrating the capability to calibrate and duplicate known flight test events that occurred on the same LF01 engine and to generate engine data to support fundamental and computational research to investigate and better understand the physics of ice crystal icing in a turbofan engine environment while duplicating known revenue service events and conducting test points while varying facility and engine parameters. During PSL calibration testing it was discovered than heated probes installed through tunnel sidewalls experienced ice buildup aft of their location due to ice crystals impinging upon them, melting and running back. Filtered city water was used in the cloud generation nozzle system to provide ice crystal nucleation sites. This resulted in mineralization forming on flow path hardware that led to a chronic degradation of performance during the month long test. Lacking internal flow path cameras, the response of thermocouples along the flow path was interpreted as ice building up. Using this interpretation, a strong correlation between total water content (TWC) and a weaker correlation between median volumetric diameter (MVD) of the ice crystal cloud and the rate of ice buildup along the instrumented flow path was identified. For this test article the engine anti-ice system was required to be turned on before ice crystal

  1. Propulsion Options For Interstellar Exploration

    NASA Technical Reports Server (NTRS)

    Johnson, Les; Leifer, Stephanie

    2000-01-01

    NASA is considering missions to explore near-interstellar space (40 - 250 Astronomical Units) early in the next decade as the first step toward a vigorous interstellar exploration program. A key enabling technology for such an ambitious science and exploration effort is a propulsion system capable of providing fast trip times, yet which has low enough mass to allow for the use of inexpensive launch vehicles. Advanced propulsion technologies that might support the First interstellar precursor mission by the end of the first decade of the new millennium include solar sails and nuclear electric propulsion. Solar sails and electric propulsion are two technology areas that may hold promise for the next generation of interstellar precursor missions as well - perhaps a thousand astronomical units traveled in a professional lifetime. Future missions to far beyond the Heliosphere will require the development of propulsion technologies that are only at the conceptual stage today. For years, the scientific community has been interested in solar sail and electric propulsion technologies to support robotic exploration of the solar system. Progress in thin-film materials fabrication and handling, and advancement in technologies that may enable the deployment of large sails in space are only now maturing to the point where ambitious interstellar precursor missions using sails can be considered. Xenon ion propulsion is now being demonstrated for planetary exploration by the Deep Space 1 mission. The primary issues for the adaptation of electric propulsion to interstellar precursor applications include the development of low specific mass nuclear power systems, engine lifetime, and high power operation. Recent studies of interstellar precursor mission scenarios that use these propulsion systems will be described, and the range of application of each technology will be explored.

  2. Nuclear Blast Response of Airbreathing Propulsion Systems. Laboratory Measurements with an Operational J-85-5 Turbojet Engine

    DTIC Science & Technology

    1980-03-31

    27 I.1 LIST OF ILLUSTRATIONS Page TABLE 1 EXPERIMENTAL CONDITIONS FOR WHICH DATA ARE AVAILABLE Figure No. 1 LUDWIEG-TUBE FACILITY...pressure ratio, turbine inlet temperature, and fuel/ air ratio) and the engine control system (measurement parameters or response characteristics...obtained under con- trolled laboratory conditions . At the present time, the major portion of the United States airbreathing propulsion system inventory

  3. Emission Inventories for Ocean-Going Vessels Using Category 3 Propulsion Engines in or Near the United States; Technical Support Document

    EPA Science Inventory

    This report describes the development of emission inventories for ocean-going vessels using Category 3 propulsion engines within the U.S. Exclusive Economic Zone. Inventories are presented for the 2002, 2020, and 2030 calendar years.

  4. Emission Inventories for Ocean-Going Vessels Using Category 3 Propulsion Engines in or Near the United States; Technical Support Document

    EPA Science Inventory

    This report describes the development of emission inventories for ocean-going vessels using Category 3 propulsion engines within the U.S. Exclusive Economic Zone. Inventories are presented for the 2002, 2020, and 2030 calendar years.

  5. Propulsion Research and Technology: Overview

    NASA Technical Reports Server (NTRS)

    Cole, John; Schmidt, George

    1999-01-01

    Propulsion is unique in being the main delimiter on how far and how fast one can travel in space. It is the lack of truly economical high-performance propulsion systems that continues to limit and restrict the extent of human endeavors in space. Therefore the goal of propulsion research is to conceive and investigate new, revolutionary propulsion concepts. This presentation reviews the development of new propulsion concepts. Some of these concepts are: (1) Rocket-based Combined Cycle (RBCC) propulsion, (2) Alternative combined Cycle engines suc2 as the methanol ramjet , and the liquid air cycle engines, (3) Laser propulsion, (4) Maglifter, (5) pulse detonation engines, (6) solar thermal propulsion, (7) multipurpose hydrogen test bed (MHTB) and other low-G cryogenic fluids, (8) Electric propulsion, (9) nuclear propulsion, (10) Fusion Propulsion, and (11) Antimatter technology. The efforts of the NASA centers in this research is also spotlighted.

  6. Electric propulsion system technology

    NASA Astrophysics Data System (ADS)

    Brophy, John R.; Garner, Charles E.; Goodfellow, Keith D.; Pivirotto, Thomas J.; Polk, James E.

    1992-11-01

    The work performed in fiscal year (FY) 1991 under the Propulsion Technology Program RTOP (Research and Technology Objectives and Plans) No. (55) 506-42-31 for Low-Thrust Primary and Auxiliary Propulsion technology development is described. The objectives of this work fall under two broad categories. The first of these deals with the development of ion engines for primary propulsion in support of solar system exploration. The second with the advancement of steady-state magnetoplasmadynamic (MPD) thruster technology at 100 kW to multimegawatt input power levels. The major technology issues for ion propulsion are demonstration of adequate engine life at the 5 to 10 kW power level and scaling ion engines to power levels of tens to hundreds of kilowatts. Tests of a new technique in which the decelerator grid of a three-grid ion accelerator system is biased negative of neutralizer common potential in order to collect facility induced charge-exchange ions are described. These tests indicate that this SAND (Screen, Accelerator, Negative Decelerator) configuration may enable long duration ion engine endurance tests to be performed at vacuum chamber pressures an order of magnitude higher than previously possible. The corresponding reduction in pumping speed requirements enables endurance tests of 10 kW class ion engines to be performed within the resources of existing technology programs. The results of a successful 5,000-hr endurance of a xenon hollow cathode operating at an emission current of 25 A are described, as well as the initial tests of hollow cathodes operating on a mixture of argon and 3 percent nitrogen. Work performed on the development of carbon/carbon grids, a multi-orifice hollow cathode, and discharge chamber erosion reduction through the addition of nitrogen are also described. Critical applied-field MPD thruster technical issues remain to be resolved, including demonstration of reliable steady-state operation at input powers of hundreds to thousands of

  7. Electric propulsion system technology

    NASA Technical Reports Server (NTRS)

    Brophy, John R.; Garner, Charles E.; Goodfellow, Keith D.; Pivirotto, Thomas J.; Polk, James E.

    1992-01-01

    The work performed in fiscal year (FY) 1991 under the Propulsion Technology Program RTOP (Research and Technology Objectives and Plans) No. (55) 506-42-31 for Low-Thrust Primary and Auxiliary Propulsion technology development is described. The objectives of this work fall under two broad categories. The first of these deals with the development of ion engines for primary propulsion in support of solar system exploration. The second with the advancement of steady-state magnetoplasmadynamic (MPD) thruster technology at 100 kW to multimegawatt input power levels. The major technology issues for ion propulsion are demonstration of adequate engine life at the 5 to 10 kW power level and scaling ion engines to power levels of tens to hundreds of kilowatts. Tests of a new technique in which the decelerator grid of a three-grid ion accelerator system is biased negative of neutralizer common potential in order to collect facility induced charge-exchange ions are described. These tests indicate that this SAND (Screen, Accelerator, Negative Decelerator) configuration may enable long duration ion engine endurance tests to be performed at vacuum chamber pressures an order of magnitude higher than previously possible. The corresponding reduction in pumping speed requirements enables endurance tests of 10 kW class ion engines to be performed within the resources of existing technology programs. The results of a successful 5,000-hr endurance of a xenon hollow cathode operating at an emission current of 25 A are described, as well as the initial tests of hollow cathodes operating on a mixture of argon and 3 percent nitrogen. Work performed on the development of carbon/carbon grids, a multi-orifice hollow cathode, and discharge chamber erosion reduction through the addition of nitrogen are also described. Critical applied-field MPD thruster technical issues remain to be resolved, including demonstration of reliable steady-state operation at input powers of hundreds to thousands of

  8. Electric Propulsion Laboratory Vacuum Chamber

    NASA Image and Video Library

    1964-06-21

    Engineer Paul Reader and his colleagues take environmental measurements during testing of a 20-inch diameter ion engine in a vacuum tank at the Electric Propulsion Laboratory (EPL). Researchers at the Lewis Research Center were investigating the use of a permanent-magnet circuit to create the magnetic field required power electron bombardment ion engines. Typical ion engines use a solenoid coil to create this magnetic field. It was thought that the substitution of a permanent magnet would create a comparable magnetic field with a lower weight. Testing of the magnet system in the EPL vacuum tanks revealed no significant operational problems. Reader found the weight of the two systems was similar, but that the thruster’s efficiency increased with the magnet. The EPL contained a series of large vacuum tanks that could be used to simulate conditions in space. Large vacuum pumps reduced the internal air pressure, and a refrigeration system created the cryogenic temperatures found in space.

  9. In-Flight Operation of the Dawn Ion Propulsion System Through Year Two of Cruise to Ceres

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.

    2014-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total (Delta)V of 11.3 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits.

  10. In-Flight Operation of the Dawn Ion Propulsion System Through Year Two of Cruise to Ceres

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.

    2014-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total (Delta)V of 11.3 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits.

  11. Performance studies on the application of four-engine and two-engine USB propulsive lift to the E-2C aircraft

    NASA Technical Reports Server (NTRS)

    Riddle, D. W.; Stevens, V. C.

    1986-01-01

    A study has been completed of the performance benefits to be derived from applying advanced upper-surface blowing (USB) propulsive-lift technology to the E-2C aircraft. The results of comparing four-engine with two-engine USB configurations are discussed, and engine sizing and aerodynamic/structural considerations pertaining to the E-2C/USB modification are examined. The effects of the modification on performance are described in detail with regard to takeoff distance and landing distance estimation in free-deck operations, operations using catapult and arresting gear, ceiling and radar surveillance missions, and range and endurance capability.

  12. Performance Evaluation of 40 cm Ion Optics for the NEXT Ion Engine

    NASA Technical Reports Server (NTRS)

    Soulas, George C.; Haag, Thomas W.; Patterson, Michael J.

    2002-01-01

    The results of performance tests with two 40 cm ion optics sets are presented and compared to those of 30 cm ion optics with similar aperture geometries. The 40 cm ion optics utilized both NSTAR and TAG (Thick-Accelerator-Grid) aperture geometries. All 40 cm ion optics tests were conducted on a NEXT (NASA's Evolutionary Xenon Thruster) laboratory model ion engine. Ion optics performance tests were conducted over a beam current range of 1.20 to 3.52 A and an engine input power range of 1.1 to 6.9 kW. Measured ion optics' performance parameters included near-field radial beam current density profiles, impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, beam divergence angles, and start-up transients. Impingement-limited total voltages for 40 cm ion optics with the NSTAR aperture geometry were 60 to 90 V lower than those with the TAG aperture geometry. This difference was speculated to be due to an incomplete burn-in of the TAG ion optics. Electron backstreaming limits for the 40 cm ion optics with the TAG aperture geometry were 8 to 19 V higher than those with the NSTAR aperture geometry due to the thicker accelerator grid of the TAG geometry. Because the NEXT ion engine provided beam flatness parameters that were 40 to 63 percent higher than those of the NSTAR ion engine, the 40 cm ion optics outperformed the 30 cm ion optics.

  13. Dual-fuel propulsion - Why it works, possible engines, and results of vehicle studies. [on earth-to-orbit Space Shuttle flights

    NASA Technical Reports Server (NTRS)

    Martin, J. A.; Wilhite, A. W.

    1979-01-01

    The reasons why dual-fuel propulsion works are discussed. Various engine options are discussed, and vehicle mass and cost results are presented for earth-to-orbit vehicles. The results indicate that dual-fuel propulsion is attractive, particularly with the dual-expander engine. A unique orbit-transfer vehicle is described which uses dual-fuel propulsion. One Space Shuttle flight and one flight of a heavy-lift Shuttle derivative are used for each orbit-transfer vehicle flight, and the payload capability is quite attractive.

  14. Ion Engine With Solid-Electrolyte Ion Generator

    NASA Technical Reports Server (NTRS)

    Richter, R.

    1984-01-01

    Working fluid utilized efficiently. Working fluid positive ions conducted through solid electrolyte to outside, then accelerated in external electric field. While in solid-electrolyte material, ions do not recombine with electrons: transported to surface with high ionization efficiency. Provides new way to generate beam of ions for implantation in semiconductors or other applications.

  15. Ion Engine With Solid-Electrolyte Ion Generator

    NASA Technical Reports Server (NTRS)

    Richter, R.

    1984-01-01

    Working fluid utilized efficiently. Working fluid positive ions conducted through solid electrolyte to outside, then accelerated in external electric field. While in solid-electrolyte material, ions do not recombine with electrons: transported to surface with high ionization efficiency. Provides new way to generate beam of ions for implantation in semiconductors or other applications.

  16. NEXT Ion Engine 2000 Hour Wear Test Results

    NASA Technical Reports Server (NTRS)

    Soulas, George C.; Kamhawi, Hani; Patterson, Michael J.; Britton, Melissa A.; Frandina, Michael M.

    2004-01-01

    The results of the NEXT 2000 h wear test are presented. This test was conducted with a 40 cm engineering model ion engine, designated EM1, at a 3.52 A beam current and 1800 V beam power supply voltage. Performance tests, which were conducted over a throttling range of 1.1 to 6.9 kW throughout the wear test, demonstrated that EM1 satisfied all thruster performance requirements. The ion engine accumulated 2038 h of operation at a thruster input power of 6.9 kW, processing 43 kg of xenon. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, was steady with no indications of performance degradation. The ion engine was also inspected following the test. This paper presents these findings.

  17. Specialized data analysis for the Space Shuttle Main Engine and diagnostic evaluation of advanced propulsion system components

    NASA Technical Reports Server (NTRS)

    1993-01-01

    The Marshall Space Flight Center is responsible for the development and management of advanced launch vehicle propulsion systems, including the Space Shuttle Main Engine (SSME), which is presently operational, and the Space Transportation Main Engine (STME) under development. The SSME's provide high performance within stringent constraints on size, weight, and reliability. Based on operational experience, continuous design improvement is in progress to enhance system durability and reliability. Specialized data analysis and interpretation is required in support of SSME and advanced propulsion system diagnostic evaluations. Comprehensive evaluation of the dynamic measurements obtained from test and flight operations is necessary to provide timely assessment of the vibrational characteristics indicating the operational status of turbomachinery and other critical engine components. Efficient performance of this effort is critical due to the significant impact of dynamic evaluation results on ground test and launch schedules, and requires direct familiarity with SSME and derivative systems, test data acquisition, and diagnostic software. Detailed analysis and evaluation of dynamic measurements obtained during SSME and advanced system ground test and flight operations was performed including analytical/statistical assessment of component dynamic behavior, and the development and implementation of analytical/statistical models to efficiently define nominal component dynamic characteristics, detect anomalous behavior, and assess machinery operational condition. In addition, the SSME and J-2 data will be applied to develop vibroacoustic environments for advanced propulsion system components, as required. This study will provide timely assessment of engine component operational status, identify probable causes of malfunction, and indicate feasible engineering solutions. This contract will be performed through accomplishment of negotiated task orders.

  18. Note: An advanced in situ diagnostic system for characterization of electric propulsion thrusters and ion beam sources

    NASA Astrophysics Data System (ADS)

    Bundesmann, C.; Tartz, M.; Scholze, F.; Leiter, H. J.; Scortecci, F.; Gnizdor, R. Y.; Neumann, H.

    2010-04-01

    We present an advanced diagnostic system for in situ characterization of electric propulsion thrusters and ion beam sources. The system uses a high-precision five-axis positioning system with a modular setup and the following diagnostic tools: a telemicroscopy head for optical imaging, a triangular laser head for surface profile scanning, a pyrometer for temperature scanning, a Faraday probe for current density mapping, and an energy-selective mass spectrometer for beam characterization (energy and mass distribution, composition). The capabilities of our diagnostic system are demonstrated with a Hall effect thruster SPT-100D EM1.

  19. Note: An advanced in situ diagnostic system for characterization of electric propulsion thrusters and ion beam sources.

