SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 44, ISSUE 8 - April 21, 2006
20 SPACECRAFT PROPULSION AND POWER
Includes main propulsion systems and components, e.g., rocket engines; and spacecraft auxiliary power sources.
For related information see also 07 Aircraft Propulsion and Power, 28 Propellants and Fuels, 15 Launch Vehicles and Launch Operations, and 44 Energy Production and Conversion.
20060010195 Princeton Univ., NJ, USA
Design, Fabrication, and Testing of Emissive Probes to Determine the Plasma Potential of the Plumes of VariousElectric Thrusters
Chen, Erinna M.; Summer Student Research Presentations; August 2005, pp. 31; In English; See also 20060010186; NoCopyright; Available from CASI only as part of the entire parent document
A significant problem in the use of electric thrusters in spacecraft is the formation of low-energy ions in the thrusterplume. Low-energy ions are formed in the plume via random collisions between high-velocity ions ejected from the thrusterand slow-moving neutral atoms of propellant effusing from the engine. The sputtering of spacecraft materials due tointeractions with low-energy ions may result in erosion or contamination of the spacecraft. The trajectory of these ions isdetermined primarily by the plasma potential of the plume. Thus, accurate characterization of the plasma potential is essentialto predicting low-energy ion contamination. Emissive probes were utilized to determine the plasma potential. When the ionand electron currents to the probe are balanced, the potential of such probes float to the plasma potential. Two emissive probeswere fabricated; one utilizing a DC power supply, another utilizing a rectified AC power source. Labview programs werewritten to coordinate and automate probe motion in the thruster plume. Employing handshaking interaction, these motionprograms were synchronized to various data acquisition programs to ensure precision and accuracy of the measurements.Comparing these experimental values to values from theoretical models will allow for a more accurate prediction oflow-energy ion interaction.Author
Electron Probes; Plumes; Propellants; Sputtering; Thrustors; Plasma Potentials; Spacecraft Contamination
20060010494 MBDA, Saint Arnoult, France
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Propulsion Systems for Hypersonic Flight
Falempin, Francois; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 9-1 - 9-28; In English;See also 20060010486; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy;Available from CASI on CD-ROM only as part of the entire parent document
On this scientific and technology basis, two families of possible application can be imagined for highspeed airbreathingpropulsion : reusable space launcher and military systems. By combining the high-speed airbreathing propulsion with aconventional rocket engine (combined cycle or combined propulsion system), it should be possible to improve the averageinstalled specific impulse along the ascent trajectory and then make possible more performing launchers and, hopefully, a fullyreusable one.
A lot of system studies have been performed on that subject within the framework of different and consecutiveprograms. Nevertheless, these studies never clearly concluded if a space launcher could take advantage of using a combinedpropulsion system or not.
Different possible military applications can be proposed :
1) tactical missile when penetration is thekey factor or when pure speed is necessary against critical time targets,
2) high speed reconnaissance drone with improvedmission safety and response time capability,
3) global range rapid intervention system based on a large body airplane, equippedwith high-speed very long range drones and missiles controlled by an on-board analysis/command team.
Considering requiredtechnology level and development risk for these both applications, it appears clearly that military application could bedeveloped more rapidly. Development of operational application, civilian or military, of the hypersonic airbreathingpropulsion depends of two key points : development of needed technologies for the fuel-cooled structure of the propulsionsystem, capability to predict with a reasonable accuracy and to optimize the aero-propulsive balance (or generalizedthrust-minus-drag balance). The most part of the technology development effort can be led with available ground test facilitiesand classical numerical simulation (thermics, mechanics...).Author
Propulsion System Configurations; Air Breathing Engines; Hypersonic Flight; Propulsion System Performance; AerodynamicDrag
20060010498 Florida Univ., Gainesville, FL, USA
Flameholding in Non-Premixed Supersonic Flows
Segal, C.; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 16-1 - 16-12; In English; See also20060010486; Original contains black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy; Available fromCASI on CD-ROM only as part of the entire parent document
A balance between the flame propagation speed and the fluid velocity must be maintained to maintain the flame in achemically reacting flow. Since the fluid velocity exceeds the flame speed in supersonic combustion applications, the flameholding issue is solved by generation of some sort of recirculation region which ensures sufficient residence time so that the processes involved - fuel-air mixing, ignition and chemical reactions propagation can take place to completion. These processes are determined by local conditions of gas composition, temperature and velocity and are substantially different innon-premixed cases, such as encountered in most practical applications, than in premixed cases which are easier to analyzeand predict.
A substantial database of flame stability exists for premixed gases (Ozawa, 1971, Huelmantell et al, 1957,Ogorodnikov et al, 2001) from which stability limits for the rich and lean flames have been obtained for a number of fuel/airsystems. The stability limit is usually cast in terms of a flameholding boundary on an equivalence ratio vs. a stability parameterplane. The stability parameter depends, in general, on the flow velocity, temperature, size and shape of the flameholder andhas received various formulations in different studies, from empirical formulations to expressions that reflect globalDamkholer numbers. In the case of non-premixed gases the determination of stability limits is less straightforward, mostly dueto the non-homogeneity of the parameters in the recirculation region behind the flameholder.
