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Title: Non-nuclear power sources for deep space

Abstract

Electric propulsion and non-nuclear power can be used in tandem as a replacement for the current chemical booster and radioisotope thermoelectric generators now in use for deep space applications (i.e., to the asteroid belt and beyond). In current generation systems, electric propulsion is usually considered to be impractical because of the lack of high power for deep space, and non-nuclear power is thought to be impractical partly due to its high mass. However, when taken in combination, a solar powered electric upper stage can provide ample power and propulsion capability for use in deep space. Radioisotope thermoelectric generator (RTG) systems have generally been selected for missions only when other systems are absolutely unavailable. The disadvantages of radioisotopes include the need for nuclear safety as another dimension of concern in payload integration; the lack of assured availability of plutonium in the post-cold-war world; the enormous cost of plutonium-238; and the system complexity introduced by the need to continuously cool the system during the pre-launch phase. A conservative estimate for the total power for the solar array at beginning of life (BOL) may be in the range of 25 kW in order to provide 500 W continuous power at Jupiter. The availabilitymore » of {approximately} 25 kW(e) in earth orbit raises the interesting possibility of coupling electric propulsion units to this free electric power. If electric propulsion is used to raise the probe from low-earth-orbit to an earth-escape trajectory, the system could actually save on low-earth orbit mass. Electric propulsion could be used by itself in a spiral trajectory orbit raising maneuver to earth escape velocity, or it could be used in conjunction with a chemical upper stage (either solid rocket or liquid), which would boost the payload to an elliptical orbit. The concept is to begin the Earth-Jupiter trip with a swing-by near the Sun close to the orbit of Venus and perhaps even closer if thermal loads can be tolerated. During the solar swing-by, much more power will be produced by the solar panels, allowing the spacecraft's velocity to be increased significantly. The outbound leg of the journey can, therefore, be made much more quickly than with the classical trajectory. For the purposes of a Jupiter mission, it is assumed that 20 km/sec total delta-v would be required. For a payload envelope of 17,304 kg, a 1,900 sec Isp capability means that 11,386 kg of propellant would have to be consumed, leaving 5,917 kg for the mass of the probe plus dry mass of the upper stage. The thruster subsystem would require 765 kg of thruster subsystem mass, and probably less. Assuming tanks, regulators and valves amount to 10% of the propellant mass (very likely a pessimistic assumption), it is possible to assign a mass of 1,150 kg for the tankage subsystem. This results in a mass allowance of at least 4,000 kg for the probe. This compares favorably with the dry mass of 1,637 kg for Galileo, for example, and suggests that more than adequate margin exists. If the payload margin is used for battery storage, flyby missions to the outer planets may be possible.« less

Authors:
; ;
Publication Date:
Research Org.:
Applied Sciences Inc., Cedarville, OH (US)
OSTI Identifier:
20000409
Resource Type:
Conference
Resource Relation:
Conference: 33rd Intersociety Energy Conversion Engineering Conference, Colorado Springs, CO (US), 08/02/1998--08/06/1998; Other Information: 1 CD-ROM, uses operating systems Windows 3.x; Windows95,98,NT; Macintosh; UNIX. All systems need 2X CD-ROM drive., PBD: 1998; Related Information: In: Proceedings of the 33. intersociety energy conversion engineering conference, by Anghaie, S. [ed.], [3000] pages.
Country of Publication:
United States
Language:
English
Subject:
14 SOLAR ENERGY; 33 ADVANCED PROPULSION SYSTEMS; SPACECRAFT POWER SUPPLIES; SOLAR ELECTRIC PROPULSION; SPACE VEHICLES; SOLAR CELL ARRAYS; PHOTOVOLTAIC POWER SUPPLIES; MASS; PROPULSION SYSTEMS; DESIGN

Citation Formats

Kennel, E B, Tang, C, and Santarius, J F. Non-nuclear power sources for deep space. United States: N. p., 1998. Web.
Kennel, E B, Tang, C, & Santarius, J F. Non-nuclear power sources for deep space. United States.
Kennel, E B, Tang, C, and Santarius, J F. Wed . "Non-nuclear power sources for deep space". United States.
@article{osti_20000409,
title = {Non-nuclear power sources for deep space},
author = {Kennel, E B and Tang, C and Santarius, J F},
abstractNote = {Electric propulsion and non-nuclear power can be used in tandem as a replacement for the current chemical booster and radioisotope thermoelectric generators now in use for deep space applications (i.e., to the asteroid belt and beyond). In current generation systems, electric propulsion is usually considered to be impractical because of the lack of high power for deep space, and non-nuclear power is thought to be impractical partly due to its high mass. However, when taken in combination, a solar powered electric upper stage can provide ample power and propulsion capability for use in deep space. Radioisotope thermoelectric generator (RTG) systems have generally been selected for missions only when other systems are absolutely unavailable. The disadvantages of radioisotopes include the need for nuclear safety as another dimension of concern in payload integration; the lack of assured availability of plutonium in the post-cold-war world; the enormous cost of plutonium-238; and the system complexity introduced by the need to continuously cool the system during the pre-launch phase. A conservative estimate for the total power for the solar array at beginning of life (BOL) may be in the range of 25 kW in order to provide 500 W continuous power at Jupiter. The availability of {approximately} 25 kW(e) in earth orbit raises the interesting possibility of coupling electric propulsion units to this free electric power. If electric propulsion is used to raise the probe from low-earth-orbit to an earth-escape trajectory, the system could actually save on low-earth orbit mass. Electric propulsion could be used by itself in a spiral trajectory orbit raising maneuver to earth escape velocity, or it could be used in conjunction with a chemical upper stage (either solid rocket or liquid), which would boost the payload to an elliptical orbit. The concept is to begin the Earth-Jupiter trip with a swing-by near the Sun close to the orbit of Venus and perhaps even closer if thermal loads can be tolerated. During the solar swing-by, much more power will be produced by the solar panels, allowing the spacecraft's velocity to be increased significantly. The outbound leg of the journey can, therefore, be made much more quickly than with the classical trajectory. For the purposes of a Jupiter mission, it is assumed that 20 km/sec total delta-v would be required. For a payload envelope of 17,304 kg, a 1,900 sec Isp capability means that 11,386 kg of propellant would have to be consumed, leaving 5,917 kg for the mass of the probe plus dry mass of the upper stage. The thruster subsystem would require 765 kg of thruster subsystem mass, and probably less. Assuming tanks, regulators and valves amount to 10% of the propellant mass (very likely a pessimistic assumption), it is possible to assign a mass of 1,150 kg for the tankage subsystem. This results in a mass allowance of at least 4,000 kg for the probe. This compares favorably with the dry mass of 1,637 kg for Galileo, for example, and suggests that more than adequate margin exists. If the payload margin is used for battery storage, flyby missions to the outer planets may be possible.},
doi = {},
url = {https://www.osti.gov/biblio/20000409}, journal = {},
number = ,
volume = ,
place = {United States},
year = {1998},
month = {7}
}

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