    PubMed

    Bundesmann, C; Tartz, M; Scholze, F; Leiter, H J; Scortecci, F; Gnizdor, R Y; Neumann, H

    2010-04-01

    We present an advanced diagnostic system for in situ characterization of electric propulsion thrusters and ion beam sources. The system uses a high-precision five-axis positioning system with a modular setup and the following diagnostic tools: a telemicroscopy head for optical imaging, a triangular laser head for surface profile scanning, a pyrometer for temperature scanning, a Faraday probe for current density mapping, and an energy-selective mass spectrometer for beam characterization (energy and mass distribution, composition). The capabilities of our diagnostic system are demonstrated with a Hall effect thruster SPT-100D EM1.

  20. Modeling of Highly Instrumented Honeywell Turbofan Engine Tested with Ice Crystal Ingestion in the NASA Propulsion System Laboratory

    NASA Technical Reports Server (NTRS)

    Veres, Joseph P.; Jorgenson, Philip C. E.; Jones, Scott M.

    2016-01-01

    The Propulsion Systems Laboratory (PSL), an altitude test facility at NASA Glenn Research Center, has been used to test a highly instrumented turbine engine at simulated altitude operating conditions. This is a continuation of the PSL testing that successfully duplicated the icing events that were experienced in a previous engine (serial LF01) during flight through ice crystal clouds, which was the first turbofan engine tested in PSL. This second model of the ALF502R-5A serial number LF11 is a highly instrumented version of the previous engine. The PSL facility provides a continuous cloud of ice crystals with controlled characteristics of size and concentration, which are ingested by the engine during operation at simulated altitudes. Several of the previous operating points tested in the LF01 engine were duplicated to confirm repeatability in LF11. The instrumentation included video cameras to visually illustrate the accretion of ice in the low pressure compressor (LPC) exit guide vane region in order to confirm the ice accretion, which was suspected during the testing of the LF01. Traditional instrumentation included static pressure taps in the low pressure compressor inner and outer flow path walls, as well as total pressure and temperature rakes in the low pressure compressor region. The test data was utilized to determine the losses and blockages due to accretion in the exit guide vane region of the LPC. Multiple data points were analyzed with the Honeywell Customer Deck. A full engine roll back point was modeled with the Numerical Propulsion System Simulation (NPSS) code. The mean line compressor flow analysis code with ice crystal modeling was utilized to estimate the parameters that indicate the risk of accretion, as well as to estimate the degree of blockage and losses caused by accretion during a full engine roll back point. The analysis provided additional validation of the icing risk parameters within the LPC, as well as the creation of models for

  1. Thermionic reactor ion propulsion system /TRIPS/ - Its multi-mission capability.

    NASA Technical Reports Server (NTRS)

    Peelgren, M. L.

    1972-01-01

    The unmanned planetary exploration to be conducted in the last two decades of this century includes many higher energy missions which tax all presently available propulsion systems beyond their limit. One candidate with the versatility and performance to meet these mission objectives is nuclear electric propulsion (NEP). Additionally, the NEP System is feasible in orbit raising operations with the Shuttle or Shuttle/Tug combination. A representative planetary mission is described (Uranus-Neptune flyby with probe), and geocentric performance and tradeoffs are discussed. The NEP System is described in more detail with particular emphasis on the power subsystem consisting of the thermionic reactor, heat rejection subsystem, and neutron shield.

  2. NASA electric propulsion technology

    NASA Technical Reports Server (NTRS)

    Berkopec, F. D.; Stone, J. R.; Aston, G.

    1985-01-01

    It is pointed out that the requirements for future electric propulsion cover an extremely large range of technical and programmatic characteristics. A NASA program is to provide options for the many potential mission applications, taking into account work on electrostatic, electromagnetic, and electrothermal propulsion systems. The present paper is concerned with developments regarding the three classes of electric propulsion. Studies concerning electrostatic propulsion are concerned with ion propulsion for primary propulsion for planetary and earth-orbit transfer vehicles, stationkeeping for geosynchronous spacecraft, and ion thruster systems. In connection with investigations related to electromagnetic propulsion, attention is given to electromagnetic launchers, the Hall current thruster, and magnetoplasmadynamic thrusters. In a discussion of electrothermal developments, space station resistojets are considered along with high performance resistojets, arcjets, and a laser thruster.

  3. Advanced space propulsion study - antiproton and beamed-power propulsion. Final report, 1 May 1986-30 June 1987

    SciTech Connect

    Forward, R.L.

    1987-10-01

    The contract objective was to monitor the research at the forefront of physics and engineering to discover new spacecraft-propulsion concepts. The major topics covered were antiproton-annihilation propulsion, laser thermal propulsion, laser-pushed lightsails, tether transportation systems, solar sails, and metallic hydrogen. Five papers were prepared and are included as appendices. They covered 1) pellet, microwave, and laser-beamed power systems for interstellar transport; 2) a design for a near-relativistic laser-pushed lightsail using near-term laser technology; 3) a survey of laser thermal propulsion, tether transportation systems, antiproton annihilation propulsion, exotic applications of solar sails, and laser-pushed interstellar lightsails; 4) the status of antiproton annihilation propulsion as of 1986, and 5) the prospects for obtaining antimatter ions heavier than antiprotons. Two additional appendices contain the first seven issues of the Mirror Matter Newsletter concerning the science and technology of antimatter, and an annotated bibliography of antiproton science and technology.

  4. Changes in the terrestrial atmosphere-ionosphere-magnetosphere system due to ion propulsion for solar power satellite placement

    NASA Technical Reports Server (NTRS)

    Curtis, S. A.; Grebowsky, J. M.

    1979-01-01

    Preliminary estimates of the effects massive Ar(+) injections on the ionosphere-plasmasphere system with specific emphasis on potential communications disruptions are given. The effects stem from direct Ar(+) precipitation into the atmosphere and from Ar(+) beam induced precipitation of MeV radiation belt protons. These injections result from the construction of Solar Power Satellites using earth-based materials in which sections of a satellite must be lifted from low earth to geosynchronous orbit by means of ion propulsion based on the relatively abundant terrestrial atmospheric component, Ar. The total amount of Ar(+) injected in transporting the components for each Solar Power Satellite is comparable to the total ion content of the ionosphere-plasmasphere system while the total energy injected is larger than that of this system. It is suggested that such effects may be largely eliminated by using lunar-based rather than earth-based satellite construction materials.

  5. Engineering systems designs for a recirculating heavy ion induction accelerator

    SciTech Connect

    Newton, M.A.; Barnard, J.J.; Reginato, L.L.; Yu, S.S.

    1991-05-01

    Recirculating heavy ion induction accelerators are being investigated as possible drivers for heavy ion fusion. Part of this investigation has included the generation of a conceptual design for a recirculator system. This paper will describe the overall engineering conceptual design of this recirculator, including discussions of the dipole magnet system, the superconducting quadrupole system and the beam acceleration system. Major engineering issues, evaluation of feasibility, and cost tradeoffs of the complete recirculator system will be presented and discussed. 5 refs., 4 figs.

  6. Nuclear blast response of airbreathing propulsion systems: laboratory measurements with an operational J-85-5 turbojet engine

    SciTech Connect

    Dunn, M.G.; Rafferty, J.M.

    1982-07-01

    This paper describes an experimental technique for controlled laboratory measurements of the nuclear blast response of airbreathing propulsion systems. The experiments utilize an available G.E. J-855 turbojet engine located in the test section of the Calspan Ludwieg-tube facility. Significant modifications were made to this facility in order to adapt it to the desired configuration. The J-85 engine had previously been used at Calspan for other purposes and thus came equipped with eight pressure transducers at four axial locations along the compressor section. These transducers have a frequency response on the order of 40 KHz. Pressure histories obtained at several circumferential and axial locations along the compressor are presented for blastwave equivalent overpressures up to 17.2 kPa (2.5 psi) at corrected engine speeds on the order of 94 percent of maximum speed.

  7. Electric propulsion system technology

    NASA Astrophysics Data System (ADS)

    Brophy, John R.; Garner, Charles E.; Goodfellow, Keith D.

    1991-12-01

    The work performed on the Ion Propulsion System Technology Task in FY90 is described. The objectives of this work fall under two broad categories. The first of these deals with issues associated with the application of xenon ion thrusters for primary propulsion of planetary spacecraft, and the second with the investigation of technologies which will facilitate the development of larger, higher power ion thrusters to support more advanced mission applications. Most of the effort was devoted to investigation of the critical issues associated with the use of ion thrusters for planetary spacecraft. These issues may be succinctly referred to as life time, system integration, and throttling. Chief among these is the engine life time. If the engines do not have sufficient life to perform the missions of interest, then the other issues become unimportant. Ion engine life time was investigated through two experimental programs: an investigation into the reduction of ion engine internal sputter erosion through the addition of small quantities of nitrogen, and a long duration cathode life test. In addition, a literature review and analysis of accelerator grid erosion were performed. The nitrogen addition tests indicated that the addition of between 0.5 and 1.0 percent of nitrogen by mass to the xenon propellant results in a reduction in the sputter erosion of discharge chamber components by a factor of between 20 and 50, with negligible reduction in thruster performance. The long duration test of a 6.35-mm dia. xenon hollow cathode is still in progress, and has accumulated more than 4,000 hours of operation at an emission current of 25 A at the time of this writing. One of the major system integration issues concerns possible interactions of the ion thruster produced charge exchange plasma with the spacecraft. A computer model originally developed to describe the behavior of mercury ion thruster charge exchange plasmas was resurrected and modified for xenon propellant. This

  8. Electric propulsion system technology

    NASA Technical Reports Server (NTRS)

    Brophy, John R.; Garner, Charles E.; Goodfellow, Keith D.

    1991-01-01

    The work performed on the Ion Propulsion System Technology Task in FY90 is described. The objectives of this work fall under two broad categories. The first of these deals with issues associated with the application of xenon ion thrusters for primary propulsion of planetary spacecraft, and the second with the investigation of technologies which will facilitate the development of larger, higher power ion thrusters to support more advanced mission applications. Most of the effort was devoted to investigation of the critical issues associated with the use of ion thrusters for planetary spacecraft. These issues may be succinctly referred to as life time, system integration, and throttling. Chief among these is the engine life time. If the engines do not have sufficient life to perform the missions of interest, then the other issues become unimportant. Ion engine life time was investigated through two experimental programs: an investigation into the reduction of ion engine internal sputter erosion through the addition of small quantities of nitrogen, and a long duration cathode life test. In addition, a literature review and analysis of accelerator grid erosion were performed. The nitrogen addition tests indicated that the addition of between 0.5 and 1.0 percent of nitrogen by mass to the xenon propellant results in a reduction in the sputter erosion of discharge chamber components by a factor of between 20 and 50, with negligible reduction in thruster performance. The long duration test of a 6.35-mm dia. xenon hollow cathode is still in progress, and has accumulated more than 4,000 hours of operation at an emission current of 25 A at the time of this writing. One of the major system integration issues concerns possible interactions of the ion thruster produced charge exchange plasma with the spacecraft. A computer model originally developed to describe the behavior of mercury ion thruster charge exchange plasmas was resurrected and modified for xenon propellant. This

  9. Initial Test Firing Results for Solid CO/GOX Cryogenic Hybrid Rocket Engine for Mars ISRU Propulsion Applications

    NASA Technical Reports Server (NTRS)

    Rice, Eric E.; St. Clair, Christopher P.; Chiaverini, Martin J.; Knuth, William H.; Gustafson, Robert J.; Gramer, Daniel J.

    1999-01-01

    ORBITEC is developing methods for producing, testing, and utilizing Mars-based ISRU fuel/oxidizer combinations to support low cost, planetary surface and flight propulsion and power systems. When humans explore Mars we will need to use in situ resources that are available, such as: energy (solar); gases or liquids for life support, ground transportation, and flight to and from other surface locations and Earth; and materials for shielding and building habitats and infrastructure. Probably the easiest use of Martian resources to reduce the cost of human exploration activities is the use of the carbon and oxygen readily available from the CO2 in the Mars atmosphere. ORBITEC has conducted preliminary R&D that will eventually allow us to reliably use these resources. ORBITEC is focusing on the innovative use of solid CO as a fuel. A new advanced cryogenic hybrid rocket propulsion system is suggested that will offer advantages over LCO/LOX propulsion, making it the best option for a Mars sample return vehicle and other flight vehicles. This technology could also greatly support logistics and base operations by providing a reliable and simple way to store solar or nuclear generated energy in the form of chemical energy that can be used for ground transportation (rovers/land vehicles) and planetary surface power generators. This paper describes the overall concept and the test results of the first ever solid carbon monoxide/oxygen rocket engine firing.

  10. Initial Test Firing Results for Solid CO/GOX Cryogenic Hybrid Rocket Engine for Mars ISRU Propulsion Applications

    NASA Technical Reports Server (NTRS)

    Rice, Eric E.; St. Clair, Christopher P.; Chiaverini, Martin J.; Knuth, William H.; Gustafson, Robert J.; Gramer, Daniel J.

    1999-01-01

    ORBITEC is developing methods for producing, testing, and utilizing Mars-based ISRU fuel/oxidizer combinations to support low cost, planetary surface and flight propulsion and power systems. When humans explore Mars we will need to use in situ resources that are available, such as: energy (solar); gases or liquids for life support, ground transportation, and flight to and from other surface locations and Earth; and materials for shielding and building habitats and infrastructure. Probably the easiest use of Martian resources to reduce the cost of human exploration activities is the use of the carbon and oxygen readily available from the CO2 in the Mars atmosphere. ORBITEC has conducted preliminary R&D that will eventually allow us to reliably use these resources. ORBITEC is focusing on the innovative use of solid CO as a fuel. A new advanced cryogenic hybrid rocket propulsion system is suggested that will offer advantages over LCO/LOX propulsion, making it the best option for a Mars sample return vehicle and other flight vehicles. This technology could also greatly support logistics and base operations by providing a reliable and simple way to store solar or nuclear generated energy in the form of chemical energy that can be used for ground transportation (rovers/land vehicles) and planetary surface power generators. This paper describes the overall concept and the test results of the first ever solid carbon monoxide/oxygen rocket engine firing.

  11. PC programs for the prediction of the linear stability behavior of liquid propellant propulsion systems and application to current MSFC rocket engine test programs, volume 1

    NASA Technical Reports Server (NTRS)

    Doane, George B., III; Armstrong, W. C.

    1990-01-01

    Research on propulsion stability (chugging and acoustic modes), and propellant valve control was investigated. As part of the activation of the new liquid propulsion test facilities, it is necessary to analyze total propulsion system stability. To accomplish this, several codes were built to run on desktop 386 machines. These codes enable one to analyze the stability question associated with the propellant feed systems. In addition, further work was adapted to this computing environment and furnished along with other codes. This latter inclusion furnishes those interested in high frequency oscillatory combustion behavior (that does not couple to the feed system) a set of codes for study of proposed liquid rocket engines.

  12. The methodology of variable management of propellant fuel consumption by jet-propulsion engines of a spacecraft

    NASA Astrophysics Data System (ADS)

    Kovtun, V. S.

    2012-12-01

    Traditionally, management of propellant fuel consumption on board of a spacecraft is only associated with the operation of jet-propulsion engines (JPE) that are actuator devices of motion control systems (MCS). The efficiency of propellant fuel consumption depends not only on the operation of the MCS, but also, to one extent or another, on all systems functioning on board of a spacecraft, and on processes that occur in them and involve conversion of variable management of propellant fuel consumption by JPEs as a constituent part of the control of the complex process of spacecraft flight.

  13. Electric propulsion - A high energy capability for solar system exploration

    NASA Technical Reports Server (NTRS)

    Atkins, K. L.

    1976-01-01

    The principles of spacecraft electric (ion thruster) propulsion are briefly reviewed. Attention is given to the inner and outer planet applications of electric (and solar electric) propulsion. Electric propulsion is considered as a stepping stone to nuclear electric propulsion.

  14. Additive Manufacturing of Aerospace Propulsion Components

    NASA Technical Reports Server (NTRS)

    Misra, Ajay K.; Grady, Joseph E.; Carter, Robert

    2015-01-01

    The presentation will provide an overview of ongoing activities on additive manufacturing of aerospace propulsion components, which included rocket propulsion and gas turbine engines. Future opportunities on additive manufacturing of hybrid electric propulsion components will be discussed.

  15. Numerical Prediction of Grid Erosion of Ion Engine

    NASA Astrophysics Data System (ADS)

    Miyasaka, Takeshi; Kobayashi, Tsutomu; Asato, Katsuo

    With the increase of long-term space missions, the evaluation of lifetime of ion engines by numerical analyses becomes important. In order to develop a numerical code for the evaluation of ion engine lifetime, JIEDI (JAXA Ion Engine Development Initiative) tool development has been started. To evaluate the validities of boundary conditions such as upstream discharge region and downstream region conditions, a 3-dimensional full-particle code was developed. In the present study, the effects of the electron mass model introduced for shortening the calculation time were investigated. We found that there are differences in the distributions of charged particles and electric potential profiles in the downstream region among different electron masses. Consequently, the effects of electron mass on the energy peak of the ions impacting on the grid and the erosion distribution on the downstream surface of the accel grid were observed.

  16. Technologies to improve ion propulsion system performance, life and efficiency for NEP

    NASA Technical Reports Server (NTRS)

    Katz, I.; Brophy, J. R.; Anderson, J. R.; Polk, J. E.; Goebel, D. M.