It is difficult to estimate the spatial species concentration and temperature distribution in the recirculation regions of these flows due to the presence oflarge gradients and the complex, three-dimensional, flow structure. These difficulties are compounded by the uncertainty inthe shape of the recirculation region, which depends on the amount of heat release which, in turn, is dictated by the localmixing and combustion efficiencies.
The following discussion is focused on the characteristics of the flowfield in the regionof a recirculation region with implications on the flameholding analysis and modeling. Derived from text
Flame Propagation; Flow Velocity; Chemical Reactions; Reacting Flow; Supersonic Combustion; Fuel Systems; Ignition;Gas Mixtures; Flame Stability
20060011005 Indian Space Research Organization, Trivandrum, India
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New Studies Related to the Improved Performance in the Stepper Motor for Solar Array Drive Assembly (SADA) for Indian Remote Sensing (IRS) Satellites
SadasivanAchari, V. T.; Devasahayam, Robert; Joseph, C. C.; Ravichandran, M. H.; Simon, Shaji; Journal of Spacecraft Technology, Volume 16, No. 1; January 2006, pp. 31-35; In English; See also 20060011004; Copyright; Avail.: Other Sources
This paper presents the design, analysis and development of 720-step high torque stepper motor for the Solar Array Drive Assembly mechanism used in IRS class satellites. The existing motor develops a peak holding torque of 1.1 Nm at 42 V with magnetic saturation and the torque pattern has the third and fifth harmonics in addition to the fundamental component. The stepper motor magnetic circuit design is optimized with FE Tools to obtain the maximum torque and make the torque pattern free from spatial harmonics without any change in overall envelope, which is implemented in the ver$cation hardware. It is found the desaturation model develops a peak torque of 2.6 Nm as the magnetic saturation is eliminated. With the increase in permanent magnet excitation arid change in tooth profile the peak torque is further improved to 5.1 Nm. The torque pattern is smoothened by providing unequal tooth pitch in stator and rotor. Author
Stepping Motors; Torque Motors; Design Analysis; Solar Arrays
20060011035 NASA Glenn Research Center, Cleveland, OH, USA
Fabrication Materials for a Closed Cycle Brayton Turbine Wheel
Khandelwal, Suresh; Hah, Chunill; Powers, Lynn M.; Stewart, Mark E.; Suresh, Ambady; Owen, Albert K.; March 2006; 14 pp.; In English; Third International Energy Conversion Engineering Conference (IECEC), 15-18 Aug. 2005, San Francisco, CA, USA; Original contains color illustrations Contract(s)/Grant(s): WBS 22-973-80-10 Report No.(s): NASA/TM-2006-214220; E-15468; IECEC-2005-5576; Copyright; Avail.: CASI: A03, Hardcopy
A multidisciplinary analysis of a radial inflow turbine rotor is presented. This work couples high-fidelity fluid, structural, and thermal simulations in a seamless multidisciplinary analysis to investigate the consequences of material selection. This analysis extends multidisciplinary techniques previously demonstrated on rocket turbopumps and hypersonic engines. Since no design information is available for the anticipated Brayton rotating machinery, an existing rotor design (the Brayton Rotating Unit (BRU)) was used in the analysis. Steady state analysis results of a notional turbine rotor indicate that stress levels are easily manageable at the turbine inlet temperature, and stress levels anticipated using either superalloys or ceramics. Author
Brayton Cycle; Radial Flow; Turbines; Rotors; Turbine Wheels; Inlet Temperature; Thermal Simulation
20060011037 NASA Glenn Research Center, Cleveland, OH, USA
NASA’s 2004 In-Space Propulsion Refocus Studies for New Frontiers Class Missions
Witzberger, Kevin E.; Manzella, David; Oh, David; Cupples, Mike; March 2006; 32 pp.; In English; 41st Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA; Original contains color illustrations Contract(s)/Grant(s): WBS 22-800-92-71 Report No.(s): NASA/TM-2006-214228; AIAA Paper 2005-4271; E-15475; Copyright; Avail.: CASI: A03, Hardcopy
The New Frontiers (NF) program is designed to provide opportunities to fulfill the science objectives for top priority, medium class missions identified in the Decadal Solar System Exploration Survey. This paper assesses the applicability of the In-Space Propulsion s (ISP) Solar Electric Propulsion (SEP) technologies for representative NF class missions that include a Jupiter Polar Orbiter with Probes (JPOP), Comet Surface Sample Return (CSSR), and two different Titan missions. The SEP technologies evaluated include the 7-kW, 4,100-second NASA’s Evolutionary Xenon Thruster (NEXT), the 3-kW, 2,700-second Hall thruster, and two different NASA Solar Electric Propulsion Technology Readiness (NSTAR) thrusters that are variants of the Deep Space 1 (DS1) thruster. One type of NSTAR, a 2.6-kW, 3,100-second thruster, will be the primary propulsion system for the DAWN mission that is scheduled to launch in 2006; the other is an ‘enhanced’, higher power variant (3.8-kW, 4,100-second) and is so-called because it uses NEXT system components such as the NEXT power processing unit (PPU). The results show that SEP is applicable for the CSSR mission and a Titan Lander mission. In addition, NEXT has improved its applicability for these types of missions by modifying its thruster performance relative to its performance at the beginning of this study. Author
Solar Electric Propulsion; Hall Thrusters; Ion Propulsion; Ion Engines; Sample Return Missions; Deep Space 1 Mission
Source: NASA
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