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Small satellite survey mission to the second Earth moon
Available online at www.sciencedirect.com
ScienceDirect Advances in Space Research 52 (2013) 1622–1633 www.elsevier.com/locate/asr
Small satellite survey mission to the second Earth moon P. Pergola ⇑ Aerospace Engineering Department, University of Pisa, Via G. Caruso 8, 56122 Pisa, Italy Received 15 April 2013; received in revised form 24 July 2013; accepted 29 July 2013 Available online 6 August 2013
Abstract This paper presents an innovative space mission devoted to the survey of the small Earth companion asteroid by means of nano platforms. Also known as the second Earth moon, Cruithne, is the target identified for the mission. Both the trajectory to reach the target and a preliminary spacecraft budget are here detailed. The idea is to exploit high efficient ion thrusters to reduce the propellant mass fraction in such a high total impulse mission (of the order of 1e6 Ns). This approach allows for a 100 kg class spacecraft with a very small Earth escape energy (5 km2/s2) to reach the destination in about 320 days. The 31% propellant mass fraction allows for a payload mass fraction of the order of 8% and this is sufficient to embark on such a small spacecraft a couple of nano-satellites deployed once at the target to carry out a complete survey of the asteroid. Two 2U Cubesats are here considered as representative payload, but also other scientific payloads or different platforms might be considered according with the specific mission needs. The small spacecraft used to transfer these to the target guarantees the manoeuvre capabilities during the interplanetary journey, the protection against radiations along the path and the telecommunication relay functions for the data transmission with Earth stations. The approach outlined in the paper offers reliable solutions to the main issues associated with a deep space nano-satellite mission thus allowing the exploitation of distant targets by means of these tiny spacecraft. The study presents an innovative general strategy for the NEO observation and Cruithne is chosen as test bench. This target, however, mainly for its relevant inclination, requires a relatively large propellant mass fraction that can be reduced if low inclination asteroids are of interest. This might increase the payload mass fraction (e.g. additional Cubesats and/or additional scientific payloads on the main bus) for the same 100 kg class mission. Ó 2013 COSPAR. Published by Elsevier Ltd. All rights reserved. Keywords: Low thrust; Interplanetary missions; Cruithne asteroid; Cubesat; Small satellites
1. A small satellite observation mission During the last 25 years a dozen of space missions visited small bodies in our solar system. During its journey towards Jupiter, in October 1991, the Galileo spacecraft passed at 1.600 km from 951 Gaspra at a relative speed of about 8 km/s, making this asteroid the first one ever observed by a spacecraft (Belton et al., 1992). Seven years later, in 1998, 433 Eros was visited by the NEAR Shoemaker probe (Veverka et al., 2001). This probe first had a flyby with the asteroid and 2 years later, in 2000, started orbiting the same body, making 433 Eros the first asteroid ever orbited by a spacecraft. The same probe visited also ⇑ Tel.: +39 050967211.
