Mission Capability Assessment of Cubesats Using a Miniature Ion Thruster
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JOURNAL OF SPACECRAFT AND ROCKETS Vol. 50, No. 5, September–October 2013 Mission Capability Assessment of CubeSats Using a Miniature Ion Thruster Ryan W. Conversano∗ and Richard E. Wirz† University of California, Los Angeles, Los Angeles, California 90095 DOI: 10.2514/1.A32435 The successful miniaturization of many spacecraft subsystems make CubeSats attractive candidates for evermore- demanding scientific missions. A three-cell CubeSat employing the miniature xenon ion thruster, which features high efficiency and impulse capability, yields a unique spacecraft that can be optimized for a variety of missions ranging from significant inclination changes in a low Earth orbit to lunar transfers. A nominal configuration of a high-ΔV CubeSat has a dry mass of approximately 6.3 kg, including a 0.75 kg payload, margins, and contingencies. Depending on the thruster and propellant tank configuration, this CubeSat is capable of delivering mission ΔV values from 1000 to over 7000 m∕s, enabling low-Earth-orbit inclination change missions and lunar missions. A parametric analysis on a three-cell high-ΔV CubeSat bus revealed that a range of payload volumes (up to nearly 1.4 units) and masses (up to nearly 6 kg) can be accommodated depending on the ΔV requirements and mission type. Additionally, this analysis showed that a high-ΔV three-cell CubeSat in a 600 km low Earth orbit can be designed to provide an inclination change of over 80 deg. Nomenclature d = dry (mass) A = surface area, m2 IR = IR source C = battery charge rate, A IR,sc = IR source facing D = degradation in = input d = depth of discharge int = internal F = view factor loss = loss g = Earth’s acceleration of gravity, m∕s2 min = minimum I = solar intensity, W∕m2 out = outflux i = inclination, deg p = propellant L = mission life, years pwr = power m = mass, kg R = radiator m_ = mass flow rate, kg∕s rad = radiated P = power, W r = rate p = power density, W∕m2 req = required Q = heat, W S = solar; sun facing T = thrust, N; temperature, K sc = spacecraft t = time, s, h sp = specific (impulse); space facing V = velocity, m∕s t = thrust; thruster α = absorptivity tr = transfer ε = emissivity tot = total η = efficiency θ = incidence angle ρ = specific power, W∕kg I. Introduction σ = Stefan–Boltzmann constant, W∕m2K4 MBITIOUS scientific missions focused on Earth and lunar A science offer valuable data for researchers; however, the number Subscripts and frequency of such missions have significantly decreased due to Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 increasingly high costs involved with developing and operating A = albedo lit conventional spacecraft. The increasing capabilities of autonomous a = annual microsatellites developed by the aerospace community have yielded b = battery the potential for microspacecraft to complete these demanding burn = thruster burn missions. A key challenge in using microsatellites for high-ΔV c = charge; circular missions, such as geocentric transfer orbits (GTO) to geo- synchronous orbits (GEO) and lunar transfers, is generating Δ Presented as Paper 2011-6083 at the Joint Propulsion Conference and sufficient V and power with a miniaturized propulsion system. Exhibit, San Diego Convention Center, San Diego, CA, 31–3 August 2011; Additionally, the small interior dimensions of microsatellites limit received 1 June 2012; revision received 27 August 2012; accepted for their ability to carry the mission-required propellant mass, especially publication 1 September 2012; published online 24 April 2013. Copyright © for a chemical propulsion system. Nevertheless, developments in 2012 by Ryan Conversano. Published by the American Institute of miniaturization techniques have led to new electric thruster designs Aeronautics and Astronautics, Inc., with permission. Copies of this paper may that can be used by microsatellites. One such thruster is Jet be made for personal or internal use, on condition that the copier pay the Propulsion Laboratory/University of California, Los Angeles’s (JPL/ $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood UCLA’s) miniature xenon ion (MiXI) thruster, on the scale of Drive, Danvers, MA 01923; include the code 1533-6794/13 and $10.00 in correspondence with the CCC. centimeters in diameter, which can operate below 100 W, while *Graduate Student, Department of Mechanical and Aerospace Engineer- maintaining the same high specific impulse and precision as larger ing; [email protected]. Student Member AIAA. ion thrusters [1–4]. This investigation aims to present the process by †Assistant Professor, Department of Mechanical and Aerospace Engineer- which a high-ΔV mission CubeSat can be designed and to ing; [email