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 3U CubeSat. capable of providing a range of specific impulses Isp from A primary limitation for high-ΔV microsatellites is the internal approximately 2000 to over 3500 s for current designs [20]. By volume available for the propulsion system and associated selecting a value for specific impulse, the thrust T produced by the propellant. The satellite configuration under consideration for high- thruster can then be calculated as ΔV missions using a single miniature ion thruster is a 3U CubeSat. This platform is of interest due to several key features, including low 2ηtPt development and mission cost, ease of construction, frequent launch T ˆ (1) Ispg opportunities, and successful heritage. By contrast, a larger satellite platform may benefit from a different set of unique characteristics where η is the thruster power efficiency, g is the acceleration due to that may provide for more ambitious mission capabilities, such as t gravity at Earth’s surface, and Pt is the input power to the thruster. increased scientific payload, higher system power, enabling higher This can be used to determine the propellant mass flow rate m_ : thrust or multiple thrusters, additional volume available for p propellant yielding higher ΔV, higher battery capacity for continuous T

Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 thrust during eclipse, and improved thermal control and margin. m_ p ˆ (2) I g It should be noted that conventional CubeSat standards currently sp allow a maximum mass of 4 kg and do not allow the use of Δ pressurized vessels beyond 1.2 atm [13]. These requirements are The maximum mission duration tburn and mission V provided by the based on the limitations of industry-standard CubeSat deployment thruster can then be estimated based on an assumed propellant mass m m systems and are aimed at reducing the risk to the primary spacecraft p and an approximate spacecraft dry mass d: sharing the same launch vehicle. Nevertheless, due to the possibility mp of mission-specific permission to launch a CubeSat that does not t ˆ (3) burn m_ fulfill these requirements, this investigation provides insight p regarding the potential for high-ΔV mission SEP CubeSats. The two primary objectives of this investigation are, first, to m ‡ m Δ d p describe in detail the first-order spacecraft and mission design V ˆ Ispg ln (4) process as it applies to high-ΔV CubeSat missions using SEP and, md second, to show the results of this process as it is applied to 3U high- ΔV CubeSat missions employing a miniature SEP system such as the The ΔV requirement for a low-thrust spiral trajectory between two MiXI thruster. circular orbits of different radii and inclination can be approximated The first objective is accomplished by examining major spacecraft from a simplified form of Edelbaum’s equation [21]: subsystems and discussing the necessary considerations and analytical relationships associated with each subsystem. The second ‡Data available online at http://www.micro-space.org [accessed objective is accomplished through a discussion of the recent findings 27 March 2012]. CONVERSANO AND WIRZ 1037

πΔi η ˆ η D (8) ΔV2 ˆ V2 ‡ V2 − 2V V cos (5) EOL BOL c1 c2 c1 c2 2 η η where EOL and BOL are the end of life and beginning of life solar cell where V and V are the initial and final circular orbit velocities, efficiencies, respectively. The power density p and specific power ρ c1 c2 η respectively, and Δi is the total inclination change required (in of the cells can be calculated by using Eq. (8) and replacing EOL with ρ η ρ radians). Equation (5) can be rearranged to determine the total pEOL or EOL and BOL with pBOL or BOL. From these values, the inclination change possible for a constant circular orbit altitude: approximate mass and solar cell array area can be determined: π V2 ‡ V2 − ΔV2 Ptot cos Δi ˆ c1 c2 (6) mSA ˆ (9) ρEOL 2 2Vc1Vc2

Calculating the mass of the propulsion subsystem is highly Ptot dependent on the specific components selected; however, some first- ASA ˆ η θ (10) order approximations can be made to achieve an initial spacecraft IS;min EOL cos s† design. An iterative process using the preceding equations yields θ the necessary propellant mass for a given mission. Margin and where s is the incidence angle to the sun and IS;min is the minimum contingency on the propellant mass, each of between 10 and 20%, are local solar intensity. An additional mass consideration may be added to the total propellant mass to allow for changes in the necessary if the solar array uses a motorized deployment system, spacecraft bus design and the mission plan during flight [20]. gimbals, extra support braces, etc. The power delivered by the solar The tankage mass, consisting of the propellant tank and its arrays PSA can now be determined at EOL and BOL as associated components, usually ranges from 5 to 15% of the required η θ propellant mass [20]. The mass of the necessary feed systems and PSA ˆ ASA IS;min cos s† (11) mounting hardware for the propellant system is approximated as an additional 20–30% of the tankage mass [20]. Again, a margin and Maintaining a fraction of the daytime orbit power during eclipse is contingency of between 10–20% should be added to these mass essential for many spacecraft and is generated using battery cells. The values [20]. mass of the batteries mb can be approximated by P t B. Power Subsystem b;req b mb ˆ ρ η (12) A satellite’s power system is highly customizable and is based on b bd the requirements of the individual spacecraft’s subsystems. The first stage in designing a satellite power system is determining the power where Pb;req is the power required from the battery, tb is the time ρ requirements for each component of the spacecraft’ subsystem, duration for the battery to supply power in hours, b is the specific η including propulsion, command and data handling (CDH), attitude energy of the battery, b is the battery power usage efficiency, and determination and control systems (ADCS), payload, etc. The input d is the depth of discharge. The battery charging rate must be considered to ensure that a complete charge can be accomplished power required Preq for a subsystem can be determined by noting that ∕η during each orbit day. The time to charge the batteries can be Preq ˆ Pin pwr, where Pin is the power input to the subsystem and η calculated by pwr is the product of all subsystem component efficiencies. The sum of these input powers yields the total power required by the spacecraft P t P . This total power must then be fed to each component by a power b;req b tot tc ˆ η (13) distribution unit (PDU), which in turn is fed by a power processing Vb bdCr unit (PPU) that connects to the high-voltage electronics assembly (HVEA). The power required at the PDU is calculated by where Vb is the battery output voltage and Cr is the charge rate of the batteries for a given charger. Cabling and harnessing, which consists PPDU ˆ Ptot∕ηPDU, where PPDU is the total power required from the of approximately 10–25% of the total mass of the power system, and PDU and ηPDU is the PDU efficiency. The power required from the PPU and HVEA can be calculated in a similar fashion based on each a margin and contingency of between 10–20% for the power system’s component’s respective efficiency. Note that the order (PDU, PPU, mass budget must also be considered [20]. then HVEA) must be maintained for correct power budgeting. A wiring efficiency (usually around 95%) is traditionally used for the C. Thermal Subsystem pre-PPU required power; the mass of the wiring and harnessing is Thermal considerations for CubeSat are perhaps the most generally 10–25% of the total power subsystem mass [20]. It is also demanding of all subsystems due to the high surface area-to-volume Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 customary to add an additional margin and contingency of between ratio of microsatellites. Internal components of the spacecraft can 10–20% onto this power budget [20]. This final value is the power often be operated between approximately −10 and 40°C, whereas the required from the solar arrays. solar arrays are usable between −150 and 100°C [20]. The survival Appropriate solar arrays must be implemented to supply the temperature depends on the specific system, but usually extends an required spacecraft power. To ensure that the solar arrays will provide additional 10°C from the operational range [20]. sufficient power throughout the CubeSat’s designed mission lifetime, To determine the temperature of the spacecraft, a thermal balance the condition of the arrays must be considered at the end of life (EOL) must be performed. To maintain an equilibrium condition in the in addition to the beginning of life (BOL). This yields a degradation spacecraft and avoid uncontrolled heating or cooling, we must have coefficient D defined as X X L Qin − Qout ˆ 0 (14) D ˆ 1 − Da† (7)