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another asteroid 253 Mathilde in the 1997 and the Galileo spacecraft itself carried out another asteroid flyby during its path with the 243 Ida in 1993. In 1991 Deep Space 1 was the first probe to test in flight an ion thruster as main propulsive system. This spacecraft reached the asteroid 9969 Braille in 1999 and the mission was extended twice to include also the encounter with Comet Borrelly in 2001 (Rayman et al., 2000). Several interplanetary missions visited, typically with some distant fly-by, an asteroid as secondary scientific target. Beside the already mentioned Galileo, also Cassini, that visited the asteroid 2685 Masursky in 2000 (Cuzzi et al., 2003), and New Horizon, that visited the asteroid 132524 APL in 2006 (Weaver et al., 2008), are worth to be mentioned. The Stardust mission, where the primary targets were to collect samples of cosmic dust and of the dust of the Wild
0273-1177/$36.00 Ó 2013 COSPAR. Published by Elsevier Ltd. All rights reserved. http://dx.doi.org/10.1016/j.asr.2013.07.043
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2 comet coma, also visited the asteroid 5535 Annefrank on November 2002 at a distance of 3079 km (Duxbury et al., 2004). The first attempt to return asteroid samples to Earth was made by the Hayabusa spacecraft. This probe, equipped with ion thrusters, visited the asteroid 25143 Itokawa in 2005. It carried out a soft landing on the body and collected some samples that were brought back to Earth by means of a re-entry capsule (Yoshikawa et al., 2007). On their way toward their targets there are currently two probes that have already visited some asteroids as side scientific targets. The Rosetta spacecraft (Lamy et al., 2003) is on its way to reach in the 2014 the comet 67P/Churyumov– Gerasimenko. In the September 2008, after two Earth and one Mars swing-bys, the probe visited for 7 min the asteroid 2867 Sˇteins at a distance of 800 km and a relatively slow speed of 8.6 km/s (Accomazzo et al., 2010). Two years later, in July 2010, Rosetta visited also the asteroid 21 Lutetia (a large main belt asteroid) after another Earth swing-by and before entering in hibernation mode (Lamy et al., 2003). The Dawn spacecraft, equipped with three Xenon-fed ion thrusters, is another low thrust mission for the exploration of Vesta and Ceres, the two largest bodies of the main asteroid belt (Russell et al., 2007) (Taylor, 2009). Vesta has been already visited by Dawn from July 2011 to September 2012 and Ceres is scheduled to be reached in February 2015 (Russell et al., 2007). All the probes considered so far reached some asteroids, although in many cases these bodies were not the primary mission targets and the asteroid analysis was carried out only by means of short or distant fly-bys. These probes range from 300 kg up to 1200 kg for the dedicated missions (up to 2500 kg if considering also probes dedicated to planetary observations, like Galileo and Cassini). The great majority of these missions used some complex interplanetary trajectories, with multiple planetary swing-bys and thus long transfer times, and were aided by very powerful launches. New Horizon, for instance, the first spacecraft launched directly into a solar escape trajectory, was launched with a hyperbolic excess energy (C3) of about 170 km2/s2 (Weaver et al., 2008). The aim of the study is to present another option for the exploration of an asteroid. A dedicated mission is considered where the target is intended to be investigated in great detail. With the aim at reducing the mission cost and complexity the paper aims at demonstrating the feasibility of such a dedicated asteroid orbiting mission with a 100 kg class spacecraft launched with a small hyperbolic excess velocity, without any planetary swing-by along its path (thus reducing the transfer times and maintaining a high launch window flexibility) and powered by an electric thruster. In addition, the mission presented considers also the possibility of a detailed exploration of the target by means of multiple nano-satellites operating once at the target. The asteroid considered is a rather challenging target and accordingly the propellant mass fraction here identified can be significantly reduced in case of different targets