protected]. Senior Member AIAA. demonstrate that such satellites have the capability of carrying useful 1035 1036 CONVERSANO AND WIRZ scientific payloads on high-ΔV missions using miniaturized ion from an investigation of high-ΔV mission CubeSats. Details thruster technology. regarding the mass budget, the power budget, and the performance of A CubeSat is a satellite whose geometry is based on cubic cells, each major subsystem are discussed. Additionally, the payload, mass, each called a unit (U), and each measuring 10 cm along each edge. and capabilities design space of a high-ΔV CubeSat is explored. These cells can then be assembled to create rectangular-prism geometries and are named based on the number of adjoined cells (i.e., a two-cell CubeSat is called a 2U). The strength of CubeSat heritage II. Methodology for First-Order High-ΔV can be seen by the recent proliferation of microsatellite missions, CubeSat Design such as the biological science microsatellites PharmaSat (May 2009) A. Propulsion Subsystem and GeneSat-1 (December 2006), the second Canadian advanced The propulsion system selected for a specific mission clearly nanospace experiment satellite CanX-2 (April 2008), the geological defines the spacecraft’s ΔV capabilities and mission longevity. science QuakeSat (June 2003), and the electron loss and fields Conventional high-ΔV missions, such as the lunar reconnaissance investigation (ELFIN) CubeSat under development at UCLA by the orbiter and the Japanese SELENE-B program, have relied on – Institute of Geophysics and Planetary Physics [5 10]. Although these chemical propulsion to provide the necessary ΔV to perform spacecraft are able to fulfill useful scientific missions, they lack trajectory and orbit maneuvers [14,15]. However, SEP systems have a propulsion system that would enable significant on-orbit a highly successful heritage for high-ΔV long-duration missions. The maneuverability or orbit changes. Dawn spacecraft, for example, is using a trio of NASA Solar The potential for microsatellites as advanced scientific platforms Technology Application Readiness (NSTAR) program xenon ion utilizing solar electric propulsion (SEP) has been explored by thrusters to rendezvous with and orbit the asteroids Vesta and Ceres numerous research teams. A joint study by JPL and the Air Force [16]. The NSTAR thruster was originally demonstrated for space Research Laboratory examined the development of miniaturized exploration missions, one of which was the successful Deep Space 1 technologies that would enable enhanced microsatellite utility [11]. spacecraft [17]. The ESA SMART-1 satellite used a PPS-1350 Hall- Avariety of mission scenario concepts, including remote sensing and effect thruster to achieve a lunar orbit while demonstrating a variety on-orbit satellite servicing, were explored using spacecraft on the of new deep-space technologies [18]. The Hayabusa spacecraft order of 100 kg in mass. The study concluded that, if such missions employed four ion thrusters in an attempt to collect an asteroid are to be possible, a continuation of miniaturization of conventional sample from 433 Eros and return to Earth [19]. The emergence of the spacecraft subsystems is necessary. A similar investigation was CubeSat led to the development of commercially available cold-gas conducted by the European Space Agency (ESA), which involved micropropulsion systems, offering low-ΔV maneuvering capabilities the advanced microsatellite mission (AMM) spacecraft [12]. to CubeSats.‡ High-ΔV chemical propulsion systems are nearly The study addressed several design parameters, such as payload impossible to implement on CubeSats due to the spacecraft’s strict mass, available power, and thermal considerations of the AMM. The size constraints and the large propellant volume required, which is a AMM design study resulted in a spacecraft concept with a total product of the relatively low specific impulse for chemical thrusters. mass of approximately 120 kg, a significant payload fraction of Miniaturized SEP systems are strong candidates for high-ΔV approximately 50%, and deployable solar arrays to power an SEP CubeSat missions. Ion thrusters, in particular, are attractive for small system. Although these proposed missions require miniaturized spacecraft applications because of their favorable scaling capabilities subsystem components and an SEP system, the spacecraft concepts and high propellant efficiencies [20]. Assuming the use of a 3U are significantly larger than a CubeSat. Therefore, this investigation CubeSat frame, as described in Sec. II.D, the appropriate ion thruster aims to build upon the findings of these groups in an effort to provide performance characteristics must be determined. Ion thrusters are ΔV a capabilities assessment for a high- mission