Mission Steering Profiles of Outer Planetary Orbiters Using Radioisotope Electric Propulsion

Fiehler, Douglas; Oleson, Steven R.

doi:10.1063/1.1649581

Cited by 6 publications

(5 citation statements)

References 7 publications

(7 reference statements)

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“…The launch architecture chosen for the IIE provides more capability, but at a higher cost, than the previous studies 27,28,29,30,31,32 . displays the differences between the capabilities of these launch architectures 33,34,35,36 .…”

Section: IImentioning

confidence: 71%

“…Previous REP trajectory designs 27,28,29,30,31,32 showed that because of the low-acceleration capability of REP a certain mission architecture is optimal for outer solar system missions. This approach, consisting of a high-excess escape energy (C 3 ) launch from Earth (C 3 100 km 2 /s 2 ) followed by a long period of electric propulsion (EP) thrusting, has been shown to allow rendezvous of a small class spacecraft (dry mass less than 1000 kg) with many bodies throughout the outer solar system 32 .…”

Section: IImentioning

confidence: 99%

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Mission Design for the Innovative Interstellar Explorer Vision Mission

Fiehler

McNutt

2006

Journal of Spacecraft and Rockets

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Section: IImentioning

confidence: 71%

Section: IImentioning

confidence: 99%

Mission Design for the Innovative Interstellar Explorer Vision Mission

Fiehler

McNutt

2006

Journal of Spacecraft and Rockets

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“…In addition, these ARPSs could potentially enable small spacecraft to orbit outer planets with trip times comparable to flyby missions. Recent studies have shown that such an application is possible with ARPSs having specific mass of 100 -150 kg/kWe (or specific power of 6.67 to 10 We/kg) (Fiehler and Oleson, 2004).…”

Section: Introductionmentioning

confidence: 99%

Cascaded Thermoelectric Converters for Advanced Radioisotope Power Systems

El‐Genk¹,

Saber²

2004

AIP Conference Proceedings

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Three Cascaded Thermoelectric Converters (CTCs) are optimized for potential use in Multi-Mission Advanced Radioisotope Power Systems (MM-ARPS) for electrical powers up to 1 kW e , or even higher, in support of 7-10 year missions. The peak efficiencies of these CTCs of 9.43% to 14.32% are 40% to 110% higher than that of SiGe in State-of-the-Art (SOA) Radioisotope Thermoelectric Generators (RTGs). Such high efficiencies would significantly reduce the amount of 238 PuO 2 fuel and the total system mass for a lower mission cost. Each CTC is comprised of a SiGe top unicouple that is thermally, but not electrically, coupled to a bottom unicouple with one of the following three choices of thermoelectric materials: (a) p-leg of TAGS-85 and n-leg of 2N-PbTe; (b) p-leg of CeFe 3.5 Co 0.5 Sb 12 and nleg of CoSb 3 ; and (c) segmented p-leg of CeFe 3.5 Co 0.5 Sb 12 and Zn 4 Sb 3 and n-leg of CoSb 3 . The length of the top and bottom unicouples is 10 mm, but the cross-sectional areas of the n-and p-legs of the unicouples are optimized for maximum efficiency operation. They vary with the thermal power inputs of 1, 2, and 3 W th per SiGe unicouple, and the heat rejection temperature of 375 K, 475 K, and 575 K, from the bottom unicouple. Such geometrical optimization is at nominal hot shoe temperature of 1273 K for the SiGe unicouple and cold shoe temperature of either 780 K or 980 K, depending on the materials of the bottom unicouples. The hot shoe temperature of the bottom unicouples is 20 K lower than the cold shoe of the top SiGe unicouple, but the rate of heat input is the same as the rate of heat rejection from the top unicouple. The present results are conservative as they assume a contact resistance of 150 µΩ-cm 2 per leg for the top and the bottom unicouples in the CTCs; however, decreasing this resistance to 50 µΩ-cm 2 per leg could increase the current efficiency estimates by an additional 1 -2 percentage points.

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“…One similarity for all of the missions analyzed is that most of the energy required for the spacecraft to reach its destination is provided by the launch vehicle, via high C 3 , with the EP system to provide the remainder of the energy and then, for the missions that require it, accomplish the rendezvous. 8 More details of these missions can be found in the following sections.…”

Section: Mission Performancementioning

confidence: 99%

“…To obtain the shorter trip times, high excess escape energy launches are used. 8 Because of the very low thrust available to the spacecraft (electric propulsion at up to approximately 1 kW e ); the orbit nearest to Earth that an REP spacecraft can achieve is that equivalent to Jupiter's orbit about the Sun. However, beyond Jupiter's orbit, the spacecraft can perform missions to many outer planetary bodies.…”

Section: Introductionmentioning

confidence: 99%