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allowing for an even smaller launch mass or a larger payload mass fraction. In this perspective the mission proposed has to be intended as based on a set of conservative assumptions. Cruithne is chosen as the scientific target of the small survey mission proposed. As remnant brick of the solar system formation, for its near Earth orbit and for its peculiar orbital resonance, this asteroid might provide a significant scientific outcome. It is worth mentioning that also other studies (Wells et al., 2006; Giovannetti et al., 2011; PROBA-IP, 2009) proposed small asteroid survey missions, but these are different from the one detailed in the paper for the specific target chosen, the kind of payload considered and thus the exploration phase, the electric propulsion approach and the order of magnitude of the envisaged spacecraft mass. By exploiting the high thrust efficiency of electric thrusters and the current miniaturization trend of space electronic devices, it is proven that Cruithne can be reached by a 100 kg class spacecraft. Once at the target the main spacecraft bus acts mainly as relay system for telecommunication of the two nano-satellites deployed; however, a few kg of additional scientific payload can be also added to the main probe to complete the asteroid science phase. This would result in a larger launch mass thus in longer transfer times and additional propellant mass as detailed in Section 3.2. The spacecraft deploys a couple of 2U Cubesats to obtain a multiple observation of the asteroid surface and interior. These nano-satellites can operate in different wavelengths and/or be equipped with charged particles probes, mass spectrometers, accelerometers and so onto obtain a complete picture of the whole body. The aim of the paper is to demonstrate the feasibility of this innovative mission concept rather than a specific sizing of these nano-platforms. With such approach the complete asteroid exploration can be carried out by means of a single small scale mission. In the paper the mission design phase (see Section 3) is detailed to estimate the mission Dv and the propellant mass fraction required for the subsequent phase of preliminary spacecraft sizing (see Section 4). All the main subsystems are sized by using COTS and state of the art figures to provide a conservative design. 2. Cruithne Also known with the official designation of asteroid 3753, or 1986 TO, Cruithne is a 5 km wide Near Earth Asteroid (NEO) (Namouni et al., 1999). The body is locked in a 1:1 mean motion resonance with the Earth, thus the two bodies appear to follow each other on their heliocentric paths. For this reason Cruhitne is also called the second Earth moon although it is not directly affected by the Earth gravity field. Despite this orbital resonance, at the closest approach, which happens every 285 years, Cruithne comes only within 12.5e6 km, sufficiently far to be not considered as a possible impactor, but enough close to have its orbit perturbed by the Earth (Brasser et al., 2004).
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As seen from Earth, Cruithne follows a bean-shaped horseshoe orbit that lasts 364 days (Namouni et al., 1999). Since its period is slightly shorter than 1 year, the Earth remains behind the bean-shaped orbit a little more each year. As a consequence from our point of view the asteroid orbit is not really closed, but it looks more like a spiral loop moving away from Earth (Namouni et al., 1999). As seen in a heliocentric frame, the asteroid moves in a very elongated elliptic path (crosses the Mars orbit and glazes the Mercury one) with 0.4840 AU perihelion, 1.511 AU aphelion and 19.8° of inclination. This last aspect, in particular, makes the mission design rather challenging and requiring a total impulse significantly larger than any almost equatorial NEO. The asteroid orbit as seen in the inertial heliocentric frame is shown from three perspectives in Fig. 1. To obtain the trajectory shape in a reference frame rotating with Earth, the instantaneous relative position of the two bodies is required (Fig. 2, left). These distance vectors are then translated into the Earth position at a date (Fig. 2, centre) and their envelope finally gives the beanshape orbit in magenta in the rightmost plot of Fig. 2. Changing date, the Earth position changes and the apparent asteroid orbit moves describing a three dimensional horse-shoe orbit covering a torus around Earth’s orbit. The bean-shape orbit completes such a torus each 770 years (Brasser et al., 2004).