where Da is the approximate degradation per year (percent) and L is where Qin and Qout are the total heat into and out of the spacecraft, the mission duration in years. The performance of photovoltaic solar respectively. There are four sources of heat to the spacecraft: solar – cells ranges from 12 25% with several new commercially available energy Qs, Earth and moon albedo QA, internal heat production Qint, § designs that are 30% efficient [20]. The performance of these cells at and IR energy from planet shine QIR. The heat from the sun can be EOL can be calculated as calculated from α θ §Dataavailableonlineatwww.spectrolab.com[accessed17February2012]. QS ˆ ISASFS cos s† (15) 1038 CONVERSANO AND WIRZ

where QS is the influx of heat, α is the spacecraft absorptivity, AS is axis control system. It is a proven platform with a solid heritage and ¶ ** the spacecraft sunlit surface area, and FS is the spacecraft sun-facing readily available components. , Additionally, the industry-standard view factor. The heat from albedo can be approximated by Poly Picosat Orbital Deployer (P-POD) CubeSat deployment system accepts 3U-frame satellites, allowing for easy integration onto launch QA ˆ αρAISAAFA (16) vehicles and increased space-flight opportunities [22]. Although a variety of 3U CubeSat designs exist, several consistent guidelines can be found for their structural components. The frame where ρA is the albedo intensity factor of the celestial body being elements are generally machined from aluminum due to its lightweight, orbited, IS is the local solar intensity, AA is the spacecraft albedo lit surface area, and F is the albedo view factor. This equation can be high-strength-to-weight ratio, ease of machining, and low cost [20]. A – applied to the Earth and the moon, if necessary. The internal heat Commercially available 3U space frames have a mass of 200 300 g production of the spacecraft is depending on the design specifications, materials, etc. Additional braces and brackets, ranging in mass from 5–50 g depending on design, X X may be required for structural integrity of the frame and to provide Q ˆ Q ˆ P 1 − η † (17) int loss req p mounting points for the CubeSat hardware and payload [10,22]. The exterior panels are constructed from either metal or circuit board where Qloss is the heat dissipation by each powered component of the fiberglass and have a mass of approximately 50 g per 10 × 10 cm area spacecraft due to power inefficiencies. of coverage. They are fitted to the 3U frame to shield the internal The three sources of heat loss from the spacecraft are spacecraft hardware and offer mounting points for either fixed or deployable solar surface panel radiation to space Qsp, IR heat radiated to and from an arrays. It is also necessary to add a margin and contingency of between − IR source (Qrad;IR QIR), and heat radiated by the use of radiators or 15 and 25% for changes in the mass of these components [20]. special radiating surfaces QR. These are defined as follows:

4 4 E. Trajectory Analysis Qsp ˆ σεspAspFsp Tsc − Tsp† (18) A major component of a first-order mission design is the spacecraft’s proposed flight trajectory. Spacecraft trajectory analysis − σε ε 4 − 4 tools or custom trajectory models may be implemented to provide Qrad;IR QIR ˆ IR;sc IRAIRFIR Tsc TIR† (19) highly accurate models of a mission’s flight path. A first-order approximation of the mission duration ttr can be determined via an 4 4 ’ QR ˆ σεRARFsp Tsc − Tsp† (20) impulse calculation using the thruster s generated thrust T, the propellant exit velocity ve, and the change of mass of the spacecraft Δ where σ is the Stefan–Boltzmann constant, ε is the spacecraft during its mission m (i.e., propellant mass): SP Z space-facing emissivity, Asp is the spacecraft space-facing surface t Δ tr t area, FSP is the space-facing view factor, Tsc is the spacecraft surface m ˆ dt (22) 0 v temperature, Tsp is the temperature of free space (approximately ε ε 2.7 K), IR;sc is the IR source-facing spacecraft surface emissivity, IR If a constant thrust is assumed during the transfer, we can simplify is the IR source emissivity, AIR is the spacecraft surface area facing approximate duration of the transfer as the IR source, FIR is the IR source-facing view factor, TIR is the IR source surface temperature, εR is the radiator emissivity, and AR is the Δmv t ˆ e (23) radiator surface area. tr T The surface temperature of the spacecraft body or solar array during daylight operations can be determined from inputting Depending on the chosen trajectory, radiation hardening of the – Eqs. (15 20) into Eq. (14) and solving for Tsc: spacecraft may be necessary due to time spent in the Earth’sVan