    2003-01-01

    In this paper we present ion thruster design concepts created using the new computer codes that model performance limiting and erosion mechanisms. Presently, the codes model extraction grid ion optics and both discharge and neutralizer hollow cathodes.

  17. Advanced Propulsion Research Interest in Materials for Propulsion

    NASA Technical Reports Server (NTRS)

    Cole, John

    2003-01-01

    This viewgraph presentation provides an overview of material science and technology in the area of propulsion energetics. The authors note that conventional propulsion systems are near peak performance and further refinements in manufacturing, engineering design and materials will only provide incremental increases in performance. Energetic propulsion technologies could potential solve the problems of energy storage density and energy-to-thrust conversion efficiency. Topics considered include: the limits of thermal propulsion systems, the need for energetic propulsion research, emerging energetic propulsion technologies, materials research needed for advanced propulsion, and potential research opportunities.

  18. Some Thoughts About Water Analysis in Shipboard Steam Propulsion Systems for Marine Engineering Students.

    ERIC Educational Resources Information Center

    Schlenker, Richard M.; And Others

    Information is presented about the problems involved in using sea water in the steam propulsion systems of large, modern ships. Discussions supply background chemical information concerning the problems of corrosion, scale buildup, and sludge production. Suggestions are given for ways to maintain a good water treatment program to effectively deal…

  19. Some Thoughts About Water Analysis in Shipboard Steam Propulsion Systems for Marine Engineering Students.

    ERIC Educational Resources Information Center

    Schlenker, Richard M.; And Others

    Information is presented about the problems involved in using sea water in the steam propulsion systems of large, modern ships. Discussions supply background chemical information concerning the problems of corrosion, scale buildup, and sludge production. Suggestions are given for ways to maintain a good water treatment program to effectively deal…

  20. Performance and Operating Characteristics of a Turbine Engine Propulsion Simulator at Simulated Flight Conditions

    DTIC Science & Technology

    1976-05-01

    Downstream Dump TDS Supply TDV TDH ’BV Turbine Drive Venturi Drive Manifold Bleed Manifold Turbine Bleed Venturi TDL Downstream Dump Fs...Propulsion simulator instrumentation. AEOC-TR-76-76 Sym Wall Static Pressure Dynamic Static Pressure Total Pressure Total Temperature Station

  1. Space Electric Research Test in the Electric Propulsion Laboratory

    NASA Image and Video Library

    1964-06-21

    Technicians prepare the Space Electric Research Test (SERT-I) payload for a test in Tank Number 5 of the Electric Propulsion Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis researchers had been studying different methods of electric rocket propulsion since the mid-1950s. Harold Kaufman created the first successful engine, the electron bombardment ion engine, in the early 1960s. These electric engines created and accelerated small particles of propellant material to high exhaust velocities. Electric engines have a very small amount of thrust, but once lofted into orbit by workhorse chemical rockets, they are capable of small, continuous thrust for periods up to several years. The electron bombardment thruster operated at a 90-percent efficiency during testing in the Electric Propulsion Laboratory. The package was rapidly rotated in a vacuum to simulate its behavior in space. The SERT-I mission, launched from Wallops Island, Virginia, was the first flight test of Kaufman’s ion engine. SERT-I had one cesium engine and one mercury engine. The suborbital flight was only 50 minutes in duration but proved that the ion engine could operate in space. The Electric Propulsion Laboratory included two large space simulation chambers, one of which is seen here. Each uses twenty 2.6-foot diameter diffusion pumps, blowers, and roughing pumps to remove the air inside the tank to create the thin atmosphere. A helium refrigeration system simulates the cold temperatures of space.

  2. Application of electrical propulsion for an active debris removal system: a system engineering approach

    NASA Astrophysics Data System (ADS)

    Covello, Fabio

    2012-10-01

    One of the main challenge in the design of an active removal system for space debris is the high ΔV required both to approach space debris lying in different orbits and to de-orbit/re-orbit them. Indeed if the system does not target a number of objects during its lifetime the cost of the removal will be far too high to be considered as the basis of an economically viable business case. Using a classical chemical propulsion (CP) system, the ΔV is limited by the mass of propellant that the system can carry. This limitation is greatly reduced if electrical propulsion is considered. Electrical propulsion (EP) systems are indeed characterized by low propellant mass requirements, however this comes at the cost of higher electrical power and, typically, higher complexity and mass of the power supply system. Because of this, the use of EP systems has been, therefore, primarily limited to station keeping maneuvers. However in the recent past, the success of missions using EP as primary propulsion (e.g. GOCE, SMART-1, Artemis, Deep Spcae1, Hayabusa) makes this technology a suitable candidate for providing propulsion for an active debris removal system. This study case will provide the analysis of the possible application of electrical propulsion systems in such a context, presenting a number of possible mission profiles. This paper will start with the description of possible mission concepts and the assessment of the EP technology, comparing near-term propulsion options, that best fits the mission. A more detailed analysis follows with the relevant trade-off to define the characteristics of the final system and its size in terms of mass and power required. A survey of available space qualified EP systems will be performed with the selection of the best candidates to be used and/or developed for an active debris removal system. The results of a similar analysis performed for a classical CP system are then presented and the two options are compared in terms of total cost of

  3. Analysis of the Magneto-Hydrodynamic (MHD) Energy Bypass Engine for High-Speed Air-Breathing Propulsion

    NASA Technical Reports Server (NTRS)

    Riggins, David W.

    2002-01-01

    The performance of the MHD energy bypass air-breathing engine for high-speed propulsion is analyzed in this investigation. This engine is a specific type of the general class of inverse cycle engines. In this paper, the general relationship between engine performance (specific impulse and specific thrust) and the overall total pressure ratio through an engine (from inlet plane to exit plane) is first developed and illustrated. Engines with large total pressure decreases, regardless of cause or source, are seen to have exponentially decreasing performance. The ideal inverse cycle engine (of which the MHD engine is a sub-set) is then demonstrated to have a significant total pressure decrease across the engine; this total pressure decrease is cycle-driven, degrades rapidly with energy bypass ratio, and is independent of any irreversibility. The ideal MHD engine (inverse cycle engine with no irreversibility other than that inherent in the MHD work interaction processes) is next examined and is seen to have an additional large total pressure decrease due to MHD-generated irreversibility in the decelerator and the accelerator. This irreversibility mainly occurs in the deceleration process. Both inherent total pressure losses (inverse cycle and MHD irreversibility) result in a significant narrowing of the performance capability of the MHD bypass engine. The fundamental characteristics of MHD flow acceleration and flow deceleration from the standpoint of irreversibility and second-law constraints are next examined in order to clarify issues regarding flow losses and parameter selection in the MM modules. Severe constraints are seen to exist in the decelerator in terms of allowable deceleration Mach numbers and volumetric (length) required for meaningful energy bypass (work interaction). Considerable difficulties are also encountered and discussed due to thermal/work choking phenomena associated with the deceleration process. Lastly, full engine simulations utilizing inlet

  4. Center for Advanced Space Propulsion Second Annual Technical Symposium Proceedings

    NASA Technical Reports Server (NTRS)

    1990-01-01

    The proceedings for the Center for Advanced Space Propulsion Second Annual Technical Symposium are divided as follows: Chemical Propulsion, CFD; Space Propulsion; Electric Propulsion; Artificial Intelligence; Low-G Fluid Management; and Rocket Engine Materials.

  5. An overview of the VASIMR engine: High power space propulsion with RF plasma generation and heating

    NASA Astrophysics Data System (ADS)

    Díaz, F. R. Chang

    2001-10-01

    The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) is a high power, radio frequency-driven magnetoplasma rocket, capable of exhaust modulation at constant power. While the plasma is produced by a helicon discharge, the bulk of the energy is added in a separate downstream stage by ion cyclotron resonance heating (ICRH). Axial momentum is obtained by the adiabatic expansion of the plasma in a magnetic nozzle. Exhaust variation in the VASIMR is primarily achieved by the selective partitioning of the RF power to the helicon and ICRH systems, with the proper adjustment of the propellant flow. However, other complementary techniques are also being studied. Operational and performance considerations favor the light gases. The physics and engineering of this device have been under study since the late 1970s. A NASA-led, research effort, involving several terms in the United States, continues to explore the scientific and technological foundations of this concept. The research involves theory, experiment, engineering design, mission analysis, and technology development. Experimentally, high density, stable plasma discharges have been generated in Helium, Hydrogen and Deuterium, as well as mixtures of these gases. Key issues involve the optimization of the helicon discharge for high-density operation and the efficient coupling of ICRH to the plasma, prior to acceleration by the magnetic nozzle. Theoretically, the dynamics of the magnetized plasma are being studied from kinetic and fluid perspectives. Plasma acceleration by the magnetic nozzle and subsequent detachment has been demonstrated in numerical simulations. These results are presently undergoing experimental verification. A brisk technology development effort for space-qualified, compact, solid-state RF equipment, and high temperature superconducting magnets is under way in support of this project. A conceptual point design for an early space demonstrator on the International Space Station has been defined

  6. Study of LH2-fueled topping cycle engine for aircraft propulsion

    NASA Technical Reports Server (NTRS)

    Turney, G. E.; Fishbach, L. H.

    1983-01-01

    An analytical investigation was made of a topping cycle aircraft engine system which uses a cryogenic fuel. This system consists of a main turboshaft engine which is mechanically coupled (by cross-shafting) to a topping loop which augments the shaft power output of the system. The thermodynamic performance of the topping cycle engine was analyzed and compared with that of a reference (conventional-type) turboshaft engine. For the cycle operating conditions selected, the performance of the topping cycle engine in terms of brake specific fuel consumption (bsfc) was determined to be about 12 percent better than that of the reference turboshaft engine. Engine weights were estimated for both the topping cycle engine and the reference turboshaft engine. These estimates were based on a common shaft power output for each engine. Results indicate that the weight of the topping cycle engine is comparable to that of the reference turboshaft engine.

  7. Characterization of advanced electric propulsion systems

    NASA Technical Reports Server (NTRS)

    Ray, P. K.

    1982-01-01

    Characteristics of several advanced electric propulsion systems are evaluated and compared. The propulsion systems studied are mass driver, rail gun, MPD thruster, hydrogen free radical thruster and mercury electron bombardment ion engine. These are characterized by specific impulse, overall efficiency, input power, average thrust, power to average thrust ratio and average thrust to dry weight ratio. Several important physical characteristics such as dry system mass, accelerator length, bore size and current pulse requirement are also evaluated in appropriate cases. Only the ion engine can operate at a specific impulse beyond 2000 sec. Rail gun, MPD thruster and free radical thruster are currently characterized by low efficiencies. Mass drivers have the best performance characteristics in terms of overall efficiency, power to average thrust ratio and average thrust to dry weight ratio. But, they can only operate at low specific impulses due to large power requirements and are extremely long due to limitations of driving current. Mercury ion engines have the next best performance characteristics while operating at higher specific impulses. It is concluded that, overall, ion engines have somewhat better characteristics as compared to the other electric propulsion systems.

  8. An Introduction to Thermodynamic Performance Analysis of Aircraft Gas Turbine Engine Cycles Using the Numerical Propulsion System Simulation Code

    NASA Technical Reports Server (NTRS)

    Jones, Scott M.

    2007-01-01

    This document is intended as an introduction to the analysis of gas turbine engine cycles using the Numerical Propulsion System Simulation (NPSS) code. It is assumed that the analyst has a firm understanding of fluid flow, gas dynamics, thermodynamics, and turbomachinery theory. The purpose of this paper is to provide for the novice the information necessary to begin cycle analysis using NPSS. This paper and the annotated example serve as a starting point and by no means cover the entire range of information and experience necessary for engine performance simulation. NPSS syntax is presented but for a more detailed explanation of the code the user is referred to the NPSS User Guide and Reference document (ref. 1).

  9. System EMC Qualification: The Incremental Approach- The BepiColombo Power Subsystem with Ion Propulsion

    NASA Astrophysics Data System (ADS)

    Kempkens, K.

    2016-05-01

    System EMC testing generally verifies the entire spacecraft performance, operating the complete satellite in most emissive mode and measuring the power bus quality.On BepiColombo spacecraft, system tests are split into two parts, first operating the power subsystem of the transfer module together with the electric propulsion running in a special vacuum chamber, and then, running the entire spacecraft with suitable thruster simulators. The paper describes the Split System EMC test approach in detail, and presents the results of the first step, confirming the validity of the approach.

  10. Planned flight test of a mercury ion auxiliary propulsion system. I - Objectives, systems descriptions, and mission operations

    NASA Technical Reports Server (NTRS)

    Power, J. L.

    1978-01-01

    A planned flight test of an 8-cm diameter, electron-bombardment mercury ion thruster system is described. The primary objective of the test is to flight qualify the 5 mN thruster system for auxiliary propulsion applications. A seven year north-south stationkeeping mission was selected as the basis for the flight test operating profile. The flight test, which will employ two thruster systems, will also generate thruster system space performance data, measure thruster-spacecraft interactions, and demonstrate thruster operation in a number of operating modes. The flight test is designated as SAMSO-601 and will be flown aboard the Shuttle-launched Air Force Space Test Program P80-1 satellite in 1981. The spacecraft will be 3-axis stabilized in its final 740 km circular orbit, which will have an inclination of at least 73 degrees. The spacecraft design lifetime is three years.

  11. Planned flight test of a mercury ion auxiliary propulsion system. 1: Objectives, systems descriptions, and mission operations

    NASA Technical Reports Server (NTRS)

    Power, J. C.

    1978-01-01

    A planned flight test of an 8 cm diameter, electron-bombardment mercury ion thruster system is described. The primary objective of the test is to flight qualify the 5 mN (1 mlb.) thruster system for auxiliary propulsion applications. A seven year north-south stationkeeping mission was selected as the basis for the flight test operating profile. The flight test, which will employ two thruster systems, will also generate thruster system space performance data, measure thruster-spacecraft interactions, and demonstrate thruster operation in a number of operating modes. The flight test is designated as SAMSO-601 and will be flown aboard the shuttle-launched Air Force space test program P80-1 satellite in 1981. The spacecraft will be 3- axis stabilized in its final 740 km circular orbit, which will have an inclination of approximately greater than 73 degrees. The spacecraft design lifetime is three years.

  12. Organic positive ions in aircraft gas-turbine engine exhaust

    NASA Astrophysics Data System (ADS)

    Sorokin, Andrey; Arnold, Frank

    Volatile organic compounds (VOCs) represent a significant fraction of atmospheric aerosol. However the role of organic species emitted by aircraft (as a consequence of the incomplete combustion of fuel in the engine) in nucleation of new volatile particles still remains rather speculative and requires a much more detailed analysis of the underlying mechanisms. Measurements in aircraft exhaust plumes have shown the presence of both different non-methane VOCs (e.g. PartEmis project) and numerous organic cluster ions (MPIK-Heidelberg). However the link between detected organic gas-phase species and measured mass spectrum of cluster ions is uncertain. Unfortunately, up to now there are no models describing the thermodynamics of the formation of primary organic cluster ions in the exhaust of aircraft engines. The aim of this work is to present first results of such a model development. The model includes the block of thermodynamic data based on proton affinities and gas basicities of organic molecules and the block of non-equilibrium kinetics of the cluster ions evolution in the exhaust. The model predicts important features of the measured spectrum of positive ions in the exhaust behind aircraft. It is shown that positive ions emitted by aircraft engines into the atmosphere mostly consist of protonated and hydrated organic cluster ions. The developed model may be explored also in aerosol investigations of the background atmosphere as well as in the analysis of the emission of fine aerosol particles by automobiles.

  13. Justification of the Impact of the Use PPS (Plasmic Propulsion System) on Li-Ion VES140S/VES180 Batteries

    NASA Astrophysics Data System (ADS)

    Borthomieu, Yannick; Prevot, Didier

    2014-08-01

    Lithium-ion (Li-ion) battery has been since the beginning of 2000's with the support of ESA, CNES but also the European primes Astrium, (now Airbus Space and Defense) and Thalès Alénia Space. This technology replaced quickly the previous NiH2 system mainly for GEO applications thanks to the numerous advantage brought by this promising technology in terms of technical, industrial and cost aspects.The use of the Plasmic Propulsion System has been considered very early in the VES Saft Li-Ion cell development program, and included in the first life tests that run.The objective of this document is to present the impact of the use of the PPS (plasmic propulsion system also called IPS : ionic propulsion system or XPS : Xenon propulsion system) on the Saft VES140/180 Li-Ion batteries on board GEO telecommunication satellites. The PPS battery impacts have been tested since 2000 on VES140 cells and since 2006 on VES180. More than 12 years feedback on this new type of battery use on- board GEO satellites allows giving significant justification of the use of the PPS power on the battery.

  14. The supersonic fan engine: An advanced concept in supersonic cruise propulsion

    NASA Technical Reports Server (NTRS)

    Franciscus, L. C.