3. Trajectory design With the aim at reducing the fuel mass fraction, thus allowing a small spacecraft (100 kg class) to perform an asteroid observation mission, an electric propulsion transfer is considered. The path is designed in the heliocentric two body system minimizing the transfer time. This approach has been preferred with respect to a minimum mass transfer to limit the ground segment and to avoid that an extended permanence in space might jeopardize the two nano-spacecraft (or, in general, the payload). The main challenge associated with the transfer sought is the relatively high inclination of Cruithne. This forces the electric transfer to significantly change both the velocity in the ecliptic plane and the velocity perpendicular to that plane. Still following the idea to present a relatively cheap mission, the spacecraft hyperbolic excess energy has been arbitrarily limited to 15 km2/s2. A 100 kg class spacecraft with such C3 and with reduced dimensions (less than 2 m, see Section 4) is compatible with the great majority of the small/medium launcher fairings, even as piggyback payload. The mission analysis presented in Section 3.2 has been refined after the spacecraft design detailed in Section 4. An initial mass of 100 kg has been used and the spacecraft has been considered equipped with two twin T5-like ion thrusters.1 This thruster is used only for providing some 1 http://earth.esa.int/workshops/goce04/participants/101/ poster_T5_Ion_Propulsion_A0.pdf
reference values and any other thruster can be considered changing the spacecraft characteristics and accordingly the transfer time and propellant mass consumption. This specific ion thruster is actually able to provide a thrust between 0.6 and 20.6 mN with a specific impulse reaching up 3500 s and with a power requirement between 55 and 585 W across the thrust range (Edwards et al., 2004). 3200 s of specific impulse and 18 mN thrust were used as nominal values (Goebel and Katz, 2008) and two thrusters have been assumed firing in parallel for the whole transfer. With the aim at outlining a small and affordable mission, no cold redundancy on the thruster has been considered. This means that if one of the thrusters fails the whole mission might still be accomplished with significantly longer transfer times and a new trajectory design. 3.1. Approach The indirect optimization approach here used is well known to be rather sensitive to the initial conditions supplied (Bryson and Ho, 1975). To limit such sensitiveness a preliminary, more robust gradient optimization scheme has been implemented. This step is not required to be convergent and just few iterations, typically less than 10, are sufficient to obtain an initial guess for the continuous and discrete Lagrange multipliers. Starting from these values an indirect optimization scheme has been implemented. The method integrates together the equations of motion and the adjoint equations (Betts, 1998). To increase the numerical stability of the numerical routine and to avoid any singularity associated with the classical orbital elements, the equinoctial elements have been used. The calculus of variation problem is posed in the Mayer form, where the transfer time is the fnctional to be minimized, and the control vector is composed by the continuous and discrete Lagrange multipliers parameterizing the in-plane and out-of-plane thrust angels. The iterative scheme aims at obtaining a stationary value of the functional (Bryson and Ho, 1975). The boundary conditions to be satisfied derive from imposing that the final state of the propulsive phase belongs to the asteroid orbit (Betts, 1998). Together with these boundary constraints also the transversality condition arising from the indirect problem formulation has to be satisfied (Betts, 1998). The anomaly of the encounter point is not fixed in advance but results from the optimization scheme as well as the total transfer time. Since, to reduce the transfer time, coasting arcs were not allowed, from the transfer time the propellant mass consumption can be directly computed and from this the Tsiolkowsky equations allows for computing the mission total Dv. The initial spacecraft velocity is modelled as the Earth heliocentric velocity augmented of some energy assumed to be delivered by the launcher. This initial Dv is always considered as tangential to the Earth heliocentric velocity. A grid search on the departing energy has been carried out considering escape C3 energies between 1 and 15 km2/s2.
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Fig. 1. 3753 Cruithne orbit in three views in the heliocentric inertial frame.
Fig. 2. Construction of the Cruithne orbit relative to Earth. The bean-shape orbit in the rightmost plot represents the envelope of the relative EarthCruithne position vectors over one revolution centred in the Earth position at a date.
The iterations have been stopped when a tolerance at least of the order of 1e 6 was reached on the accuracy of the final constraints. The equations of motions have been numerically integrated in non dimensional units using a variable step Runge–Kutta 4–5 order integrator with absolute and relative tolerance of the order of 1e 9.