α θ ρ σε 4 σε ε 4 1 Is ‰AsFs cos s†‡ AAAFAŠ‡Qint ‡ spAspFspTsp ‡ IR;sc IRAIRFIRTIR 4 Tsc ˆ σ ε ε ε (21) spAspFsp ‡ IR;sc IRAIRFIR†

The temperature of the solar arrays is calculated using the same Allan belts or other radiation environments. Determining the orbit Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 process outlined earlier; however, the amount of solar energy trajectory can be accomplished with either professional software or converted into power and used for the spacecraft must be subtracted in-house algorithms. from the total heat rejected by the arrays. This is accomplished by multiplying the solar power received from the sun by a factor of − η η F. Attitude Determination and Control Systems, Command and Data (1 s), where s is the solar power conversion efficiency of the Handling, and Communication Systems array. Additionally, the different absorptivities and emissivities for Although not a primary focus of this investigation, a complete first- the front and the back of the array must be considered. For operations order spacecraft design requires consideration of the ADCS and CDH during eclipse, heat terms relating to the solar energy and albedo must subsystems. To maintain the power required for a high-ΔV mission, a be neglected. three-axis stability system (rather than a spinning design) is necessary to ensure that the solar arrays remain accurately pointed toward the D. Structure sun during the orbit day. Although some commercially available An analysis of the 3U frame suggests that it offers sufficient miniature three-axis attitude control systems exist, most designs are interior volume for the hardware necessary for CubeSat operations, still too large for use on CubeSats and many are not yet flight †† along with a complete solar electric propulsion system and a three- qualified. The power requirements and mass of the electronics package supporting the three-axis control system must also be considered. The CDH subsystem, which acts as the brain of the ¶Details available online at http://www.cubesatkit.com [accessed 17 February 2012]. **Details available online at http://www.isispace.nl [accessed 17 Febru- ††Data available online at www.cubesatshop.com [accessed 17 Febru- ary 2012]. ary 2012]. CONVERSANO AND WIRZ 1039

Table 1 MEL for the HiVel spacecraft

Subsystem Component Mass, kg No. % Inert mass Structures 1.340 21.20 Frame 0.200 1 3.16 Brackets 0.005 20 0.08 Braces 0.005 8 0.08 30 × 10 cm exterior panel 0.150 4 2.37 10 × 10 cm exterior panel 0.050 2 0.79 Solar array gimbal and mounting system 0.100 3 1.58 Propulsion 0.650 10.29 MiXI thruster 0.250 1 3.96 Thruster gimbal 0.150 1 2.37 Tankage system 0.200 1 3.16 Feed system 0.050 1 0.79 Power 1.111 17.58 Solar arrays 0.087 3 1.38 Battery 0.250 1 3.96 Battery charger 0.050 1 0.79 PPU 0.150 1 2.37 PDU 0.100 1 1.58 HVEA 0.150 1 2.37 Wiring and harnesses 0.150 — 2.37 ACS 0.500 7.91 Three-axis momentum wheel system 0.400 1 6.33 Control system 0.100 1 1.58 Thermal 0.000 0.00 Communications 0.180 2.85 Antennas 0.015 2 0.24 Radios 0.075 2 1.19 CDH 0.150 2.37 CDH board 0.050 1 0.79 EPS board 0.100 1 1.58 Bus Dry Mass 3.931 62.21 Payload (assumed) 0.750 11.87 Inert Mass 4.681 74.07 20% margin 0.936 14.81 15% contingency 0.702 11.11 Inert mass w/margin and contingency Propellant 6.319 100.00 High-ΔV mission 1.410 22.31 Lunar mission 2.330 36.87 High-ΔV wet mass 7.729 Moon mission wet mass 8.649