    1981-01-01

    Engine performance and mission studies were carried out for turbofan engines equipped with supersonic through-flow fans. The mission was for a commercial supersonic transport with a Mach 2.32 capability. The advantages of the supersonic fan engines are discussed in terms of mission range comparisons with other engine types. The effects of fan efficiency, inlet losses, and engine weight on engine performance and mission range are shown. The range of a supersonic transport with supersonic fan engines could be 10 to 20 percent better than with other types having the same technology core.

  15. Emerging Propulsion Technologies

    NASA Technical Reports Server (NTRS)

    Keys, Andrew S.

    2006-01-01

    The Emerging Propulsion Technologies (EPT) investment area is the newest area within the In-Space Propulsion Technology (ISPT) Project and strives to bridge technologies in the lower Technology Readiness Level (TRL) range (2 to 3) to the mid TRL range (4 to 6). A prioritization process, the Integrated In-Space Transportation Planning (IISTP), was developed and applied in FY01 to establish initial program priorities. The EPT investment area emerged for technologies that scored well in the IISTP but had a low technical maturity level. One particular technology, the Momentum-eXchange Electrodynamic-Reboost (MXER) tether, scored extraordinarily high and had broad applicability in the IISTP. However, its technical maturity was too low for ranking alongside technologies like the ion engine or aerocapture. Thus MXER tethers assumed top priority at EPT startup in FY03 with an aggressive schedule and adequate budget. It was originally envisioned that future technologies would enter the ISP portfolio through EPT, and EPT developed an EPT/ISP Entrance Process for future candidate ISP technologies. EPT has funded the following secondary, candidate ISP technologies at a low level: ultra-lightweight solar sails, general space/near-earth tether development, electrodynamic tether development, advanced electric propulsion, and in-space mechanism development. However, the scope of the ISPT program has focused over time to more closely match SMD needs and technology advancement successes. As a result, the funding for MXER and other EPT technologies is not currently available. Consequently, the MXER tether tasks and other EPT tasks were expected to phased out by November 2006. Presentation slides are presented which provide activity overviews for the aerocapture technology and emerging propulsion technology projects.

  16. Nuclear Thermal Propulsion engine based on Particle Bed Reactor using light water steam as a propellant

    SciTech Connect

    Powell, J.R.; Ludewig, H.; Maise, G.

    1993-06-01

    In this paper the possibility of configuring a water cooled Nuclear Thermal Propulsion (NTP) rocket, based on a Particle Bed Reactor (PBR) is investigated. This rocket will be used to operate on water obtained from near earth objects. The conclusions reached in this paper indicate that it is possible to configure a PBR based NTP rocket to operate on water and meet the mission requirements envisioned for it. No insurmountable technology issues have been identified.

  17. Results of a 2000 hour wear test of the NEXIS ion engine

    NASA Technical Reports Server (NTRS)

    Snyder, John Steven; Goebel, Dan M.; Polk, James E.; Schneider, Analyn C; Sengupta, Anita

    2005-01-01

    The Nuclear Electric Xenon Ion System (NEXIS) ion thruster was developed for potential outer planet robotic missions under NASA's Prometheus program. This engine was designed to operate at power levels ranging from 16 to over 20 kWe at specific impulses of 6000 to 7500 s for burn times of up to 10 years, satisfying the requirements of nuclear electric propulsion systems such as that on the proposed Prometheus 1 mission to explore the icy moons of Jupiter. State-of-the-art performance and life assessment tools were used to design the thruster. Following the successful performance validation of a Laboratory Model thruster, Development Model hardware was fabricated and subjected to vibration and wear testing. The results of a 2000-hour wear test are reported herein. Thruster performance achieved the target requirements and was steady for the duration of the test. Ion optics performance was similarly stable. Discharge loss increases of 6 eV/ion were observed in the first 500 hours of the test and were attributed to primary electron energy decreases due to cathode insert conditioning. Relatively high recycle rates were observed and were identified to be high-voltage-to-ground arcs in the back of the thruster caused by wire insulation outgassing and electron penetration through the plasma screen. Field emission of electrons between the accelerator and screen grids was observed and attributed to evolution of field emitter sites at accelerator grid aperture edges caused by ion bombardment. Preliminary modeling and analysis indicates that the NEXIS engine can meet mission performance requirements over the required lifetime. Finally, successful validation of the NEXIS design methodology, design tools, and technologies with the results of the wear test and companion performance and vibration tests presents significant applicability of the NEXIS development effort to missions of near-term as well as long-term interest for NASA.

  18. Results of a 2000 Hour Wear Tof the NEXIS Ion Engine

    NASA Technical Reports Server (NTRS)

    Snyder, John Steven; Goebel, Dan M.; Polk, James E.; Schneider, Analyn C; Sengupta, Anita

    2005-01-01

    The Nuclear Electric Xenon Ion System (NEXIS) ion thruster was developed for potential outer planet robotic missions under NASA's Prometheus program. This engine was designed to operate at power levels ranging from 16 to over 20 kWe at specific impulses of 6000 to 7500 s for burn times of up to 10 years, satisfying the requirements of nuclear electric propulsion systems such as that on the proposed Prometheus 1 mission to explore the icy moons of Jupiter. State-of-the-art performance and life assessment tools were used to design the thruster. Following the successful performance validation of a Laboratory Model thruster, Development Model hardware was fabricated and subjected to vibration and wear testing. The results of a 2000-hour wear test are reported herein. Thruster performance achieved the target requirements and was steady for the duration of the test. Ion optics performance was similarly stable. Discharge loss increases of 6 eV/ion were observed in the first 500 hours of the test and were attributed to primary electron energy decreases due to cathode insert conditioning. Relatively high recycle rates were observed and were identified to be high-voltage-to-ground arcs in the back of the thruster caused by wire insulation outgassing and electron penetration through the plasma screen. Field emission of electrons between the accelerator and screen grids was observed and attributed to evolution of field emitter sites at accelerator grid aperture edges caused by ion bombardment. Preliminary modeling and analysis indicates that the NEXIS engine can meet mission performance requirements over the required lifetime. Finally, successful validation of the NEXIS design methodology, design tools, and technologies with the results of the wear test and companion performance and vibration tests presents significant applicability of the NEXIS development effort to missions of near-term as well as long-term interest for NASA.

  19. Results of a 2000 Hour Wear Tof the NEXIS Ion Engine

    NASA Technical Reports Server (NTRS)

    Snyder, John Steven; Goebel, Dan M.; Polk, James E.; Schneider, Analyn C; Sengupta, Anita

    2005-01-01

    The Nuclear Electric Xenon Ion System (NEXIS) ion thruster was developed for potential outer planet robotic missions under NASA's Prometheus program. This engine was designed to operate at power levels ranging from 16 to over 20 kWe at specific impulses of 6000 to 7500 s for burn times of up to 10 years, satisfying the requirements of nuclear electric propulsion systems such as that on the proposed Prometheus 1 mission to explore the icy moons of Jupiter. State-of-the-art performance and life assessment tools were used to design the thruster. Following the successful performance validation of a Laboratory Model thruster, Development Model hardware was fabricated and subjected to vibration and wear testing. The results of a 2000-hour wear test are reported herein. Thruster performance achieved the target requirements and was steady for the duration of the test. Ion optics performance was similarly stable. Discharge loss increases of 6 eV/ion were observed in the first 500 hours of the test and were attributed to primary electron energy decreases due to cathode insert conditioning. Relatively high recycle rates were observed and were identified to be high-voltage-to-ground arcs in the back of the thruster caused by wire insulation outgassing and electron penetration through the plasma screen. Field emission of electrons between the accelerator and screen grids was observed and attributed to evolution of field emitter sites at accelerator grid aperture edges caused by ion bombardment. Preliminary modeling and analysis indicates that the NEXIS engine can meet mission performance requirements over the required lifetime. Finally, successful validation of the NEXIS design methodology, design tools, and technologies with the results of the wear test and companion performance and vibration tests presents significant applicability of the NEXIS development effort to missions of near-term as well as long-term interest for NASA.

  20. Results of a 2000 hour wear test of the NEXIS ion engine

    NASA Technical Reports Server (NTRS)

    Snyder, John Steven; Goebel, Dan M.; Polk, James E.; Schneider, Analyn C; Sengupta, Anita

    2005-01-01

    The Nuclear Electric Xenon Ion System (NEXIS) ion thruster was developed for potential outer planet robotic missions under NASA's Prometheus program. This engine was designed to operate at power levels ranging from 16 to over 20 kWe at specific impulses of 6000 to 7500 s for burn times of up to 10 years, satisfying the requirements of nuclear electric propulsion systems such as that on the proposed Prometheus 1 mission to explore the icy moons of Jupiter. State-of-the-art performance and life assessment tools were used to design the thruster. Following the successful performance validation of a Laboratory Model thruster, Development Model hardware was fabricated and subjected to vibration and wear testing. The results of a 2000-hour wear test are reported herein. Thruster performance achieved the target requirements and was steady for the duration of the test. Ion optics performance was similarly stable. Discharge loss increases of 6 eV/ion were observed in the first 500 hours of the test and were attributed to primary electron energy decreases due to cathode insert conditioning. Relatively high recycle rates were observed and were identified to be high-voltage-to-ground arcs in the back of the thruster caused by wire insulation outgassing and electron penetration through the plasma screen. Field emission of electrons between the accelerator and screen grids was observed and attributed to evolution of field emitter sites at accelerator grid aperture edges caused by ion bombardment. Preliminary modeling and analysis indicates that the NEXIS engine can meet mission performance requirements over the required lifetime. Finally, successful validation of the NEXIS design methodology, design tools, and technologies with the results of the wear test and companion performance and vibration tests presents significant applicability of the NEXIS development effort to missions of near-term as well as long-term interest for NASA.

  1. Comprehensive Modeling and Analysis of Rotorcraft Variable Speed Propulsion System With Coupled Engine/Transmission/Rotor Dynamics

    NASA Technical Reports Server (NTRS)

    DeSmidt, Hans A.; Smith, Edward C.; Bill, Robert C.; Wang, Kon-Well

    2013-01-01

    This project develops comprehensive modeling and simulation tools for analysis of variable rotor speed helicopter propulsion system dynamics. The Comprehensive Variable-Speed Rotorcraft Propulsion Modeling (CVSRPM) tool developed in this research is used to investigate coupled rotor/engine/fuel control/gearbox/shaft/clutch/flight control system dynamic interactions for several variable rotor speed mission scenarios. In this investigation, a prototypical two-speed Dual-Clutch Transmission (DCT) is proposed and designed to achieve 50 percent rotor speed variation. The comprehensive modeling tool developed in this study is utilized to analyze the two-speed shift response of both a conventional single rotor helicopter and a tiltrotor drive system. In the tiltrotor system, both a Parallel Shift Control (PSC) strategy and a Sequential Shift Control (SSC) strategy for constant and variable forward speed mission profiles are analyzed. Under the PSC strategy, selecting clutch shift-rate results in a design tradeoff between transient engine surge margins and clutch frictional power dissipation. In the case of SSC, clutch power dissipation is drastically reduced in exchange for the necessity to disengage one engine at a time which requires a multi-DCT drive system topology. In addition to comprehensive simulations, several sections are dedicated to detailed analysis of driveline subsystem components under variable speed operation. In particular an aeroelastic simulation of a stiff in-plane rotor using nonlinear quasi-steady blade element theory was conducted to investigate variable speed rotor dynamics. It was found that 2/rev and 4/rev flap and lag vibrations were significant during resonance crossings with 4/rev lagwise loads being directly transferred into drive-system torque disturbances. To capture the clutch engagement dynamics, a nonlinear stick-slip clutch torque model is developed. Also, a transient gas-turbine engine model based on first principles mean

  2. Engineered ion channels as emerging tools for chemical biology.

    PubMed

    Mayer, Michael; Yang, Jerry

    2013-12-17

    Over the last 25 years, researchers have developed exogenously expressed, genetically engineered, semi-synthetic, and entirely synthetic ion channels. These structures have sufficient fidelity to serve as unique tools that can reveal information about living organisms. One of the most exciting success stories is optogenetics: the use of light-gated channels to trigger action potentials in specific neurons combined with studies of the response from networks of cells or entire live animals. Despite this breakthrough, the use of molecularly engineered ion channels for studies of biological systems is still in its infancy. Historically, researchers studied ion channels in the context of their own function in single cells or in multicellular signaling and regulation. Only recently have researchers considered ion channels and pore-forming peptides as responsive tools to report on the chemical and physical changes produced by other biochemical processes and reactions. This emerging class of molecular probes has a number of useful characteristics. For instance, these structures can greatly amplify the signal of chemical changes: the binding of one molecule to a ligand-gated ion channel can result in flux of millions of ions across a cell membrane. In addition, gating occurs on sub-microsecond time scales, resulting in fast response times. Moreover, the signal is complementary to existing techniques because the output is ionic current rather than fluorescence or radioactivity. And finally, ion channels are also localized at the membrane of cells where essential processes such as signaling and regulation take place. This Account highlights examples, mostly from our own work, of uses of ion channels and pore-forming peptides such as gramicidin in chemical biology. We discuss various strategies for preparing synthetically tailored ion channels that range from de novo designed synthetic molecules to genetically engineered or simply exogenously expressed or reconstituted wild

  3. Carbon-carbon grid for ion engines

    NASA Technical Reports Server (NTRS)

    Garner, Charles E. (Inventor)

    1995-01-01

    A method and apparatus of manufacturing a grid member for use in an ion discharge apparatus provides a woven carbon fiber in a matrix of carbon. The carbon fibers are orientated to provide a negatibe coefficient of thermal expansion for at least a portion of the grid member's operative range of use.

  4. Carbon-carbon grid for ion engines

    NASA Technical Reports Server (NTRS)

    Garner, Charles E. (Inventor)

    1993-01-01

    A method and apparatus of manufacturing a grid member for use in an ion discharge apparatus provides a woven carbon fiber in a matrix of carbon. The carbon fibers are orientated to provide a negatibe coefficient of thermal expansion for at least a portion of the grid member's operative range of use.

  5. Development of the Engineering Test Satellite-3 (ETS-3) ion engine system

    NASA Technical Reports Server (NTRS)

    Kitamura, S.

    1984-01-01

    The ion engine system onboard the ETS-3 is discussed. The system consists of two electron bombardment type mercury ion engines with 2 mN thrust and 2,000 sec specific impulse and a power conditioner with automatic control functions. The research and development of the system, development of its EM, PM and FM, the system test and the technical achievements leading up to final launch are discussed.

  6. Retention of Sputtered Molybdenum on Ion Engine Discharge Chamber Surfaces

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Dever, Joyce A.; Power, John L.

    2001-01-01

    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.

  7. Radioisotope Electric Propulsion (REP): A Near-Term Approach to Nuclear Propulsion

    NASA Technical Reports Server (NTRS)

    Schmidt, George R.; Manzella, David H.; Kamhawi, Hani; Kremic, Tibor; Oleson, Steven R.; Dankanich, John W.; Dudzinski, Leonard A.

    2009-01-01

    Studies over the last decade have shown radioisotope-based nuclear electric propulsion to be enhancing and, in some cases, enabling for many potential robotic science missions. Also known as radioisotope electric propulsion (REP), the technology offers the performance advantages of traditional reactor-powered electric propulsion (i.e., high specific impulse propulsion at large distances from the Sun), but with much smaller, affordable spacecraft. Future use of REP requires development of radioisotope power sources with system specific powers well above that of current systems. The US Department of Energy and NASA have developed an advanced Stirling radioisotope generator (ASRG) engineering unit, which was subjected to rigorous flight qualification-level tests in 2008, and began extended lifetime testing later that year. This advancement, along with recent work on small ion thrusters and life extension technology for Hall thrusters, could enable missions using REP sometime during the next decade.

  8. Propulsion controls

    NASA Technical Reports Server (NTRS)

    Harkney, R. D.

    1980-01-01

    Increased system requirements and functional integration with the aircraft have placed an increased demand on control system capability and reliability. To provide these at an affordable cost and weight and because of the rapid advances in electronic technology, hydromechanical systems are being phased out in favor of digital electronic systems. The transition is expected to be orderly from electronic trimming of hydromechanical controls to full authority digital electronic control. Future propulsion system controls will be highly reliable full authority digital electronic with selected component and circuit redundancy to provide the required safety and reliability. Redundancy may include a complete backup control of a different technology for single engine applications. The propulsion control will be required to communicate rapidly with the various flight and fire control avionics as part of an integrated control concept.

  9. Electric Propulsion Performance from Geo-transfer to Geosynchronous Orbits

    NASA Technical Reports Server (NTRS)

    Dankanich, John W.; Carpenter, Christian B.

    2007-01-01

    For near-Earth application, solar electric propulsion advocates have focused on Low Earth Orbit (LEO) to Geosynchronous (GEO) low-thrust transfers because of the significant improvement in capability over chemical alternatives. While the performance gain attained from starting with a lower orbit is large, there are also increased transfer times and radiation exposure risk that has hindered the commercial advocacy for electric propulsion stages. An incremental step towards electric propulsion stages is the use of integrated solar electric propulsion systems (SEPS) for GTO to GEO transfer. Thorough analyses of electric propulsion systems options and performance are presented. Results are based on existing or near-term capabilities of Arcjets, Hall thrusters, and Gridded Ion engines. Parametric analyses based on "rubber" thruster and launch site metrics are also provided.