3.2. Results The resulting trajectory implements a direct transfer that lasts less than one year and reaches the asteroid orbit close to its aphelion (the slowest point). The thrust law is characterised by an in-plane thrust angle aiming at increasing the spacecraft velocity and an out-of-plane angle direct constantly toward the asteroid aphelion (negative z semispace) to modify the orbital inclination. The resulting transfer trajectory is shown in the heliocentric inertial frame in Fig. 3. After the gradient first guess generator the forward shooting approach typically converges in less than 500 iterations with a total CPU time of less than 8 min on a 2.2 GHz laptop with 4 GB ram. The transfer lasts 322 days are requires 31.9 kg of propellant mass. The mission total Dv is about 12.01 km/s and the total impulse is about 1.01e6 Ns. The total impulse
a single T5 is able to deliver is of the order of 1.5e6 Ns2 and up to 8500 on/off cycles can be handled (not stringent limit in the planned mission as coasting arcs are not allowed and only few eclipses have to be considered), thus the couple of thrusters should already offer a safe design for the planned mission. The T5 total impulse, however, is the actual limit to take into account in case of mission re-planning for a single thruster failure. The grid search on the departing energy results in a rather flat behaviour in terms of transfer time and propellant mass consumption. This is mainly due to the high orbital inclination of the target that requires some time to be reached (the initial velocity is always supplied tangential to the Earth velocity and thus in the ecliptic plane). This causes also the transfers characterised by a high C3 to require a longer time (more mass) since the initial velocity in the ecliptic plane does not allow reaching the Cruithne orbit in a convenient point. The minimum time (that is also minimum mass as no coasting arcs are considered) trajectory, see Fig. 3, among the results obtained has a C3 energy of 5 km2/s2.
2 http://earth.esa.int/workshops/goce04/participants/101/ poster_T5_Ion_Propulsion_A0.pdf
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Fig. 3. Electric transfer trajectory from Earth to Cruithne orbit as seen in the heliocentric inertial frame.
As the main aim of the study is to present an innovative mission scenario making use of nano platforms to explore NEO objects, the specific mission might differ from the case here presented in particular for the launch mass. Accordingly a sensitive analysis with respect to a variation of the launch mass, from 100 to 150 kg has been carried out. Fig. 4 shows the variation of the transfer time and propellant mass as a function of the launch mass. All the cases have been optimised with the approach described and the launch C3 has been adapted to allow for the convergence of the numerical procedure. In particular up to 110 kg of launch mass 5 km2/s2 proved to be sufficient; from 110 to 130 kg, instead, 10 km2/s2 were considered and from 130 to 150 kg a C3 of 15 km2/s2 was used. It has been observed in this analysis the encounter point with Cruithne orbit moves toward its perihelion. This is a consequence of the lower acceleration level provided by the thruster that
requires a longer time to match Cruithne orbital parameters. From this launch mass sensitivity analysis it results that the propellant mass fraction remains between the 28% and the 32% of the initial mass, thus leaving in all cases about the 70% for the spacecraft systems and the payload. In addition, up to approximately 120 kg of launch mass, the transfer takes less than one year. To provide an impulsive comparison the same transfer has been assessed by means of a rough grid search over one year of Earth departure dates and one year of transfer times. The Lambert problem has been solved for this grid of data and it results that the same target can be reached with a standard bi-impulsive manoeuvre in approximately 8 months and with 17 km/s of total Dv. This value, in particular, is comparable with the electric transfer Dv as in this case the significant inclination change has to be performed
Fig. 4. Transfer time and propellant mass variation with respect to a variation in the launch mass.
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Fig. 5. The spacecraft during the operative phase with the ion thrusters operating, one Cubesat fully deployed and the other one ready to be delivered.
anyway and we cannot take advantage of the negligible gravity field of the target that might have reduced the magnitude of the second impulse. The 84% of the total chemical Dv has to be provided at the departure and it realises the largest part of the inclination change manoeuvre. This Dv might be provided by the launcher, although it is significantly larger than the 5 km2/s2 C3 used for the low thrust analysis.