spacecraft, receives telemetry information and governs the Van Allan belts and would lead to decreased mission times due to the spacecraft’s actions when not directly controlled by a human excess ΔV granted by the launch vehicle. A payload mass of 0.75 kg operator. Finally, a comprehensive autonomous navigation software with a power requirement of 15 W was assumed for all ensuing mass package is necessary for the spacecraft to operate successfully. and power budgets unless otherwise stated. The payload was The communications subsystem is a CubeSat’s link to Earth. It is assumed be operational only when the propulsion system was composed of the satellite’s onboard antennas, radios, and data circuit inactive. These values are loosely based on UCLA’s ELFIN CubeSat boards. A key difficulty with lunar missions is transmitting and payload requirements of approximately 13% of the total spacecraft receiving information over the large distances between the Earth and dry mass and 20% of the total spacecraft power [9,10]. Table 1 the moon (approximately 384,000 km). A 3U CubeSat’s limited displays the mass equipment list (MEL) of the HiVel CubeSat; each internal dimensions require a space-efficient, high-powered receiver subsystem's components and first-order design is herein discussed. and transmitter in conjunction with any appropriate antennas to Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 deliver the performance necessary for a lunar transfer; this hardware A. Structure and supporting software is currently entering the commercial market ‡‡ The HiVel spacecraft was designed to use a 3U chassis. Many of [23]. HiVel’s structural elements are based on the ELFIN CubeSat, and although ELFIN is a 3U spinner-type spacecraft, the basic design III. Results for High-ΔV CubeSat Investigation elements and overall mass of the chassis is analogous to HiVeland are suitable for initial sizing and mass calculations [9,10]. The spacecraft design methodology discussed earlier was used to The majority of HiVel’s structures, including the frame, brackets, develop a first-order high-ΔV mission CubeSat concept. The satellite and braces, are assumed to be machined from 6061-T6 aluminum to under investigation will herein be referred to as the HiVel CubeSat to ensure a strong, lightweight spacecraft. The 3U frame designed for Δ represent a high- velocity CubeSat. The spacecraft was assumed to HiVel has an estimated mass of 200 g, requiring 20 brackets and eight begin its missions in a 600 km low Earth orbit (LEO) with an braces, each with an approximate mass of 5 g. The exterior panels Δ inclination of 60 deg and no excess V granted by the launch vehicle were designed to be constructed from 47- or 63-mm-thick circuit or deployment mechanism. This orbit was chosen due to the board fiberglass or machined from thin-wall 6061 T6 aluminum. The frequency of CubeSat launch opportunities to similar orbits and to mass per 10 × 10 cm section of either of these panel materials is Δ demonstrate the impressive V capabilities of the HiVel bus. If the approximately 50 g; a 3U CubeSat such as HiVel requires four selected mission requires a lunar transfer, beginning the mission in a 10 × 30 cm sections and two 10 × 10 cm sections for complete ’ high-energy GTO would be ideal to minimize time spent in Earth s external coverage. These mass values were taken from approximations made by the ELFIN program (CAD analysis) and the aforementioned ‡‡Data available online at ipp.gsfc.nasa.gov [accessed 20 May 2012]. commercially available CubeSat developer kits [10]. 1040 CONVERSANO AND WIRZ

Fig. 1 MiXI thruster with miniature hollow cathode fitted to a two-axis gimbal concept. Fig. 2 HiVel bus (ΔV of 6000 m∕s) in outboard (top) and inboard (bottom) configurations. The mass for the gimbals and mounting hardware for each of the three solar arrays was approximated at 40% of the total solar array begin its mission with no excess ΔV provided by the launch vehicle mass, resulting in a mass of approximately 100 g per array. This is a or deployment mechanism, thereby generating all of its ΔV from the conservative estimate when compared with the reported mass of the MiXI thruster. Given these initial conditions, the propellant mass for HaWK gimbaled solar array system, which has a total mass of a high-ΔV mission requiring 4500 m∕s, for example, or a lunar approximately 280 g for two full arrays, including solar cells [24]. A mission transfer, requiring a ΔV of approximately 7100 m∕s, can be dry mass margin of 20% and a contingency of 15% were applied, determined. The HiVel spacecraft, with an approximate dry mass of yielding an approximate total structural mass of 1.85 kg. 6.3 kg (including margin and contingency), requires approximately 1.41 kg of propellant for the aforementioned high-ΔV mission, B. Propulsion Subsystem whereas the lunar mission requires approximately 2.33 kg of The demanding propulsion requirements and long durations of a propellant; both of these values include a 10% margin, 15% high-ΔV mission led to the implementation of an SEP system on the contingency, and a 5% propellant fraction remaining after the transfer HiVelspacecraft. The MiXI thruster is well suited for use on a CubeSat for end-of-mission or emergency maneuvers. The tankage fraction mission due to its small 3 cm diameter and favorable performance was assumed to be approximately 10% of the required propellant characteristics [1–3]. Other thrusters on this scale that may be mass and the feed system was assumed to be approximately 25% of considered for CubeSat missions include miniature radio-frequency the tankage mass. This results in an approximate required tankage ion thrusters, miniature microwave ion thrusters, and several others mass of 200 g and an approximate feed system mass of 50 g. A [25–29]. For this investigation, however, the MiXI thruster is used as propellant tank consisting of a cylindrical middle section with the baseline. The MiXI thruster has demonstrated thrust values of 0.1– hemispherical ends was assumed. The semispherical ends have a 1.553 mN at a specific impulse of over 3000 s and is capable of maximum radius of 4.5 cm (due to volumetric constraints of the 3U providing a ΔV of over 7000 m∕s for certain HiVel configurations [1– CubeSat frame), whereas the length of the tank can be varied 3]. Although the thruster can be operated anywhere from 20–60 W, a depending on the ΔV requirement for the mission. The tank’s inner model with a 30 W nominal operating power was selected (primarily diameter was assumed to be 95% of the outer diameter, which results due to power considerations) to generate a thrust of approximately in a conservatively robust design for the tank. 1.43 mN with an assumed specific impulse of 3000 s. The thruster was assumed to deliver a conservative total efficiency of 45% (proven C. Power Subsystem values up to 56%) and a conservative propellant utilization efficiency of 70% (proven values of up to 82%) [30,31]. The thruster’smassis Power generation requirements for the HiVel spacecraft depend approximately 250 g, based on the measured masses of an engineering primarily on the input power for the MiXI thruster with additional model of the MiXI thruster and associated components. considerations for the communications, CDH, and payload To maintain the desired thrusting vector, the MiXI thruster will be subsystems. HiVel is designed to use spring-deployed, gimbaled mounted onto a two-axis gimbal system (see Fig. 1). The gimbal was solar arrays to collect the necessary power for daylight operations. Commercially available gimbaled solar array systems for CubeSats, Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 assumed to have a mass of 150 g and a power requirement of 2 W, ’ which are scaled-down values based on a larger design by Asadurian such as the HaWK, provide insufficient power to meet HiVel s high [32]. The thruster can be mounted in either an inboard or outboard power requirements; therefore, a similar gimbal system affixed to a configuration. The inboard configuration assumes that the entire larger multipanel solar array is necessary [24]. As mentioned in propulsion system and thruster will fit within the interior volume of a Sec. III.A, the solar panel mounting and gimbal system was assumed 3U CubeSat. The outboard configuration extends the MiXI thruster to have a mass of 100 g per array and a power requirement of outside of the rear of the spacecraft. The two configurations approximately 2 W, which was provided for day and night operations are shown in Fig. 2. The inboard variant benefits from meeting as an aid to the ADCS system. The high-power demands of the MiXI the current CubeSat requirements of fitting inside a standard thruster prevent its operation during eclipse; however, it was assumed ’ P-POD deployment system, whereas the outboard variant provides that 10 W of battery power is used to keep the thruster s hollow significant benefits to the thermal control of the spacecraft. The cathodes warm during eclipse, which generates greater reliability in outboard CubeSat configuration requires a nonstandard deployment thruster startup. ’ system; however, the modified P-POD system used for the GeneSat-1 A combination of the spacecraft s three-axis control system, the mission, which allows the CubeSat to use the volume inside gimbaled solar arrays, and the two-axis gimbaled thruster will of the deployment spring, would accommodate the outboard HiVel provide accurate sun pointing, while maintaining a desired thrust configuration [33]. The additional volume is cylindrical, with a vector. The three-axis system, a scaled version of the miniature four- diameter of approximately 6.5 cm and a depth of approximately 4 cm, wheel XACT system by Blue Canyon Technologies, was assumed to §§ which is sufficient for MiXI and its associated hardware. have a total power requirement of 5 W. Highly efficient methods for As previously stated, a launch inclination of 60 deg and an initial orbit of 600 km were assumed. The HiVel spacecraft was assumed to §§Data available online at bluecanyontech.com. [accessed 20 May 2012] CONVERSANO AND WIRZ 1041