  10. Electric Propulsion Performance from Geo-transfer to Geosynchronous Orbits

    NASA Technical Reports Server (NTRS)

    Dankanich, John W.; Carpenter, Christian B.

    2007-01-01

    For near-Earth application, solar electric propulsion advocates have focused on Low Earth Orbit (LEO) to Geosynchronous (GEO) low-thrust transfers because of the significant improvement in capability over chemical alternatives. While the performance gain attained from starting with a lower orbit is large, there are also increased transfer times and radiation exposure risk that has hindered the commercial advocacy for electric propulsion stages. An incremental step towards electric propulsion stages is the use of integrated solar electric propulsion systems (SEPS) for GTO to GEO transfer. Thorough analyses of electric propulsion systems options and performance are presented. Results are based on existing or near-term capabilities of Arcjets, Hall thrusters, and Gridded Ion engines. Parametric analyses based on "rubber" thruster and launch site metrics are also provided.

  11. Electrostatic Propulsion Beam Divergence Effects on Spacecraft Surfaces. Volume 2, Addendum 1: Ion Time-of-flight Determinations of Doubly to Singly Ionized Mercury Ion Ratios from a Mercury Electron Bombardment Discharge

    NASA Technical Reports Server (NTRS)

    Sellen, J. M., Jr.; Kemp, R. F.; Hall, D. F.

    1973-01-01

    The analysis of ion exhaust beam current flow for multiply charged ion species and the application to propellant utilization for the thruster are discussed. The ion engine in use in the experiments is a twenty centimeter diameter electromagnet electron bombardment engine. The experimental technique to determine the multiply charged ion abundance ratios using ion time of flight is described. An analytical treatment of the discharge action in producing various ion species has been carried out.

  12. Diagnostic system design for the Ion Auxiliary Propulsion System (IAPS). Flight tests of two 8 cm mercury ion

    NASA Technical Reports Server (NTRS)

    Hurst, E. B.; Thomas, G. Z.

    1981-01-01

    The mechanical, thermal, electrical design and the ground test results of four types of detectors are explained. The DSS is designed to measure the thruster efflux material deposition and S/C potential relative to the local plasma in the vicinity of two 8 cm mercury ion thrusters. The DSS consists of two quartz crystal microbalance (QCM) detectors, one potential probe, nine solar cell arrays, seven ion collectors and two electronic packages.

  13. Role of Air-Breathing Pulse Detonation Engines in High Speed Propulsion

    NASA Technical Reports Server (NTRS)

    Povinelli, Louis A.; Lee, Jin-Ho; Anderberg, Michael O.

    2001-01-01

    In this paper, the effect of flight Mach number on the relative performance of pulse detonation engines and gas turbine engines is investigated. The effect of ram and mechanical compression on combustion inlet temperature and the subsequent sensible heat release is determined. Comparison of specific thrust, fuel consumption and impulse for the two engines show the relative benefits over the Mach number range.

  14. Advanced nuclear propulsion concepts

    SciTech Connect

    Howe, S.D.

    1994-12-31

    A preliminary analysis has been carried out for two potential advanced nuclear propulsion systems: a contained pulsed nuclear propulsion engine and an antiproton initiated ICF system. The results of these studies indicate that both concepts have a high potential to help enable manned planetary exploration but require substantial development.

  15. Nuclear thermal propulsion

    NASA Technical Reports Server (NTRS)

    Bennett, Gary L.

    1991-01-01

    This document is presented in viewgraph form, and the topics covered include the following: (1) the direct fission-thermal propulsion process; (2) mission applications of direct fission-thermal propulsion; (3) nuclear engines for rocket vehicles; (4) manned mars landers; and (5) particle bed reactor design.

  16. Diagnostic of plasma streams from ion thrusters for space propulsion using emissive probes

    NASA Astrophysics Data System (ADS)

    Conde, L.; Tierno, S. P.; Domenech-Garret, J. L.; Donoso, J. M.; Castillo, M. A.; Eíriz, I.; Sáez de Ocáriz, I.

    2016-10-01

    The emissive probes are employed for the determination of the local plasma potential of plasma streams produced by ion thrusters. Its operation basically relies on electron collection and emission and are less sensitive to the ion motion than collecting probes. The diagnostic using emissive probes is reviewed with emphasis in low density plasmas. Our results support the conclusion that potential structures around the probe, as virtual cathodes, would be responsible for the operation of emissive probes in low density plasmas.

  17. The QED engine spectrum - Fusion-electric propulsion for air-breathing to interstellar flight

    NASA Technical Reports Server (NTRS)

    Bussard, Robert W.; Jameson, Lorin W.

    1993-01-01

    A new inertial-electrostatic-fusion direct electric power source can be used to drive a relativistic e-beam to heat propellant. The resulting system is shown to yield specific impulse and thrust/mass ratio 2-3 orders of magnitude larger than from other advanced propulsion concepts. This QED system can be applied to aerospace vehicles from air-breathing to near-interstellar flight. Examples are given for Earth/Mars flight missions, that show transit times of 40 d with 20 percent payload in single-stage vehicles.

  18. The QED engine spectrum - Fusion-electric propulsion for air-breathing to interstellar flight

    NASA Technical Reports Server (NTRS)

    Bussard, Robert W.; Jameson, Lorin W.

    1993-01-01

    A new inertial-electrostatic-fusion direct electric power source can be used to drive a relativistic e-beam to heat propellant. The resulting system is shown to yield specific impulse and thrust/mass ratio 2-3 orders of magnitude larger than from other advanced propulsion concepts. This QED system can be applied to aerospace vehicles from air-breathing to near-interstellar flight. Examples are given for Earth/Mars flight missions, that show transit times of 40 d with 20 percent payload in single-stage vehicles.

  19. Ring cusp/hollow cathode discharge chamber performance studies. [ion propulsion

    NASA Technical Reports Server (NTRS)

    Vaughn, J. A.; Wilbur, Paul J.

    1988-01-01

    An experimental study was performed to determine the effects of hollow cathode position, anode position, and ring cusp magnetic field configuration and strength on discharge chamber performance. The results are presented in terms of comparative plasma ion energy cost, extracted ion fraction, and beam profile data. Such comparisons are used to demonstrate whether changes in performance are caused by changes in the loss rate of primary electrons to the anode or the loss rate of ions to discharge chamber walls or cathode and anode surfaces. Results show: (1) the rate of primary electron loss to the anode decreases as the anode is moved downstream of the ring cusp toward the screen grid; (2) the loss rate of ions to hollow cathode surfaces are excessive if the cathode is located upstream of a point of peak magnetic flux density at the discharge chamber centerline; and (3) the fraction of the ions produced that are lost to discharge chamber walls and ring magnet surfaces is reduced by positioning of the magnet rings so the plasma density is uniform over the grid surface, and adjusting their strength to a level where it is sufficient to prevent excessive ion losses by Bohm diffusion.

  20. Design study of RL10 derivatives. Volume 2: Engine design characteristics. [application of rocket engine to space tug propulsion

    NASA Technical Reports Server (NTRS)

    Adams, A.

    1973-01-01

    The design characteristics of the RL-10 rocket engine are discussed. The results from critical elements evaluation, baseline engine design, parametric and special study tasks are presented. Critical element evaluation established the feasibility of various engine features such as tank head idle, pumped idle, autogenous tank pressurization, and two-phase pumping. Three baseline engines, derived from the RL-10 were conceptually designed. Parametric life and performance data were generated. Special studies were conducted to establish the impact on the engine design of environment, safety, interchangeability, and maintenance.

  1. Minority University System Engineering: A Small Satellite Design Experience Held at the Jet Propulsion Laboratory During the Summer of 1996

    NASA Technical Reports Server (NTRS)

    Ordaz, Miguel Angel

    1997-01-01

    The University of Texas at El Paso (UTEP) in conjunction with the Jet Propulsion Laboratory (JPL), North Carolina A&T and California State University of Los Angeles participated during the summer of 1996 in a prototype program known as Minority University Systems Engineering (MUSE). The program consisted of a ten week internship at JPL for students and professors of the three universities. The purpose of MUSE as set forth in the MUSE program review August 5, 1996 was for the participants to gain experience in the following areas: 1) Gain experience in a multi-disciplinary project; 2) Gain experience working in a culturally diverse atmosphere; 3) Provide field experience for students to reinforce book learning; and 4) Streamline the design process in two areas: make it more financially feasible; and make it faster.

  2. In-flight Operation of the Dawn Ion Propulsion System Through Year One of Cruise to Ceres

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.; Brophy, John R.

    2013-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, 4 Vesta, and the dwarf planet 1 Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 rocket that placed the 1218-kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide an additional delta- V of approximately 11 km/s for the heliocentric transfers to each body and for all orbit transfers including orbit capture/escape and transition to the various science orbits. Deterministic thrusting to Vesta began in December 2007 and concluded with orbit capture at Vesta in July 2011. The transfer to Vesta included a Mars gravity assist flyby in February 2009 that provided an additional delta-V of 2.6 km/s and was the only postlaunch mission delta-V not provided by IPS. During the mission at Vesta the IPS was used for all orbit transfers which included six different near-polar science mapping orbits. Thrusting for departure from Vesta and the start of cruise to Ceres began on July 25, 2012 with escape from Vesta occurring on September 5, 2012. To date the IPS has been operated for approximately 31,000 hours, consumed approximately 300 kg of xenon, and provided a delta-V of approximately 8.3 km/s. IPS performance characteristics are very close to the expected performance based on analysis and testing performed prelaunch. Thrusting for cruise to Ceres will continue until the spring of 2015, with a planned arrival date at Ceres in April 2015. This paper provides an overview of Dawn's mission objectives and the results of Dawn IPS mission operations from Vesta departure through the first year of cruise to Ceres.

  3. NASA Propulsion Investments for Exploration and Science

    NASA Technical Reports Server (NTRS)

    Smith, Bryan K.; Free, James M.; Klem, Mark D.; Priskos, Alex S.; Kynard, Michael H.

    2008-01-01

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

  4. Electric propulsion technology

    NASA Technical Reports Server (NTRS)

    Finke, R. C.

    1980-01-01

    The advanced electric propulsion program is directed towards lowering the specific impulse and increasing the thrust per unit of ion thruster systems. In addition, electrothermal and electromagnetic propulsion technologies are being developed to attempt to fill the gap between the conventional ion thruster and chemical rocket systems. Most of these new concepts are exagenous and are represented by rail accelerators, ablative Teflon thrusters, MPD arcs, Free Radicals, etc. Endogenous systems such as metallic hydrogen offer great promise and are also being pursued.

  5. Test stand performance of a convertible engine for advanced V/STOL and rotorcraft propulsion

    NASA Technical Reports Server (NTRS)

    Mcardle, Jack G.

    1987-01-01

    A variable inlet guide vane (VIGV) convertible engine that could be used to power future high-speed V/STOL and rotorcraft was tested on an outdoor stand. The engine ran stably and smoothly in the turbofan, turboshaft, and dual (combined fan and shaft) power modes. In the turbofan mode with the VIGV open, fuel consumption was comparable to that of a conventional turbofan engine. In the turboshaft mode with the VIGV closed, fuel consumption was higher than that of present turboshaft engines because power was wasted in churning fan-tip air flow. In dynamic performance tests with a specially built digital engine control and using a waterbrake dynamometer for shaft load, the engine responded effectively to large steps in thrust command and shaft torque.

  6. Advanced rocket propulsion

    NASA Technical Reports Server (NTRS)

    Obrien, Charles J.

    1993-01-01

    Existing NASA research contracts are supporting development of advanced reinforced polymer and metal matrix composites for use in liquid rocket engines of the future. Advanced rocket propulsion concepts, such as modular platelet engines, dual-fuel dual-expander engines, and variable mixture ratio engines, require advanced materials and structures to reduce overall vehicle weight as well as address specific propulsion system problems related to elevated operating temperatures, new engine components, and unique operating processes. High performance propulsion systems with improved manufacturability and maintainability are needed for single stage to orbit vehicles and other high performance mission applications. One way to satisfy these needs is to develop a small engine which can be clustered in modules to provide required levels of total thrust. This approach should reduce development schedule and cost requirements by lowering hardware lead times and permitting the use of existing test facilities. Modular engines should also reduce operational costs associated with maintenance and parts inventories.

  7. NACA Conference on Turbojet Engines for Supersonic Propulsion. A Compilation of Technical Material Presented

    DTIC Science & Technology

    1953-10-01

    turbojet Pngine with a turbine cooled by compressor air involves several design pruilems that do not e~ist in an uncooled turbo - jet engine . Careful...facilitate testing the sheet-metal blades in the turbojet engine , bases were formed by removing the solid airfoil portion from the standard turbine blade ...OF TURBINE BLADES by J. C. Freche 6. APPLICATION AND OPERATION OF AIR-COOLED TURBINES IN TURBOJET ENGINES

  8. The NASA Electric Propulsion program

    NASA Technical Reports Server (NTRS)

    Byers, D. C.

    1984-01-01

    It is pointed out that the NASA Electric Propulsion program is aimed at providing technology for auxiliary and primary propulsion functions for earth-orbital and planetary space missions. Efforts in electrostatic propulsion include analyses of ion propulsion for Geosynchronous (GEO) and planetary spacecraft, continued preflight efforts associated with the Ion Auxiliary Propulsion System (IAPS), and research and technology for advanced and simplified ion thruster systems. In the area of electromagnetic propulsion, studies were conducted regarding the feasibility and impacts of the use of electromagnetic launchers. Research on magnetoplasmadynamic (MPD) thrusters, electromagnetic launchers, and Hall current thrusters was also performed. Studies in the electrothermal sector included an evaluation of electric propulsion options for the Space Station, taking into account also resistojets, a pulsed electrothermal thruster, and arc jets.

  9. Nuclear Thermal Propulsion (NTP)

    NASA Image and Video Library

    NASA's history with nuclear thermal propulsion (NTP) technology goes back to the earliest days of the Agency. The Manned Lunar Rover Vehicle and the Nuclear Engine for Rocket Vehicle Applications p...

  10. Ion engine propelled Earth-Mars cycler with nuclear thermal propelled transfer vehicle, volume 2

    NASA Technical Reports Server (NTRS)

    Meyer, Rudolf X.; Baker, Myles; Melko, Joseph

    1994-01-01

    The goal of this project was to perform a preliminary design of a long term, reusable transportation system between earth and Mars which would be capable of providing both artificial gravity and shelter from solar flare radiation. The heart of this system was assumed to be a Cycler spacecraft propelled by an ion propulsion system. The crew transfer vehicle was designed to be propelled by a nuclear-thermal propulsion system. Several Mars transportation system architectures and their associated space vehicles were designed.

  11. Application of SDI technology in space propulsion

    SciTech Connect

    Klein, A.J.

    1992-01-01

    Numerous technologies developed by the DOD within the SDI program are now available for adaptation to the requirements of commercial spacecraft; SDI has accordingly organized the Technology Applications Information System data base, which contains nearly 2000 nonproprietary abstracts on SDI technology. Attention is here given to such illustrative systems as hydrogen arcjets, ammonia arcjets, ion engines, SSTO launch vehicles, gel propellants, lateral thrusters, pulsed electrothermal thrusters, laser-powered rockets, and nuclear propulsion.

  12. Effect of Surface Impulsive Thermal Loads on Fatigue Behavior of Constant Volume Propulsion Engine Combustor Materials

    NASA Technical Reports Server (NTRS)

    Zhu, Dongming; Fox, Dennis S.; Miller, Robert A.; Ghosn, Louis J.; Kalluri, Sreeramesh

    2004-01-01

    The development of advanced high performance constant-volume-combustion-cycle engines (CVCCE) requires robust design of the engine components that are capable of enduring harsh combustion environments under high frequency thermal and mechanical fatigue conditions. In this study, a simulated engine test rig has been established to evaluate thermal fatigue behavior of a candidate engine combustor material, Haynes 188, under superimposed CO2 laser surface impulsive thermal loads (30 to 100 Hz) in conjunction with the mechanical fatigue loads (10 Hz). The mechanical high cycle fatigue (HCF) testing of some laser pre-exposed specimens has also been conducted under a frequency of 100 Hz to determine the laser surface damage effect. The test results have indicated that material surface oxidation and creep-enhanced fatigue is an important mechanism for the surface crack initiation and propagation under the simulated CVCCE engine conditions.