4. Spacecraft layout To contain dimensions and to better exploit the internal volume, the spacecraft is conceived as an octagonal prism, like the Lisa Pathfinder bus (McNamara and Racca, 2009). The upper surface of the bus is reserved for solar arrays and the bottom surface for the propulsion system. The lateral surfaces are used for the telecommunication system, the radiators and for the Cubesat deploying system. Fig. 5 presents a conceptual view of the probe during the operative phase with the two ion thrusters active (for instance for orbit maintenance purposes), one Cubesat already fully deployed and the other one still in its bay. The main bus structure assures the mechanical support to the spacecraft subsystem and to the deploying system of the two 2U Cubesats. It is an octagonal prism with an outer diameter of 1.4 m (1.6 m is the diameter of the upper face) and 0.7 m height. The upper face is completely covered by a sunshield panel supporting the body mounted solar array. To produce the 1140 W required to operate the two electric thrusters and to satisfy the spacecraft power needs (200 W assumed, see Section 4.1), two additional solar arrays half the size of the body mounted one are deployed after the Earth escape. This 8 m2 panel produces 1430 W End Of Life (EOL) at 1.511 AU distance from the Sun (Cruithne aphelion). The internal spacecraft structure is composed by a large central cylinder from which a set of transversal shear walls connects the inner cylinder to the outer panels, see Fig. 6. The whole structure is assumed to be constructed of sandwich panels with carbon fibre laminate skins bonded to aluminium honeycomb core. The internal cylinder accommodates the ATK 80194-1 18.8 L Xenon tank. This tank has been chosen after a tradeoff between required propellant and available volume. The tank itself has a mass of 7.67 kg, is of spherical shape
with a radius of 340 mm and operates at a pressure of 250 bar.3 At such pressure, and considering a storage temperature around 310 K, approximately 17.8 L are required to store 35 kg of propellant. Of these almost 32 kg are reserved for the transfer and 3 are considered for the attitude control system and for eventual additional contingency. The two ion thrusters are located on the bottom plate of the bus. These are mounted with a small, 7°, angle to have the thrust direction passing through the spacecraft centre of mass. A cylindrical shell is placed all around the thrusters to limit the plume interaction with the other spacecraft surfaces. Each thruster weights about 1.7 kg and has a cylindrical shape 17 25 mm. The control electronics of the thruster is placed in one of the internal bus bays. It is composed by two main units. The proportional Xenon feed system includes the pressure control unit and the flow control section (bang bang pressure regulation valve and Xenon flow control unit). The module receives Xenon directly from the tank (5–250 bar) and regulates it down to approximately 2.5 bar. The module includes also particle filters, pressure transducers and insulation valves (Wallace et al., 2011). The ion propulsion control unit considered is a highly integrated module. It includes the control electronics (MIL-1553 interfaces and timing synchronisation with the spacecraft OBC), the DC-AC converter with low voltage (LV) and high voltage (HV) output, the converter from spacecraft DC to the HV DC source required for the ion beam, LV and HV output control for internal function and for the power distribution to the other units. The control module runs also on an independent micro processor (like the ERC 324) to implement the thrust control algorithm of the two ion thrusters. This unit is estimated to weight approximately 12 kg, requires up to 20 W and has estimate dimensions of 450 200 200. The thermal control system must guarantee a very stable thermal environment during the whole transfer and at destination. During the path the Sun heating is shielded by the thermal shield and the solar array mounted on the upper surface. Two radiators are mounted on two sides of the 3
http://www.psi-pci.com/Data_Sheets_Library/DS194.pdf http://microelectronics.esa.int/erc32/chipset/Hardware%20and% 20Documentation%20Status%20of%20the%20ERC32%203-Chipset% 20i2r1a.pdf 4
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Fig. 6. A global view of the complete spacecraft with deployed arrays (upper) and two views of the internal spacecraft structures where some internal walls have been hidden to show the internal subsystems.