Fig. 3 HiVel spacecraft in stowed configuration (left) and deployed configuration (right). Both the outboard configuration (top) and inboard configuration (bottom) are shown.

maintaining three-axis control using a two-axis solar array gimbal, a gimbaled electric thruster, and a momentum wheel system have been developed by Randolph et al.; such methods may be implemented for a high-performance CubeSat such as HiVel [34]. The HiVel spacecraft uses Spectrolab UTJ (ultra-triple-junction solar cells which are capable of providing a power density of over 380 W∕m2, a specific power of over 450 W∕kg, and an efficiency of approximately 28% at BOL [22]. A degradation of 7.8%, which accounts for assembly, design, and storage-time losses, was assumed to occur before launch with an annual degradation of 0.5% during the mission [20]. The view factor to the sun was approximated as 0.9 and a solar cell packing fraction of 0.9 was assumed [20]. With an assumed average solar incidence angle of 30 deg (a conservative estimate given gimbaled solar arrays), these values lead to a necessary solar array area of approximately 0.33 m2 with a total solar array mass of approximately 260 g. As can be seen from Fig. 3, no solar cells are mounted on the spacecraft body to avoid overheating of the Fig. 4 HiVel spacecraft in the outboard configuration with the spacecraft bus caused by the high thermal absorptivity of the cells. simplified three-axis momentum wheel system featured. HiVel’s PPU has an approximate mass of 150 g with a 92% efficiency, whereas the PDU has a mass of approximately 100 g and has an efficiency of 85%. These values represent a very conserva- tive design for the power subsystem. The HVEA’s mass was approximated as 150 g with an efficiency of 97%. These values are based on approximations given in New SMAD and mass quotes from spacecraft electronics manufacturers [20]. A set of battery cells are included in the design to keep the MiXI thruster’s cathodes warm during eclipse. The batteries are also designed to provide a minimum

Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 of 1 h of emergency power for flight-critical subsystems during eclipse. The required battery mass is 250 g; a battery charger was also included at a mass of 50 g with an assumed power requirement of 5 W (additional power may be supplied while the propulsion system is inactive if necessary). Wiring and cabling mass was found to be approximately 180 g, with an assumed electrical efficiency of 95%. A margin of 20% and a contingency of 15% were applied to the final power budget. The entire power electronics and battery assembly is represented by the cyan box in Figs. 2, 4, and 5. The power budget for HiVel was divided into two mission phases: Fig. 5 HiVel spacecraft in the outboard configuration with simplified cruise and science/communication. The cruise phase assumes circuit boards featured. These boards control the communications, CDH, EPS, and ADCS systems. operation of the propulsion subsystem, the ADCS subsystem, the communications subsystem, the battery charger from the power subsystem, and the solar array gimbals from the structures sub- system. The science/communications phase assumes operation of Although this is a high power budget for a 3U CubeSat, commercially the communications subsystem, the CDH subsystem, the ADCS available CubeSat solar array designs are capable of providing subsystem, the solar array gimbals from the structures subsystem, approximately 70 W orbit average power, suggesting that 77 W is and the payload subsystem. The cruise phase requires a total power of certainly possible in the near future. A complete summary of all approximately 77 W from the solar arrays, whereas the science/ subsystem power requirements for each mission phase is provided in communications phase requires a total power of approximately 63 W. Table 2 in Sec. IV. 1042 CONVERSANO AND WIRZ