  13. NASA Glenn Research Center, Propulsion Systems Laboratory: Plan to Measure Engine Core Flow Water Vapor Content

    NASA Technical Reports Server (NTRS)

    Oliver, Michael

    2014-01-01

    This presentation will be made at the 92nd AIAA Turbine Engine Testing Working Group (TETWoG), a semi-annual technical meeting of turbine engine testing professionals. The objective is to describe an effort by NASA to measure the water vapor content on the core airflow in a full scale turbine engine ice crystal icing test and to open a discussion with colleagues how to accurately conduct the measurement based on any previous collective experience with the procedure, instruments and nature of engine icing testing within the group. The presentation lays out the schematics of the location in the flow path from which the sample will be drawn, the plumbing to get it from the engine flow path to the sensor and several different water vapor measurement technologies that will be used: Tunable diode laser and infrared spectroscopy.

  14. Alternate Propulsion Energy Sources.

    DTIC Science & Technology

    1983-06-01

    Marietta - cryogenic oscillator Dr. Brice Cassenti, UTRC - antimatter propulsion A. E. Mensing, UTRC - nuclear lightbulb engine Michael Fowler, UTRC...sails, laser propulsion, tethers, fusion rockets, antimatter rockets Z9 BSTRACT (Continue on reverse aide if necessary and identify by block number) This...thorough literature search and carry out an intense technical assessment of the latest concepts in science and engineering that show promise of leading to a

  15. Evaluation of High-Power Solar Electric Propulsion using Advanced Ion, Hall, MPD, and PIT Thrusters for Lunar and Mars Cargo Missions

    NASA Technical Reports Server (NTRS)

    Frisbee, Robert H.

    2006-01-01

    This paper presents the results of mission analyses that expose the advantages and disadvantages of high-power (MWe-class) Solar Electric Propulsion (SEP) for Lunar and Mars Cargo missions that would support human exploration of the Moon and Mars. In these analyses, we consider SEP systems using advanced Ion thrusters (the Xenon [Xe] propellant Herakles), Hall thrusters (the Bismuth [Bi] propellant Very High Isp Thruster with Anode Layer [VHITAL], magnetoplasmadynamic (MPD) thrusters (the Lithium [Li] propellant Advanced Lithium-Fed, Applied-field Lorentz Force Accelerator (ALFA2), and pulsed inductive thruster (PIT) (the Ammonia [NH3] propellant Nuclear-PIT [NuPIT]). The analyses include comparison of the advanced-technology propulsion systems (VHITAL, ALFA2, and NuPIT) relative to state-of-theart Ion (Herakles) propulsion systems and quantify the unique benefits of the various technology options such as high power-per-thruster (and/or high power-per-thruster packaging volume), high specific impulse (Isp), high-efficiency, and tankage mass (e.g., low tankage mass due to the high density of bismuth propellant). This work is based on similar analyses for Nuclear Electric Propulsion (NEP) systems.

  16. Evaluation of High-Power Solar Electric Propulsion using Advanced Ion, Hall, MPD, and PIT Thrusters for Lunar and Mars Cargo Missions

    NASA Technical Reports Server (NTRS)

    Frisbee, Robert H.

    2006-01-01

    This paper presents the results of mission analyses that expose the advantages and disadvantages of high-power (MWe-class) Solar Electric Propulsion (SEP) for Lunar and Mars Cargo missions that would support human exploration of the Moon and Mars. In these analyses, we consider SEP systems using advanced Ion thrusters (the Xenon [Xe] propellant Herakles), Hall thrusters (the Bismuth [Bi] propellant Very High Isp Thruster with Anode Layer [VHITAL], magnetoplasmadynamic (MPD) thrusters (the Lithium [Li] propellant Advanced Lithium-Fed, Applied-field Lorentz Force Accelerator (ALFA2), and pulsed inductive thruster (PIT) (the Ammonia [NH3] propellant Nuclear-PIT [NuPIT]). The analyses include comparison of the advanced-technology propulsion systems (VHITAL, ALFA2, and NuPIT) relative to state-of-theart Ion (Herakles) propulsion systems and quantify the unique benefits of the various technology options such as high power-per-thruster (and/or high power-per-thruster packaging volume), high specific impulse (Isp), high-efficiency, and tankage mass (e.g., low tankage mass due to the high density of bismuth propellant). This work is based on similar analyses for Nuclear Electric Propulsion (NEP) systems.

  17. YIP - Ultrasensitive Infrared Spectroscopy of Molecular Ions of Importance in Atmospheric Chemistry and Propulsion

    DTIC Science & Technology

    2010-08-23

    State University, the University of Wisconsin at Madison, the Massachusetts Institute of Technology, the Harvard-Smithsonian Center for Astrophysics ...awarded the 2009 Coblentz Award at the 64th International Symposium on Molecular Spectroscopy. Prior to this effort, I was selected for a Presidential Early Career Award for Scientists and Engineers (PECASE) by the White House.

  18. Engineering Heteromaterials to Control Lithium Ion Transport Pathways

    SciTech Connect

    Liu, Yang; Vishniakou, Siarhei; Yoo, Jinkyoung; Dayeh, Shadi A.

    2015-12-21

    Safe and efficient operation of lithium ion batteries requires precisely directed flow of lithium ions and electrons to control the first directional volume changes in anode and cathode materials. Understanding and controlling the lithium ion transport in battery electrodes becomes crucial to the design of high performance and durable batteries. Recent work revealed that the chemical potential barriers encountered at the surfaces of heteromaterials play an important role in directing lithium ion transport at nanoscale. Here, we utilize in situ transmission electron microscopy to demonstrate that we can switch lithiation pathways from radial to axial to grain-by-grain lithiation through the systematic creation of heteromaterial combinations in the Si-Ge nanowire system. Lastly, our systematic studies show that engineered materials at nanoscale can overcome the intrinsic orientation-dependent lithiation, and open new pathways to aid in the development of compact, safe, and efficient batteries.

  19. Engineering Heteromaterials to Control Lithium Ion Transport Pathways

    DOE PAGES

    Liu, Yang; Vishniakou, Siarhei; Yoo, Jinkyoung; ...

    2015-12-21

    Safe and efficient operation of lithium ion batteries requires precisely directed flow of lithium ions and electrons to control the first directional volume changes in anode and cathode materials. Understanding and controlling the lithium ion transport in battery electrodes becomes crucial to the design of high performance and durable batteries. Recent work revealed that the chemical potential barriers encountered at the surfaces of heteromaterials play an important role in directing lithium ion transport at nanoscale. Here, we utilize in situ transmission electron microscopy to demonstrate that we can switch lithiation pathways from radial to axial to grain-by-grain lithiation through themore » systematic creation of heteromaterial combinations in the Si-Ge nanowire system. Lastly, our systematic studies show that engineered materials at nanoscale can overcome the intrinsic orientation-dependent lithiation, and open new pathways to aid in the development of compact, safe, and efficient batteries.« less

  20. Engineering Heteromaterials to Control Lithium Ion Transport Pathways

    SciTech Connect

    Liu, Yang; Vishniakou, Siarhei; Yoo, Jinkyoung; Dayeh, Shadi A.

    2015-12-21

    Safe and efficient operation of lithium ion batteries requires precisely directed flow of lithium ions and electrons to control the first directional volume changes in anode and cathode materials. Understanding and controlling the lithium ion transport in battery electrodes becomes crucial to the design of high performance and durable batteries. Some recent work revealed that the chemical potential barriers encountered at the surfaces of heteromaterials play an important role in directing lithium ion transport at nanoscale. We utilize in situ transmission electron microscopy to demonstrate that we can switch lithiation pathways from radial to axial to grain-by-grain lithiation through the systematic creation of heteromaterial combinations in the Si-Ge nanowire system. Furthermore, our systematic studies show that engineered materials at nanoscale can overcome the intrinsic orientation-dependent lithiation, and open new pathways to aid in the development of compact, safe, and efficient batteries.

  1. Engineering Heteromaterials to Control Lithium Ion Transport Pathways

    DOE PAGES

    Liu, Yang; Vishniakou, Siarhei; Yoo, Jinkyoung; ...

    2015-12-21

    Safe and efficient operation of lithium ion batteries requires precisely directed flow of lithium ions and electrons to control the first directional volume changes in anode and cathode materials. Understanding and controlling the lithium ion transport in battery electrodes becomes crucial to the design of high performance and durable batteries. Some recent work revealed that the chemical potential barriers encountered at the surfaces of heteromaterials play an important role in directing lithium ion transport at nanoscale. We utilize in situ transmission electron microscopy to demonstrate that we can switch lithiation pathways from radial to axial to grain-by-grain lithiation through themore » systematic creation of heteromaterial combinations in the Si-Ge nanowire system. Furthermore, our systematic studies show that engineered materials at nanoscale can overcome the intrinsic orientation-dependent lithiation, and open new pathways to aid in the development of compact, safe, and efficient batteries.« less

  2. Engineering Heteromaterials to Control Lithium Ion Transport Pathways

    PubMed Central

    Liu, Yang; Vishniakou, Siarhei; Yoo, Jinkyoung; Dayeh, Shadi A.

    2015-01-01

    Safe and efficient operation of lithium ion batteries requires precisely directed flow of lithium ions and electrons to control the first directional volume changes in anode and cathode materials. Understanding and controlling the lithium ion transport in battery electrodes becomes crucial to the design of high performance and durable batteries. Recent work revealed that the chemical potential barriers encountered at the surfaces of heteromaterials play an important role in directing lithium ion transport at nanoscale. Here, we utilize in situ transmission electron microscopy to demonstrate that we can switch lithiation pathways from radial to axial to grain-by-grain lithiation through the systematic creation of heteromaterial combinations in the Si-Ge nanowire system. Our systematic studies show that engineered materials at nanoscale can overcome the intrinsic orientation-dependent lithiation, and open new pathways to aid in the development of compact, safe, and efficient batteries. PMID:26686655

  3. Ion engineering of embedded nanostructures: From spherical to facetted nanoparticles

    SciTech Connect

    Rizza, G.; Dawi, E. A.; Vredenberg, A. M.; Monnet, I.

    2009-07-27

    We show that the high-energy ion irradiation of embedded metallic spherical nanoparticles (NPs) is not limited to their transformation into prolate nanorods or nanowires. Depending on their pristine size, the three following morphologies can be obtained: (i) nanorods, (ii) facettedlike, and (iii) almost spherical nanostructures. Planar silica films containing nearly monodisperse gold NPs (8-100 nm) were irradiated with swift heavy ions (5 GeV Pb) at room temperature for fluences up to 5x10{sup 13} cm{sup -2}. The experimental results are accounted for by considering a liquid-solid transformation of the premelted NP surface driven by the in-plane stress within the ion-deformed host matrix. This work demonstrates the interest of using ion-engineering techniques to shape embedded nanostructures into nonconventional configurations.

  4. Interstellar precursor missions using advanced dual-stage ion propulsion systems

    NASA Astrophysics Data System (ADS)

    Fearn, Dave G.; Walker, Roger

    2006-08-01

    In this paper it is shown that an advanced form of gridded ion thruster, employing a novel 4-grid ion extraction and acceleration system rather than the usual two or three grid variants, can provide a velocity increment and specific impulse of interest to interstellar precursor missions, extending to a few hundred astronomical units from the sun. In this it is assumed that a nuclear power source is available with a mass-to-power ratio of 15 to 35 kg/kW and an output of at least several tens of kW. Mission durations are of about 25 years and the velocity increment provided exceeds 37 km/s. The paper includes a description of the technical approach adopted to achieving the required values of specific impulse, thrust density and power consumption, and presents for the first time the data obtained from an experimental programme conducted at ESTEC to verify the principles on which these theoretical predictions are based.

  5. Rocket Propulsion Through Multiply-Charged Ions From a Mirror Plasma

    SciTech Connect

    Leung, L.; Petty, C. C.

    2007-09-28

    This paper evaluates a new type of ambipolar plasma thruster that uses multiply-charged ions as propellant. Beginning with an electron cyclotron heated (ECH) mirror plasma with energetic electrons, the ion charge state distribution and confining electrostatic potentials are self-consistently modeled for different fill gases using the particle, charge, and energy conservation equations and Pastukhov-flow confinement. The specific impulse is found to be high ({approx}6000 s) and easily varied. Although the thrust efficiency is low, 24% for double-ended operation and 45% for single-ended operation, this ambipolar thruster is capable of producing high thrust in a compact source because ECH mirror plasmas can operate at high density.

  6. Ion Engine Plume Interaction Calculations for Prototypical Prometheus 1

    NASA Technical Reports Server (NTRS)

    Mandell, Myron J.; Kuharski, Robert A.; Gardner, Barbara M.; Katz, Ira; Randolph, Tom; Dougherty, Ryan; Ferguson, Dale C.

    2005-01-01

    Prometheus 1 is a conceptual mission to demonstrate the use of atomic energy for distant space missions. The hypothetical spacecraft design considered in this paper calls for multiple ion thrusters, each with considerably higher beam energy and beam current than have previously flown in space. The engineering challenges posed by such powerful thrusters relate not only to the thrusters themselves, but also to designing the spacecraft to avoid potentially deleterious effects of the thruster plumes. Accommodation of these thrusters requires good prediction of the highest angle portions of the main beam, as well as knowledge of clastically scattered and charge exchange ions, predictions for grid erosion and contamination of surfaces by eroded grid material, and effects of the plasma plume on radio transmissions. Nonlinear interactions of multiple thrusters are also of concern. In this paper we describe two- and three-dimensional calculations for plume structure and effects of conceptual Prometheus 1 ion engines. Many of the techniques used have been validated by application to ground test data for the NSTAR and NEXT ion engines. Predictions for plume structure and possible sputtering and contamination effects will be presented.

  7. An Engineered Palette of Metal Ion Quenchable Fluorescent Proteins

    PubMed Central

    Yu, Xiaozhen; Strub, Marie-Paule; Barnard, Travis J.; Noinaj, Nicholas; Piszczek, Grzegorz; Buchanan, Susan K.; Taraska, Justin W.

    2014-01-01

    Many fluorescent proteins have been created to act as genetically encoded biosensors. With these sensors, changes in fluorescence report on chemical states in living cells. Transition metal ions such as copper, nickel, and zinc are crucial in many physiological and pathophysiological pathways. Here, we engineered a spectral series of optimized transition metal ion-binding fluorescent proteins that respond to metals with large changes in fluorescence intensity. These proteins can act as metal biosensors or imaging probes whose fluorescence can be tuned by metals. Each protein is uniquely modulated by four different metals (Cu2+, Ni2+, Co2+, and Zn2+). Crystallography revealed the geometry and location of metal binding to the engineered sites. When attached to the extracellular terminal of a membrane protein VAMP2, dimeric pairs of the sensors could be used in cells as ratiometric probes for transition metal ions. Thus, these engineered fluorescent proteins act as sensitive transition metal ion-responsive genetically encoded probes that span the visible spectrum. PMID:24752441

  8. The QED engine - Fusion-electric propulsion for Cis-Oort/Quasi-Interstellar (QIS) flight

    NASA Technical Reports Server (NTRS)

    Bussard, Robert W.; Jameson, Lorin W.; Froning, H. D., Jr.

    1993-01-01

    A summary is presented of QED fusion-direct-electric engine systems, their features, and performance ranges. The principles and characteristics of inertial-electrostatic-fusion (IEF) power source systems are then reviewed, and their application to the diluted-fusion-product (DFP) engine concept for QIS missions is discussed. Particular attention is given to vehicle performance over a range of very high specific impulses and to specifications of a typical candidate DFP/IEF engine and a single-stage vehicle for rapid flight to 550 AU.

  9. The QED engine - Fusion-electric propulsion for Cis-Oort/Quasi-Interstellar (QIS) flight

    NASA Technical Reports Server (NTRS)

    Bussard, Robert W.; Jameson, Lorin W.; Froning, H. D., Jr.

    1993-01-01

    A summary is presented of QED fusion-direct-electric engine systems, their features, and performance ranges. The principles and characteristics of inertial-electrostatic-fusion (IEF) power source systems are then reviewed, and their application to the diluted-fusion-product (DFP) engine concept for QIS missions is discussed. Particular attention is given to vehicle performance over a range of very high specific impulses and to specifications of a typical candidate DFP/IEF engine and a single-stage vehicle for rapid flight to 550 AU.

  10. Outer-Planet Mission Analysis Using Solar-Electric Ion Propulsion

    NASA Technical Reports Server (NTRS)

    Woo, Byoungsam; Coverstone, Victoria L.; Hartmann, John W.; Cupples, Michael

    2003-01-01

    Outer-planet mission analysis was performed using three next generation solar-electric ion thruster models. Optimal trajectories are presented that maximize the delivered mass to the designated outer planet. Trajectories to Saturn and Neptune with a single Venus gravity assist are investigated. For each thruster model, the delivered mass versus flight time curve was generated to obtain thruster model performance. The effects of power to the thrusters and resonance ratio of Venutian orbital periods to spacecraft period were also studied. Multiple locally optimal trajectories to Saturn and Neptune have been discovered in different regions of the parameter search space. The characteristics of each trajectory are noted.