spacecraft bus for a total surface of 1.4 m2 (total emissivity 0.8, solar absorbance 0.21). The whole external structure is folded with a 10 layers Multi Layer Insulator (MLI) (large conductance and emissivity, 0.4 kg/m3 density), while the inner structures are covered with a 20 layer MLI (high radiation protection, 0.6 kg/m3 density). A nominal operation temperature of 290 K is considered with a maximum admitted variation of 10°. The main internal heat sources considered are the two ion thrusters; for these a thrust efficiency of 55% is considered. This means that the 45% of the incoming power is transformed into heat to be dissipated. From experimental tests, however, it is known that this heat flow is mainly radiated away with the plume and only the 40% is the heat directly transferred to the structure (about 170 W in total). Active thermal control system elements are considered for the most critical elements like electronic modules and the tank. As general philosophy, it is assumed to provide the minimum necessary heater power in the cold case scenario so that the lower temperature of each unit is maintained near the bottom of the allowable range. Depending on the science phase duration, the most critical situation (hot case) is when Cruithne is at its perihelion. In this situation rather high temperatures can be reached and it is assumed to maintain the spacecraft attitude to have sun pointing solar panels, which offer also the proper thermal shielding. The spacecraft is a three-axis controlled spacecraft. The reaction wheels are desaturated by means of the two ion
thrusters and a set of 6 small (50 g each) Xenon-fed 10 W resistojets mounted on the structure corners. 3 kg of additional propellant are embarked for this purpose. The vertical wheel is placed on top of the Xenon tank, the horizontal one on one side of the same and the other wheel along the spacecraft principal axis, together with the spare (inclined) one and the associated electronics on a bus bay, see Fig. 6. Each wheel weights about 0.9 kg (e.g. RW 905). The attitude and orbit control system is completed with one Earth horizon sensor, two autonomous star trackers and two digital Sun sensors for the Sun acquisition (McNamara and Racca, 2009). These sensors are mounted on the bottom bus plate with the proper orientation (star trackers are mounted where there is the radiator and below one of the two solar array wings, a spacecraft side that should be constantly pointing the deep space), see Fig. 7. During the transfer the attitude is constrained by the thrust steering law, thus it may happen that the solar panels are not perfectly Sun pointing. Actually an incidence Sun light angle up to 40° can be reached, thus the solar array surface has been properly oversized to take into account this situation. This option has been preferred to a Solar Array Driver Assembly (SADA) system as it is lightweight, cheap and more affordable for a small mission.
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http://www.astrofein.com/2728/dwnld/admin/Datenblatt_RW90.pdf
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Fig. 7. Internal views of the spacecraft from the top (left) and form the bottom (right) the upper solar panel and the bottom covering surfaces have been hidden to show the interior structure and components.
The power subsystem guarantees a stable regulated bus with a constant voltage of 28 V. Because of its low mass and simple management, a pack of three Li-ion batteries has been chosen as energy storage subsystem. The battery pack provides a total of 300 Wh considered for the early mission phases before the two wings are flipped, for the short eclipses and for the peak power mode requests. This battery pack is placed in a spacecraft bay, see Fig. 7, and it is equipped with the proper heaters. Two of these bays are protected against radiations with a 1 mm aluminium shield. In these bays are placed all the sensitive electronic components and the battery pack. These two bays are also the two directly behind the radiators to assure an easier thermal path to control the electronic component temperature, see Fig. 6 and 7. The On-Board Computer (OBC) is in charge of handling the on-board data, the attitude and orbit control system, the management of the platform and the Cubesat deploying system once at destination. Furthermore the control system performs also monitoring functions of the main systems and provides safe system reconfiguration capabilities. The OBC is composed of two processing modules of 33 MHz each (e.g. RAD60006) and a redundant solid state 100 GB memory device. The OBC interface with the spacecraft is assumed to be based on a standard MIL-bus 1553B, while the internal backplane link uses the Space-Wire standard. Dedicate modules for telecommand, telemetry, reconfiguration and for actuator and sensor interfaces are also considered. The OBC primary computer is estimated to weight about 1.5 kg7 and 0.8 kg are considered for the data recorder unit and cabling. 6