Table 2 PEL for the HiVel spacecraft

Subsystem Component Cruise phase, W Science/Communications phase, W Efficiency Structures 2.61 2.61 Solar array gimbal and mounting system 2.00 2.00 0.77 Propulsion 35.02 0.00 MiXI thruster 30.00 0.92 Thruster gimbal 2.00 0.83 Power 6.54 0.00 Battery charger 5.00 0.77 ADCS 8.07 8.07 Three-axis ACS 2.00 2.00 0.60 Control system 3.00 3.00 0.64 Thermal 0.00 0.00 0.00 Communications 2.00 3.14 Antennas 1.00 1.00 0.64 Radios 1.00 1.00 0.64 CCH 0.00 15.69 CDH board 5.00 0.64 EPS board 5.00 0.64 Payload 0.00 15.00 Power Required 54.24 44.51 Wiring/cabling 0.95 Pre-HVEA power required 57.10 46.85 20% margin 11.42 9.37 15% contingency 8.56 7.03 Power required from solar arrays 77.08 63.24

aEfficiency column includes component and PDU/PPU efficiencies (where applicable).

A key challenge for developing the necessary power for a high-ΔV 245 K (−28°C). Appropriate orbit selection, the addition of heaters mission is avoiding shadowing during eclipse. These effects are on specific subsystem components, or the use of battery waste heat, greatest for low-altitude orbits and decrease with increasing orbit however, may reduce or eliminate this problem. radius. Maintaining a desired thrust vector will dictate solar array With a larger range of operational and survivable temperatures, the pointing; however, sufficient power generation will be accomplished solar arrays are more thermally robust. The Spectrolab UTJ solar cells by the use of the three-axis momentum wheel system, gimbaled solar have a front emissivity and absorptivity of approximately 0.85 and arrays, and a two-axis thruster gimbal. Although not thoroughly 0.92, respectively [22]. The back of the solar panel will be coated in explored in this investigation, shadowing can also be minimized by black paint with an emissivity and absorptivity of 0.89 and 0.92, placing the spacecraft in a high-power generating orbit, such as GTO respectively [20]. For this configuration, the arrays reach a or a dawn–dusk LEO orbit. temperature of approximately 325 K (52°C) during daylight operations and drop to approximately 234 K (−38°C) during eclipse. D. Thermal Subsystem These values are well within the operational and survivable As previously discussed, the thermal considerations for an SEP 3U temperatures of typical solar arrays [20]. CubeSat are critical to ensure survivability of the spacecraft. HiVel’s small size makes it highly susceptible to large temperature fluctua- E. Trajectory Analysis tions during the course of a single orbit. HiVel currently employs no As previously suggested, it is assumed that the HiVel spacecraft will external radiators. To maintain thermal control, while keeping the be inserted into a 600 km LEO at an inclination of 60 deg with no design as simple as possible, all exterior panels of the spacecraft were excess ΔV. For lunar-mission transfers, the moon was assumed to have assumed to be coated in aluminized Kapton, which has an an altitude of 384,000 km and an inclination of 23 deg. The total absorptivity range of 0.34–0.46 and an emissivity range of 0.55–0.86 transfer time for a lunar-mission transfer is approximately 450 days; depending on the thickness selected [20]. Mounting the MiXI high-ΔV missions starting from LEO, GTO, and GEO can achieve thruster in an outboard configuration (see Figs. 2 and 3) results in a similar mission durations. These long mission times may require Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 significant reduction in internal heat production. By performing a radiation hardening of the spacecraft due to the time spent in the Earth’s calculation of the view factor of the thruster to the spacecraft using a Van Allan belts and outside the magnetosphere. The increased mass of solid angle approximation centered at the thruster discharge, radiation-hardened equipment and shielding materials will be approximately 8∕3π steradians of the maximum 4π are not intersec- considered in future studies, but for now it is assumed to be included ting the spacecraft bus (66%). Therefore, it was conservatively in the design margins used in this investigation. assumed that 60% of the waste heat produced by the thruster A spiral trajectory model developed by Tarzi et al. at UCLA was was radiated directly to space. Using a thermal balance as outlined used to confirm the approximate mission and spacecraft parameters in Sec. II.D, an average body surface temperature for the of HiVelfrom the Earth to the moon [35]. The model uses the findings aforementioned design is approximately 291 K (18°C) during of Edelbaum [21] to develop approximate spiral trajectories between daylight operations, and 275 K (2°C) during eclipse is expected. Both two celestial bodies. It accepts inputs relating to initial orbit of these values drop approximately 30 K during emergency orientation and spacecraft properties and generates graphical operations (only necessary subsystems remain powered and representations of the spiral transfer, propellant mass requirements, operational when a hardware or software error occurs, limiting transfer time, etc. internal heat generation). By comparison, the inboard configuration yields a sunlit temperature of 307 K (34°C), which lies beyond the F. Attitude Determination and Control Systems, Command and Data upper operational limits of many subsystem components; the same Handling, and Communications Systems eclipse temperature is observed. The eclipse emergency operations Navigation to the moon and accurate pointing of the solar arrays is condition may cause survival issues with certain subsystem assumed to be accomplished by a three-axis momentum wheel components as temperatures drop to a minimum of approximately system in conjunction with gimbaled solar array panels and CONVERSANO AND WIRZ 1043