  11. 13-kV Ion-Extraction System Being Developed for Inert Gas Ion Engines

    NASA Technical Reports Server (NTRS)

    Rawlin, Vincent K.; Williams, George J.; Wilbur, Paul

    2002-01-01

    A high-voltage ion optics design was chosen for an assumed outer planet or interstellar precursor mission that would require a long-life, high-power, high-specific-impulse krypton ion engine. Such an engine could support energetic space missions to the outer planets or beyond. Detailed performance and lifetime analyses and several inexpensive subscale grid tests were conducted at the NASA Glenn Research Center and at the Colorado State University under a NASA Glenn grant. A subscale grid set of the selected geometry shown was tested at voltages up to 13,000 V. This yielded a krypton ion beam current that would, when scaled to a full-size 50-cm diameter, produce an ion beam with a power of 30 kW at a specific impulse over 14,000 sec. The operational ion beam focusing limits, as a function of ion current per hole, were found to impose requirements of high uniformity on the discharge chamber plasma density. A full-size set of two-grid, 50-cm-diameter titanium ion optics has been fabricated and awaits testing.

  12. Emergent Propulsion Systems

    NASA Astrophysics Data System (ADS)

    El-Fakdi Sencianes, Andres

    2002-01-01

    almost an Engineer (2002 will be my last year as student) and the studies that I'm now ending here, in Girona, are closely related not only with science and technology subjects but with optimization and economic result obtention, too. Huge distances that separate us from everything in space have launched scientists and engineers into a new challenge: How to reach maximum speeds keeping high ratios payload/total spacecraft mass? The key limitation of chemical rockets is that their exhaust velocity is relatively low. Because achieving Earth orbit requires a high velocity change a rocket must carry far more propellant than payload. The answer to all this complications seems to stare in one way: electric propulsion systems and the possibility of taking advantatge of solar winds to thrust our crafts. possible solutions, some of them have been studied for years and now they are not a project but a reality; also newest theories bring us the possibility of dream. Improve of commom propellants, search of new ones: Investigators continued research on use of atomic species as high-energy-density propellants, which could increase the specific impulse of hydrogen/oxygen rockets by 50-150%. Nuclear fission propulsion: Centered in development of reactors for nearterm nuclear electric propulsion aplications. Multimegawatt systems based on vapor core reactors and magnetohydrodynamic power conversion. Engineers investigated new fuels for compact nuclear thermal propulsion systems. What is called plasma state?: When a gas is heated to tens of thousands or millions of degrees, atoms lose their electrons. The result is a "soup" of charged particles, or plasma, made up of negatively charged electrons and positively charged ions. No known material can contain the hot plasma necessary for rocket propulsion, but specially designed magnetic fields can. Plasma rockets: This rockets are not powered by conventional chemical reactions as today's rockets are, but by electrical energy that heats

  13. Simulated Altitude Performance of Combustor of Westinghouse 19XB-1 Jet-Propulsion Engine

    NASA Technical Reports Server (NTRS)

    Childs, J. Howard; McCafferty, Richard J.

    1948-01-01

    A 19XB-1 combustor was operated under conditions simulating zero-ram operation of the 19XB-1 turbojet engine at various altitudes and engine speeds. The combustion efficiencies and the altitude operational limits were determined; data were also obtained on the character of the combustion, the pressure drop through the combustor, and the combustor-outlet temperature and velocity profiles. At altitudes about 10,000 feet below the operational limits, the flames were yellow and steady and the temperature rise through the combustor increased with fuel-air ratio throughout the range of fuel-air ratios investigated. At altitudes near the operational limits, the flames were blue and flickering and the combustor was sluggish in its response to changes in fuel flow. At these high altitudes, the temperature rise through the combustor increased very slowly as the fuel flow was increased and attained a maximum at a fuel-air ratio much leaner than the over-all stoichiometric; further increases in fuel flow resulted in decreased values of combustor temperature rise and increased resonance until a rich-limit blow-out occurred. The approximate operational ceiling of the engine as determined by the combustor, using AN-F-28, Amendment-3, fuel, was 30,400 feet at a simulated engine speed of 7500 rpm and increased as the engine speed was increased. At an engine speed of 16,000 rpm, the operational ceiling was approximately 48,000 feet. Throughout the range of simulated altitudes and engine speeds investigated, the combustion efficiency increased with increasing engine speed and with decreasing altitude. The combustion efficiency varied from over 99 percent at operating conditions simulating high engine speed and low altitude operation to less than 50 percent at conditions simulating operation at altitudes near the operational limits. The isothermal total pressure drop through the combustor was 1.82 times as great as the inlet dynamic pressure. As expected from theoretical

  14. Propulsion and Power Supplies for Unmanned Vehicles. Volume I. Engines for Small Propeller-Driven RPVS

    DTIC Science & Technology

    1977-11-01

    speed. Based on data of e sting small RPV’s for the smallest RPV (typo "A") a good choice for the wing loading seems: W/S - 25 kg/m . It is considered as...CC 24 Saladillo Provincia de Buenos Aires Argentine Clinton Engine Corp. MaquokeLa Iowa, 52060, USA Cuyuna Engine - Scorpion Inc. Crosby, Minnesota...Manufacturer: Scorpion Inc. Crosby, Minnesotta Type designation: Cuyuna Rated b.h.p. @ r.p.m.: 40 bhp @ 6,500 rpm Number of cylinders and layout: 2

  15. Separation of organic ion exchange resins from sludge -- engineering study

    SciTech Connect

    Duncan, J.B.

    1998-08-25

    This engineering study evaluates the use of physical separation technologies to separate organic ion exchange resin from KE Basin sludge prior to nitric acid dissolution. This separation is necessitate to prevent nitration of the organics in the acid dissolver. The technologies under consideration are: screening, sedimentation, elutriation. The recommended approach is to first screen the Sludge and resin 300 microns then subject the 300 microns plus material to elutriation.

  16. In-Flight Operation of the Dawn Ion Propulsion System Through the Preparations for Escape From Vesta

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.; Brophy, John R.; Mikes, Steven C.

    2012-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, 4 Vesta, and the dwarf planet 1 Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 rocket that placed the 1218-kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total delta-V of approximately 11 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, transfer between Ceres science orbits, and orbit maintenance maneuvers for all Vesta and Ceres science orbits. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional delta-V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. An additional 231 hours of IPS thrusting was used to enter the first Vesta science orbit, called Survey orbit, on August 3, 2011 at an altitude of about 2,735 km. The IPS was then used over the next year to transfer the spacecraft to the other science orbits: a high altitude mapping orbit (HAMO-1) in September 2011 at an altitude of approximately 673 km, a low altitude mapping orbit (LAMO) at approximately 210 km altitude, and a second high altitude mapping orbit (HAMO-2) at approximately 673 km altitude. To date the IPS has been operated for approximately 24,327 hours, consumed approximately 260 kg of xenon, and provided a delta-V of approximately 7 km/s. IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. Thrusting for escape from Vesta and

  17. In-Flight Operation of the Dawn Ion Propulsion System Through the Preparations for Escape From Vesta

    NASA Technical Reports Server (NTRS)

    Garner, Charles E.; Rayman, Marc D.; Brophy, John R.; Mikes, Steven C.

    2012-01-01

    The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, 4 Vesta, and the dwarf planet 1 Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 rocket that placed the 1218-kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total delta-V of approximately 11 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, transfer between Ceres science orbits, and orbit maintenance maneuvers for all Vesta and Ceres science orbits. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional delta-V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. An additional 231 hours of IPS thrusting was used to enter the first Vesta science orbit, called Survey orbit, on August 3, 2011 at an altitude of about 2,735 km. The IPS was then used over the next year to transfer the spacecraft to the other science orbits: a high altitude mapping orbit (HAMO-1) in September 2011 at an altitude of approximately 673 km, a low altitude mapping orbit (LAMO) at approximately 210 km altitude, and a second high altitude mapping orbit (HAMO-2) at approximately 673 km altitude. To date the IPS has been operated for approximately 24,327 hours, consumed approximately 260 kg of xenon, and provided a delta-V of approximately 7 km/s. IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. Thrusting for escape from Vesta and

  18. Propulsion health monitoring of a turbine engine disk using spin test data

    NASA Astrophysics Data System (ADS)

    Abdul-Aziz, Ali; Woike, Mark; Oza, Nikunj; Matthews, Bryan; Baakilini, George

    2010-03-01

    On line detection techniques to monitor the health of rotating engine components are becoming increasingly attractive options to aircraft engine companies in order to increase safety of operation and lower maintenance costs. Health monitoring remains a challenging feature to easily implement, especially, in the presence of scattered loading conditions, crack size, component geometry and materials properties. The current trend, however, is to utilize noninvasive types of health monitoring or nondestructive techniques to detect hidden flaws and mini cracks before any catastrophic event occurs. These techniques go further to evaluate materials' discontinuities and other anomalies that have grown to the level of critical defects which can lead to failure. Generally, health monitoring is highly dependent on sensor systems that are capable of performing in various engine environmental conditions and able to transmit a signal upon a predetermined crack length, while acting in a neutral form upon the overall performance of the engine system. Efforts are under way at NASA Glenn Research Center through support of the Intelligent Vehicle Health Management Project (IVHM) to develop and implement such sensor technology for a wide variety of applications. These efforts are focused on developing high temperature, wireless, low cost and durable products. Therefore, in an effort to address the technical issues concerning health monitoring of a rotor disk, this paper considers data collected from an experimental study using high frequency capacitive sensor technology to capture blade tip clearance and tip timing measurements in a rotating engine-like-disk-to predict the disk faults and assess its structural integrity. The experimental results collected at a range of rotational speeds from tests conducted at the NASA Glenn Research Center's Rotordynamics Laboratory will be evaluated using multiple data-driven anomaly detection techniques to identify anomalies in the disk. This study

  19. Surface Engineering of Nanostructured Titanium Implants with Bioactive Ions.

    PubMed

    Kim, H-S; Kim, Y-J; Jang, J-H; Park, J-W

    2016-05-01

    Surface nanofeatures and bioactive ion chemical modification are centrally important in current titanium (Ti) oral implants for enhancing osseointegration. However, it is unclear whether the addition of bioactive ions definitively enhances the osteogenic capacity of a nanostructured Ti implant. We systematically investigated the osteogenesis process of human multipotent adipose stem cells triggered by bioactive ions in the nanostructured Ti implant surface. Here, we report that bioactive ion surface modification (calcium [Ca] or strontium [Sr]) and resultant ion release significantly increase osteogenic activity of the nanofeatured Ti surface. We for the first time demonstrate that ion modification actively induces focal adhesion development and expression of critical adhesion–related genes (vinculin, talin, and RHOA) of human multipotent adipose stem cells, resulting in enhanced osteogenic differentiation on the nanofeatured Ti surface. It is also suggested that fibronectin adsorption may have only a weak effect on early cellular events of mesenchymal stem cells (MSCs) at least in the case of the nanostructured Ti implant surface incorporating Sr. Moreover, results indicate that Sr overrides the effect of Ca and other important surface factors (i.e., surface area and wettability) in the osteogenesis function of various MSCs (derived from human adipose, bone marrow, and murine bone marrow). In addition, surface engineering of nanostructured Ti implants using Sr ions is expected to exert additional beneficial effects on implant bone healing through the proper balancing of the allocation of MSCs between adipogenesis and osteogenesis. This work provides insight into the future surface design of Ti dental implants using surface bioactive ion chemistry and nanotopography.

  20. (abstract) Unidirectional Carbon/Carbon for Ion Engine Optics

    NASA Technical Reports Server (NTRS)

    Brown, D. Kyle

    1995-01-01

    Conventional ion engine optical grids are made from hydroformed molybdenum. Carbon/carbon has been utilized in place of molybdenum because of its lower sputter yield, which contributes a greatly increased engine life, and for its low cte, which allows more efficient engine operation. The requirements for this material are that it must have high stiffness, very tight dimensional tolerances, and can be optimized for an hexagonal hole pattern with a very high open area friction. The carbon/carbon for this application was fabricated from unidirectional tape prepreg, using pitch fiber, and was processed to a very high temperature. The use of unidirectional tape allowed for a sufficient number of plies to be used to generate a balanced three directional layup within the thickness constraints of the material, as well as providing strength and stiffness over that normally seen with fabric based carbon/carbons.

  1. F100 Engine Emissions Tested in NASA Lewis' Propulsion Systems Laboratory

    NASA Technical Reports Server (NTRS)

    Wey, Chowen C.

    1998-01-01

    Recent advances in atmospheric sciences have shown that the chemical composition of the entire atmosphere of the planet (gases and airborne particles) has been changed due to human activity and that these changes have changed the heat balance of the planet. National Research Council findings indicate that anthropogenic aerosols1 reduce the amount of solar radiation reaching the Earth's surface. Atmospheric global models suggest that sulfate aerosols change the energy balance of the Northern Hemisphere as much as anthropogenic greenhouse gases have. In response to these findings, NASA initiated the Atmospheric Effects of Aviation Project (AEAP) to advance the research needed to define present and future aircraft emissions and their effects on the Earth's atmosphere. Although the importance of aerosols and their precursors is now well recognized, the characterization of current subsonic engines for these emissions is far from complete. Furthermore, since the relationship of engine operating parameters to aerosol emissions is not known, extrapolation to untested and unbuilt engines necessarily remains highly uncertain. Tests in 1997-an engine test at the NASA Lewis Research Center and the corresponding flight measurement test at the NASA Langley Research Center-attempted to address both issues by measuring emissions when fuels containing different levels of sulfur were burned. Measurement systems from four research groups were involved in the Lewis engine test: A Lewis gas analyzer suite to measure the concentration of gaseous species 1. including NO, NOx, CO, CO2, O2, THC, and SO2 as well as the smoke number; 2. A University of Missouri-Rolla Mobile Aerosol Sampling System to measure aerosol and particulate properties including the total concentration, size distribution, volatility, and hydration property; 3. An Air Force Research Laboratory Chemical Ionization Mass Spectrometer to measure the concentration of SO2 and SO3/H2SO4; and 4. An Aerodyne Research Inc

  2. Interior of Vacuum Tank at the Electric Propulsion Laboratory

    NASA Image and Video Library

    1961-08-21

    Interior of the 20-foot diameter vacuum tank at the NASA Lewis Research Center’s Electric Propulsion Laboratory. Lewis researchers had been studying different electric rocket propulsion methods since the mid-1950s. Harold Kaufman created the first successful ion engine, the electron bombardment ion engine, in the early 1960s. These engines used electric power to create and accelerate small particles of propellant material to high exhaust velocities. Electric engines have a very small thrust, but can operate for long periods of time. The ion engines are often clustered together to provide higher levels of thrust. The Electric Propulsion Laboratory, which began operation in 1961, contained two large vacuum tanks capable of simulating a space environment. The tanks were designed especially for testing ion and plasma thrusters and spacecraft. The larger 25-foot diameter tank included a 10-foot diameter test compartment to test electric thrusters with condensable propellants. The portals along the chamber floor lead to the massive exhauster equipment that pumped out the air to simulate the low pressures found in space.

  3. Propulsion and Energetics Panel Working Group 15 on the Uniform Engine Test Programme

    DTIC Science & Technology

    1990-02-01

    among member nations in aerospace research and development: - Exchange of scientific and technical information : - Providing assistance to member nations...developed in one country often were used in airframes developed in another. Both situations require engine pertormance information which can be interpreted...earlier test of uniform aerodynamic models in wind tunnels under the auspices of the Fluid Dynamics Panel. A formal proposal was presented to the

  4. Effects of the six engine air breathing propulsion system on space shuttle orbiter subsonic stability and control characteristics

    NASA Technical Reports Server (NTRS)

    Mennell, R. C.; Soard, T.

    1974-01-01

    Experimental aerodynamic investigations were conducted on a 0.0405 scale representation of the -89B space shuttle orbiter in the 7.75 x 11.00 foot low speed wind tunnel during the time period September 4 - 14, 1973. The primary test objective was to optimize the air breathing propulsion system nacelle cowl-inlet design and to determine the aerodynamic effects of this design on the orbiter stability and control characteristics. Nacelle cowl-inlet optimization was determined from total pressure - static pressure measurements obtained from pressure rakes located in the left hand nacelle pod at the engine face station. After the optimum cow-inlet design, consisting of a 7 deg cowl lip angle, short cowl, 7 deg short diverter, and a nacelle toe-in angle of 5 deg was selected, the aerodynamic effects of various locations of this design were investigated. The 3 pod - 6 Nacelle configuration was tested both underwing and overwing in three different longitudinal locations. Orbiter control effectiveness, both with and without Nacelles, was investigated at elevon deflections of 0 deg, -10 deg and +15 deg and at aileron deflections of 0 deg and +10 deg about 0 deg elevon.

  5. A review of nuclear electric propulsion spacecraft system concepts

    NASA Technical Reports Server (NTRS)

    Deininger, W. D.; Nock, K. T.

    1990-01-01

    The last 25-30 years of system concepts and design philosophies for spacecraft employing nuclear-electric propulsion (NEP) are reviewed. NEP spacecraft-system design constraints and criteria are identified, including radiation exposure of humans and electronics, thermal control requirements, effluent contamination of spacecraft surfaces, surface erosion, launch-vehicle integration, operations and safety requirements, attitude control, EM interference, and power control and distribution. The impact on spacecraft design philosophy of these constraints and criteria is explored. Several NEP spacecraft are characterized and discussed with respect to the propulsion system used. The electric propulsion system catagories are electrothermal (arcjet), EM (magnetoplasmadynamic and pulsed-inductive thruster) and electrostatic (ion engine). A brief summary of the mission, nuclear power source, electric propulsion system, and spacecraft configuration are provided for each NEP spacecraft concept.