a two-axis thruster gimbal. A commercially available three-axis reaction wheel system by Blue Canyon Technologies requires approximately 0.5 U and a total mass of under 700 g. On the HiVel spacecraft, this system would be working in conjunction with the solar array and thruster gimbals to track the sun within an assumed incidence angle of 30 deg; therefore, it was assumed that this system could be scaled to 500 g with a total volume of 0.25 U. Navigation to the moon can be accomplished partially by using the GPS-enhanced onboard navigation system developed by NASA Goddard. This system is only useful to approximately 50 Earth radii (note, lunar orbit is ∼60RE); therefore, supplemental celestial navigation for the remainder of the transfer will be required [23]. The HiVel avionics is based on the ELFIN spacecraft for initial mass and power considerations [9,10]. The system employs a series of CDH and energy power system (EPS) circuit boards in addition to a) two radios and two antennas. These antennas may be integrated into the aft edges of the four-panel solar arrays; however, depending on the frequency band used, additional antenna length may be required and would necessitate small antenna whips to be trailed off of the solar array tips. This system has a total mass of approximately 450 g and a daylight operations power budget of approximately 26 Wafter a 20% margin and 15% contingency for both mass and power. Although this system is known to be sufficient for near-Earth missions, certain components will require modifications or sub- stitutions to ensure that the system will be able to receive and transmit information from the moon to the Earth.

IV. Summary of Results The tables summarize the key design values for the HiVel spacecraft in the form of a mass equipment list (MEL) and power b) estimation list (PEL). Fig. 6 Available payload (PL) volume vs ΔV capability for a high-ΔV 3U CubeSat with an a) outboard and b) inboard MiXI thruster configuration assuming a payload density of 1.33 kg∕U. A. High-ΔV CubeSat Design Space For HiVel to perform demanding scientific missions, it must be capable of delivering sufficient ΔV while carrying a useful payload. All of the curves show a sloping trend before becoming linear. This To understand the relationship between ΔV capabilities and payload is because the propellant tank was assumed to be spherical with a capacity, three specific design tradeoffs must be considered. First, as maximum allowable radius of 4.5 cm, thus generating a nonlinear the ΔV requirement increases for a given mission, the interior volume relationship between the volume required for the propellant tank and available for the useful payload decreases because of the increased ΔV capability. Once the 4.5 cm radius was reached, the radius was volume required for the propellant tank. This suggests a direct held constant while the spherical tank was bisected and a cylindrical Δ correlation between mission-required V and payload capacity. section was placed in between the two hemispheres. Because the Second, for a given flight mass, a relation exists between the payload volume of a cylinder varies linearly with its length, the graphs then mass available and the propellant mass required to generate a desired show a linear trend. Δ V. To generate this relation, a spacecraft dry mass (including The payload mass available in a high-ΔV 3U CubeSat as a function payload) must be assumed. Third, a direct relationship exists between ’ Δ Δ of the propellant mass is displayed in Fig. 7. The plot shows two sets the spacecraft s propellant mass and the V or inclination change i of data: The solid lines represent the results for a dry spacecraft mass capabilities. of 3 kg without payload, and the dashed line represents the results for To explore the capabilities and limitations of HiVel based on these a dry spacecraft mass of 6 kg without payload. This large mass range three key relations, a parametric analysis of the high-ΔV 3U CubeSat was considered to allow for changes in build materials, subsystem Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 was performed. Using values from the propulsion system discussed in Sec. III.C, plots were generated for a range of spacecraft dry and wet masses. By interpolating between these parameters, it is possible to establish the approximate dry mass, flight mass, ΔV, and Δi capabilities of a 3U CubeSat given only information about the required payload geometry and mass. Conversely, the approximate payload capacity and flight performance can be found based on an assumed dry mass and flight mass. The available payload volume for a high-ΔV 3U CubeSat with an outboard or inboard MiXI thruster configuration as a function of ΔV capability is show in Figs. 6a and 6b, respectively. The payload available was calculated by taking the total available volume (3 Us) and subtracting the volume required for the electronics package and ADCS (approximately 1.25 U), the MiXI thruster and supporting feed systems (approximately 0.5 U), and the volume required for the propellant tank. Seven curves are plotted, representing a range of total dry mass of the spacecraft, including payload. Given the required masses of the various subsystems as discussed earlier, a range of 4– 7 kg for the spacecraft dry mass is reasonable for a high-ΔV 3U Fig. 7 Available payload mass vs propellant mass for 3 kg (solid) and CubeSat. 6 kg (dashed) dry HiVel without payload. 1044 CONVERSANO AND WIRZ

a)

b) Fig. 8 ΔV (solid) or Δi (dashed) capability vs propellant mass for 3 kg dry HiVel without payload. Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 a)

b) Fig. 9 ΔV (solid) or Δi (dashed) capability vs propellant mass for 6 kg dry HiVel without payload. CONVERSANO AND WIRZ 1045

Table 3 Summary of parametric analysis

Dry mass w∕PL, kg MiXI configuration PL volume range, U ΔV range, ms 4 Inboard 0–0.82 0–7000‡ 4 Outboard 0–1.37 0–7000‡ 5.5 Inboard 0–0.76 0–6700 5.5 Outboard 0–1.32 0–7000‡ 7 Inboard 0–0.72 0–5400 7 Outboard 0–1.22 0–7000‡ Wet mass, kg S/C config. and dry mass w/o payload PL mass, kg ΔV range, ms Δi range, deg 7 Inboard, 6 kg 0.76–3.76 0–2400 0–20 7 Outboard, 3 kg 0.76–3.76 0–9500 0–82 8 Inboard, 6 kg 1.76–4.76 0–4800 0–37 8 Outboard, 3 kg 1.76–4.76 0–8000 0–68 9 Inboard, 6 kg 2.76–5.76 0–5400 0–42 0 Outboard, 3 kg 2.76–5.76 0–7100 0–56