  6. Propulsion System Performance Simulation (PS**2) Computer Simulation to Evaluate Tank-Automotive Engine and Transmission Performance. A User’s Guide

    DTIC Science & Technology

    1988-09-30

    U.S. ARMY TANK-AUTOMOTIVE COMMAND RESEARCH, DEVELOPMENT & ENGINEERING CENTER Warren, Michigan 48397-5000 .Y_ E I’TY CLASSIFICATION OF THIS PAG3E Form... CLASSIFICATION lb RESTRICTIVE MARKINGS Unclassified NONE 2a. SECURITY CLASSIFICATION AUTHORITY 3. DISTRIBUTION /AVAILABILITY OF REPORT N/A Approved for...SOURCE OF FUNDING NUMBERS PROGRAM PROJECT TASK WORK UNI1T ELEMENT NO NO. NO. ACCESSRt I NO. 1 1. TITLE (include Security Classification ) Propulsion System

  7. The designing of launch vehicles with liquid propulsion engines ensuring fire, explosion and environmental safety requirements of worked-off stages

    NASA Astrophysics Data System (ADS)

    Trushlyakov, V.; Shatrov, Ya.; Sujmenbaev, B.; Baranov, D.

    2017-02-01

    The paper addresses the problem of the launch vehicles (LV) with main liquid propulsion engines launch technogenic impact in different environment areas. Therefore, as the study subjects were chosen the worked-off stages (WS) with unused propellant residues in tanks, the cosmodrome ecological monitoring system, the worked-off stage design and construction solutions development system and the unified system with the "WS+the cosmodrome ecological monitoring system+design and construction solutions development system" feedback allowing to form the optimal ways of the WS design and construction parameters variations for its fire and explosion hazard management in different areas of the environment. It is demonstrated that the fire hazard effects of propellant residues in WS tanks increase the ecosystem disorder level for the Vostochny cosmodrome impact area ecosystem. Applying the system analysis, the proposals on the selection of technologies, schematic and WS design and construction solutions aimed to the fire and explosion safety improvement during the LV worked-off stages with the main liquid propulsion engines operation were formulated. Among them are the following: firstly, the unused propellant residues in tanks convective gasification based on the hot gas (heat carrier) supply in WS tanks after main liquid propulsion engines cutoff is proposed as the basic technology; secondly, the obtained unused propellant residues in WS tanks gasification products (evaporated propellant residues + pressurizing agent + heat carrier) are used for WS stabilization and orientation while descending trajectory moving. The applying of the proposed technologies allows providing fire and explosion safety requirements of LV with main liquid propulsion engines practically.

  8. Heat transfer in aerospace propulsion

    NASA Technical Reports Server (NTRS)

    Simoneau, Robert J.; Hendricks, Robert C.; Gladden, Herbert J.

    1988-01-01

    Presented is an overview of heat transfer related research in support of aerospace propulsion, particularly as seen from the perspective of the NASA Lewis Research Center. Aerospace propulsion is defined to cover the full spectrum from conventional aircraft power plants through the Aerospace Plane to space propulsion. The conventional subsonic/supersonic aircraft arena, whether commercial or military, relies on the turbine engine. A key characteristic of turbine engines is that they involve fundamentally unsteady flows which must be properly treated. Space propulsion is characterized by very demanding performance requirements which frequently push systems to their limits and demand tailored designs. The hypersonic flight propulsion systems are subject to severe heat loads and the engine and airframe are truly one entity. The impact of the special demands of each of these aerospace propulsion systems on heat transfer is explored.

  9. Propulsion Health Monitoring of a Turbine Engine Disk Using Spin Test Data

    NASA Technical Reports Server (NTRS)

    Abdul-Aziz, Ali; Woike, Mark R.; Oza, Nikunj; Matthews, Bryan; Baaklini, George Y.

    2010-01-01

    This paper considers data collected from an experimental study using high frequency capacitive sensor technology to capture blade tip clearance and tip timing measurements in a rotating turbine engine-like-disk-to predict the disk faults and assess its structural integrity. The experimental results collected at a range of rotational speeds from tests conducted at the NASA Glenn Research Center s Rotordynamics Laboratory are evaluated using multiple data-driven anomaly detection techniques to identify abnormalities in the disk. Further, this study presents a select evaluation of an online health monitoring scheme of a rotating disk using high caliber sensors and test the capability of the in-house spin system.

  10. The NASA In-Space Propulsion Technology Project, Products, and Mission Applicability

    NASA Technical Reports Server (NTRS)

    Anderson, David J.; Pencil, Eric; Liou, Larry; Dankanich, John; Munk, Michelle M.; Kremic, Tibor

    2009-01-01

    The In-Space Propulsion Technology (ISPT) Project, funded by NASA s Science Mission Directorate (SMD), is continuing to invest in propulsion technologies that will enable or enhance NASA robotic science missions. This overview provides development status, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of aerocapture, electric propulsion, advanced chemical thrusters, and systems analysis tools. Aerocapture investments improved: guidance, navigation, and control models of blunt-body rigid aeroshells; atmospheric models for Earth, Titan, Mars, and Venus; and models for aerothermal effects. Investments in electric propulsion technologies focused on completing NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6 to 7 kW throttle-able gridded ion system. The project is also concluding its High Voltage Hall Accelerator (HiVHAC) mid-term product specifically designed for a low-cost electric propulsion option. The primary chemical propulsion investment is on the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. The project is also delivering products to assist technology infusion and quantify mission applicability and benefits through mission analysis and tools. In-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations.

  11. NASA's In-Space Propulsion Technology Project Overview, Near-term Products and Mission Applicability

    NASA Technical Reports Server (NTRS)

    Dankanich, John; Anderson, David J.

    2008-01-01

    The In-Space Propulsion Technology (ISPT) Project, funded by NASA's Science Mission Directorate (SMD), is continuing to invest in propulsion technologies that will enable or enhance NASA robotic science missions. This overview provides development status, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of aerocapture, electric propulsion, advanced chemical thrusters, and systems analysis tools. Aerocapture investments improved (1) guidance, navigation, and control models of blunt-body rigid aeroshells, 2) atmospheric models for Earth, Titan, Mars and Venus, and 3) models for aerothermal effects. Investments in electric propulsion technologies focused on completing NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system. The project is also concluding its High Voltage Hall Accelerator (HiVHAC) mid-term product specifically designed for a low-cost electric propulsion option. The primary chemical propulsion investment is on the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. The project is also delivering products to assist technology infusion and quantify mission applicability and benefits through mission analysis and tools. In-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations.

  12. Technical evaluation report on Propulsion and Energetics Panel 38th Meeting on Inlets and Nozzles for Aerospace Engines

    NASA Technical Reports Server (NTRS)

    Bowditch, D. N.; Monti, R.

    1972-01-01

    The application and use of inlets and nozzles in aerospace, V/STOL, and hypersonic propulsion systems are discussed. Data cover test techniques and facilities, experimental results from small rig tests to flight tests, and theoretical analysis of propulsion system flows. The problems associated with such a system are also discussed.

  13. Pulsed Laser Propulsion.

    DTIC Science & Technology

    1978-10-01

    afforded by a pulsed laser propulsion system over a CW laser propulsion system are 1) simplicity in engine design as a result of permitting the laser...to engineering and weight considerations. The lower boundary of the corridor is set by propellant feed considerations. To the right of this boundary...example, a OOJ -5 per pulse laser operating at 7 x 10 sec between pulses (14, 285 pps) is capable of powering a 30 lb (135 Nt)thrust rocket engine that has

  14. Re-engineering the Multimission Command System at the Jet Propulsion Laboratory

    NASA Technical Reports Server (NTRS)

    Alexander, Scott; Biesiadecki, Jeff; Cox, Nagin; Murphy, Susan C.; Reeve, Tim

    1994-01-01

    The Operations Engineering Lab (OEL) at JPL has developed the multimission command system as part of JPL's Advanced Multimission Operations System. The command system provides an advanced multimission environment for secure, concurrent commanding of multiple spacecraft. The command functions include real-time command generation, command translation and radiation, status reporting, some remote control of Deep Space Network antenna functions, and command file management. The mission-independent architecture has allowed easy adaptation to new flight projects and the system currently supports all JPL planetary missions (Voyager, Galileo, Magellan, Ulysses, Mars Pathfinder, and CASSINI). This paper will discuss the design and implementation of the command software, especially trade-offs and lessons learned from practical operational use. The lessons learned have resulted in a re-engineering of the command system, especially in its user interface and new automation capabilities. The redesign has allowed streamlining of command operations with significant improvements in productivity and ease of use. In addition, the new system has provided a command capability that works equally well for real-time operations and within a spacecraft testbed. This paper will also discuss new development work including a multimission command database toolkit, a universal command translator for sequencing and real-time commands, and incorporation of telecommand capabilities for new missions.

  15. Coupled multi-disciplinary simulation of composite engine structures in propulsion environment

    SciTech Connect

    Chamis, C.C.; Singhal, S.N.

    1992-01-01

    A computational simulation procedure is described for the coupled response of multi-layered multi-material composite engine structural components which are subjected to simultaneous multi-disciplinary thermal, structural, vibration, and acoustic loadings including the effect of hostile environments. The simulation is based on a three dimensional finite element analysis technique in conjunction with structural mechanics codes and with acoustic analysis methods. The composite material behavior is assessed at the various composite scales, i.e., the laminate/ply/constituents (fiber/matrix), via a nonlinear material characterization model. Sample cases exhibiting nonlinear geometrical, material, loading, and environmental behavior of aircraft engine fan blades, are presented. Results for deformed shape, vibration frequency, mode shapes, and acoustic noise emitted from the fan blade, are discussed for their coupled effect in hot and humid environments. Results such as acoustic noise for coupled composite-mechanics/heat transfer/structural/vibration/acoustic analyses demonstrate the effectiveness of coupled multi-disciplinary computational simulation and the various advantages of composite materials compared to metals.

  16. NDE using sensor based approach to propulsion health monitoring of a turbine engine disk

    NASA Astrophysics Data System (ADS)

    Abdul-Aziz, Ali; Woike, Mark R.; Abumeri, G.; Lekki, John D.; Baaklini, George Y.

    2009-03-01

    Rotor health monitoring and on-line damage detection have been increasingly gaining interest to manufacturers of aircraft engines, primarily to increase safety of operation and lower the high maintenance costs. But health monitoring in the presence of scatter in the loading conditions, crack size, disk geometry, and material property is rather challenging. However, detection factors that cause fractures and hidden internal cracks can be implemented via noninvasive types of health monitoring and or nondestructive evaluation techniques. These evaluations go further to inspect materials discontinuities and other anomalies that have grown to become critical defects that can lead to failure. To address the bulk of these concerning issues and understand the technical aspects leading to these outcomes, a combined analytical and experimental study is being thought. Results produced from the experiments such as blade tip displacement and other data collected from tests conducted at the NASA Glenn Research Center's Rotordynamics Laboratory, a high precision spin rig, are evaluated, discussed and compared with data predicted from finite element analysis simulating the engine rotor disk spinning at various rotational speeds. Further computations using the progressive failure analysis (PFA) approach with GENOA code [6] to additionally assess the structural response, damage initiation, propagation, and failure criterion are also performed. This study presents an inclusive evaluation of an on-line health monitoring of a rotating disk and an examination for the capability of the in-house spin system in support of ongoing research under the NASA Integrated Vehicle Health Management (IVHM) program.

  17. Re-engineering the Multimission Command System at the Jet Propulsion Laboratory

    NASA Astrophysics Data System (ADS)

    Alexander, Scott; Biesiadecki, Jeff; Cox, Nagin; Murphy, Susan C.; Reeve, Tim

    1994-11-01

    The Operations Engineering Lab (OEL) at JPL has developed the multimission command system as part of JPL's Advanced Multimission Operations System. The command system provides an advanced multimission environment for secure, concurrent commanding of multiple spacecraft. The command functions include real-time command generation, command translation and radiation, status reporting, some remote control of Deep Space Network antenna functions, and command file management. The mission-independent architecture has allowed easy adaptation to new flight projects and the system currently supports all JPL planetary missions (Voyager, Galileo, Magellan, Ulysses, Mars Pathfinder, and CASSINI). This paper will discuss the design and implementation of the command software, especially trade-offs and lessons learned from practical operational use. The lessons learned have resulted in a re-engineering of the command system, especially in its user interface and new automation capabilities. The redesign has allowed streamlining of command operations with significant improvements in productivity and ease of use. In addition, the new system has provided a command capability that works equally well for real-time operations and within a spacecraft testbed. This paper will also discuss new development work including a multimission command database toolkit, a universal command translator for sequencing and real-time commands, and incorporation of telecommand capabilities for new missions.

  18. Distributed Propulsion Vehicles

    NASA Technical Reports Server (NTRS)

    Kim, Hyun Dae

    2010-01-01

    Since the introduction of large jet-powered transport aircraft, the majority of these vehicles have been designed by placing thrust-generating engines either under the wings or on the fuselage to minimize aerodynamic interactions on the vehicle operation. However, advances in computational and experimental tools along with new technologies in materials, structures, and aircraft controls, etc. are enabling a high degree of integration of the airframe and propulsion system in aircraft design. The National Aeronautics and Space Administration (NASA) has been investigating a number of revolutionary distributed propulsion vehicle concepts to increase aircraft performance. The concept of distributed propulsion is to fully integrate a propulsion system within an airframe such that the aircraft takes full synergistic benefits of coupling of airframe aerodynamics and the propulsion thrust stream by distributing thrust using many propulsors on the airframe. Some of the concepts are based on the use of distributed jet flaps, distributed small multiple engines, gas-driven multi-fans, mechanically driven multifans, cross-flow fans, and electric fans driven by turboelectric generators. This paper describes some early concepts of the distributed propulsion vehicles and the current turboelectric distributed propulsion (TeDP) vehicle concepts being studied under the NASA s Subsonic Fixed Wing (SFW) Project to drastically reduce aircraft-related fuel burn, emissions, and noise by the year 2030 to 2035.

  19. NASA's Propulsion Research Laboratory

    NASA Technical Reports Server (NTRS)

    2004-01-01

    The grand opening of NASA's new, world-class laboratory for research into future space transportation technologies located at the Marshall Space Flight Center (MSFC) in Huntsville, Alabama, took place in July 2004. The state-of-the-art Propulsion Research Laboratory (PRL) serves as a leading national resource for advanced space propulsion research. Its purpose is to conduct research that will lead to the creation and development of innovative propulsion technologies for space exploration. The facility is the epicenter of the effort to move the U.S. space program beyond the confines of conventional chemical propulsion into an era of greatly improved access to space and rapid transit throughout the solar system. The laboratory is designed to accommodate researchers from across the United States, including scientists and engineers from NASA, the Department of Defense, the Department of Energy, universities, and industry. The facility, with 66,000 square feet of useable laboratory space, features a high degree of experimental capability. Its flexibility allows it to address a broad range of propulsion technologies and concepts, such as plasma, electromagnetic, thermodynamic, and propellant propulsion. An important area of emphasis is the development and utilization of advanced energy sources, including highly energetic chemical reactions, solar energy, and processes based on fission, fusion, and antimatter. The Propulsion Research Laboratory is vital for developing the advanced propulsion technologies needed to open up the space frontier, and sets the stage of research that could revolutionize space transportation for a broad range of applications.

  20. NASA's Propulsion Research Laboratory

    NASA Technical Reports Server (NTRS)

    2004-01-01

    The grand opening of NASA's new, world-class laboratory for research into future space transportation technologies located at the Marshall Space Flight Center (MSFC) in Huntsville, Alabama, took place in July 2004. The state-of-the-art Propulsion Research Laboratory (PRL) serves as a leading national resource for advanced space propulsion research. Its purpose is to conduct research that will lead to the creation and development of innovative propulsion technologies for space exploration. The facility is the epicenter of the effort to move the U.S. space program beyond the confines of conventional chemical propulsion into an era of greatly improved access to space and rapid transit throughout the solar system. The laboratory is designed to accommodate researchers from across the United States, including scientists and engineers from NASA, the Department of Defense, the Department of Energy, universities, and industry. The facility, with 66,000 square feet of useable laboratory space, features a high degree of experimental capability. Its flexibility allows it to address a broad range of propulsion technologies and concepts, such as plasma, electromagnetic, thermodynamic, and propellant propulsion. An important area of emphasis is the development and utilization of advanced energy sources, including highly energetic chemical reactions, solar energy, and processes based on fission, fusion, and antimatter. The Propulsion Research Laboratory is vital for developing the advanced propulsion technologies needed to open up the space frontier, and sets the stage of research that could revolutionize space transportation for a broad range of applications.