components, radiation shielding, etc., in addition to offering the HiVel CubeSat if spacecraft wet masses of up to 9 kg are considered; opportunity for spacecraft design optimization. It is recognized that however, maximizing the payload minimizes ΔV capabilities and the lower bound of 3 kg without payload is an aggressive dry mass vice versa, as seen in Figs. 8 and 9. Table 3 summarizes the estimate, as suggested by the 3.9 kg dry mass (without payload) of the information shown in Figs. 8 and 9, displaying the maxima and nominal HiVel CubeSat design described earlier. However, a HiVel minima of payload mass, payload volume, ΔV, and Δi for a variety of spacecraft with a 3 kg dry mass without payload may become viable spacecraft bus masses and thruster configurations. A complete over the next few years due to novel weight-saving techniques. These analysis was also done for a 4.5 kg dry mass spacecraft; these results may include the use of carbon fiber for all of the CubeSat exterior are not presented here because, as expected, they fell linearly between body panels or the integrating of multiple subsystem circuit boards the 3 and 6 kg cases. onto a single card [36]. Each of the two assumed spacecraft dry mass datasets were plotted for a range of spacecraft flight masses from 7 to 9 kg. The upper V. Conclusions bound of 9 kg was selected because it represents the approximate The preceding analysis shows that the proposed high-ΔV CubeSat maximum spacecraft flight mass capable of carrying enough design can be considered for demanding missions requiring propellant to reach the lunar surface given the preceding 6 kg dry significant inclination changes, orbit raising, or lunar missions. mass assumption (see Fig. 7). The assumed dry mass lower bound of According to first-order spacecraft design calculations, payload 3 kg was chosen because the original standard for a CubeSat’s density volumes of nearly 1.4 U with a mass of up to 6 kg can be was 1kg∕U (recently increased to 1.33 kg∕U) [13]. The 6 kg dry accommodated on a three-cell (3U) CubeSat platform, depending on mass upper bound shows the limitations of this design parameter: at the spacecraft subsystem layout, component selection, and mission the maximum of 2.5 kg of propellant, the spacecraft is not capable of requirements. The range of ΔV capabilities, which is highly carrying a useful payload given a 9 kg spacecraft flight mass. The dependent on thruster and propellant tank configuration, spans from propellant mass was varied between 0.1 and 2.5 kg. The lower limit under 1000 to over 7000 m∕s for a range of payload volumes, for propellant mass was chosen because it provides the spacecraft payload masses, mission types, and spacecraft configurations. with a similar ΔV capability offered by currently available CubeSat Although the spacecraft is capable of adhering to the overall size cold gas thruster systems. The upper bound of 2.5 kg was selected requirements of a standard 3U CubeSat, while containing all of the because it offers a potential ΔV of over 7000 m∕s for a 9 kg flight required hardware for a high-ΔV mission, the spacecraft will mass, which is sufficient for a lunar mission. invariably exceed the typical 4 kg maximum mass for CubeSats, thus The ΔV and Δi capabilities of a high-ΔV 3U CubeSat as a function requiring a waiver. Because of packaging, available space, and of propellant mass are presented in Figs. 8 and 9 and summarized in thermal view factor, an outboard miniature xenon ion thruster Table 3. Two sets of figures are shown, assuming a spacecraft dry configuration offers superior performance and improved thermal mass of 3 and 6 kg without payload, which represent the lower and control. The miniature xenon ion thruster’s small size allows for its upper bounds for this parametric analysis. As described earlier, the use on a CubeSat, whereas its performance provides the necessary lower bound of 3 kg was selected as a possible target for the HiVel propulsion for a high-ΔV mission. The thruster’s efficiency, even CubeSat if weight-saving measures such as CubeSat structural when conservatively approximated, is sufficiency high that a wide

Downloaded by JET PROPULSION LABORATORY on December 18, 2013 | http://arc.aiaa.org DOI: 10.2514/1.A32435 materials and ultracompact electronics are considered [36]. The variety of spacecraft configurations with significant payload can be capabilities of the spacecraft were plotted over the range of useful developed to satisfy high-ΔV missions; the useful payload capacity is payload volumes and masses available as found in Figs. 6a, 6b, and 7; limited primarily by the propellant tank size, thus providing the if no payload volume or mass was available for a specific spacecraft mission designer with a payload versus ΔV tradeoff. Power and configuration, that particular design was not presented. This is thermal models suggest that the spacecraft can be designed to avoid apparent in Figs. 8a, 9a, and 9b, in which the curves end before the mission-threatening thermal conditions during normal operations, maximum propellant mass is reached; a combination of the required eclipse, or emergency operations. propellant tank dimensions and total allowable flight mass of the spacecraft disallow sufficient interior volume or mass for a useful Acknowledgments payload. The maximum propellant mass available for an inboard MiXI configuration was 2 kg, whereas the maximum for the outboard This investigation was funded in part by a grant from California configuration was 2.5 kg. The 2 kg limit was established based on the Space Grant Consortium. The authors would like to thank Chris inboard configuration because, given the aforementioned sizing Russell and Ryan Caron for their contributions to this effort. assumptions, no useful payload volume was available due to the dimensions of the propellant tank. The outboard configuration yields References a significant increase in available interior volume, allowing for over [1] Wirz, R., Sullivan, R., Przybylowski, J., and Silva, M., “Discharge 2.5 kg of propellant. The plots only extend to this value, however, Hollow Cathode and Extraction Grid Analysis for the MiXI Ion because it represents the approximate required propellant mass for a Thruster,” International Journal of Plasma Science and Engineering, 9 kg CubeSat to reach the lunar surface (ΔV of approximately Vol. 2008, 2008, p. 693825. 7100 m∕s from LEO). Awide range of missions are possible with the doi:10.1155/2008/693825 1046 CONVERSANO AND WIRZ

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