http://atc2.aut.uah.es/mprieto/asignaturas/satelites/pdf/rad6000.pdf http://www.sst-us.com/shop/satellite-subsystems/obdh/radiation-tolerant-flight-computer-obc-695
On a side bay of the spacecraft also the Cubesat deploying structure is placed. The two access gates are directly obtained in the main structure wall. The two deployers are placed on top of each other mounted on a tubular structure. The command electronics and all the security/ check systems to monitor the Cubesat health during the transfer are placed on the back of the same structure. Each deployer weights about 1.75 kg and sizes 300 130 180 mm.8 The communication system operates in X-band with an uplink frequency of 7230 MHz and 8495 MHz downlink. The system provides both for commanding and housekeeping telemetry of the main spacecraft and acts as relay system for the telemetry of the Cubesats. This system, indeed, receives the data from the 2U micro satellites and sends these to the ground station. The communication frequency with the payload is assumed to be in S-band. The spacecraft is equipped with a 1.2 m diameter high gain antenna (slightly smaller than the antenna of other deep space missions, like Hayabusa and Mars Global Surveyor) mounted on one side of the spacecraft bus (Taylor, 2009) and two low gain omni-directional antennas mounted in opposite directions to allow for the communications with the Earth station during the early transfer phases and when the main antenna in not Earth pointing, see Figs. 6 and 7. During the operation phase, instead, only the main antenna is in charge of transmitting data to Earth and the two omni-directional antennas (easier to communicate with a moving spacecraft constantly changing aspect angle) are reserved for the bi-directional communications with the Cubesats. The maximum datarate assumed is of 60 ksymbols/s for the low gain antennas and 120 ksymbols/s for
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8
http://www.isispace.nl/brochures/ISIS_ISIPOD_Brochure_v.7.11.pdf
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the high gain one (McNamara and Racca, 2009). Two Xband transponders are included in the telecommunication system dedicated electronic (placed in the bus bay behind the main antenna) for redundancy and an additional amplification module is included at the transponder output (Taylor, 2009). The telecommunication system is equipped with a dedicated solid state memory of 60 GB for the science data following a “store and forward” strategy. The high gain antenna is estimated to weight 3 kg, the low gain antenna 0.2 kg each and the telecommunication dedicated electronic (X-band solid state amplifier and solid stage data recorder) 1.3 kg. Based on the results of the trajectory analysis in the scenario considered, almost 122 days of visibility are guaranteed after the probe arrival. This means that about 4 months of direct communications are allowed before that the line-of-sight between the spacecraft and the Earth passes beyond Mercury orbit where the intense Sun radiation might affect the data relay. During this period the Earth-spacecraft distance varies between 2.14 (at the arrival point close to the Cruithne aphelion) and 0.3 AU with an associated communication delay between 2.5 and 17.8 min. The minimum Earth distance is reached 76 days after the rendezvous. To assess the telemetry/telecommand and tracking signal availability and quality during the mission phases, a preliminary link budget has been performed. Assuming to use the Deep Space Network antennas as ground receivers and maintaining a signal/noise margin of at least 2 dB, it follows that at the closest point the whole 120 ksymbols/s data rate can be used (even higher rates are allowed), while in the farthest point the maximum data rate allowed for the downlink is 7 ksymbols/s. As a consequence, assuming 8 h of ground station visibility per day, from 0.2 up to 3.2 GB per day can be downloaded. The remaining data (depending on the specific mission application) have to be stored in the 60 GB solid state memory and forwarded at some closer point. 4.1. Mass and power budgets The structure mass is estimated according with state of the art figures as the 10% of the launch mass. This mass includes the carbon fibre honeycomb bus structure, the support plates for the subsystems and the 1 mm shield for the two bays where the more critical components are placed. The propulsion system mass is about 25 kg including the two thrusters, the tank and the main electronic units. The power condition unit and the associated modules are estimated to weight around 12 kg (Tato and de la Cruz, 2007), 1 kg for the bang-bang pressure regulation unit and 1 kg for the Xenon flow control unit together with the pre-regulation valve. To deliver the required 1140 W EOL at Cruithne aphelion, almost 6.4 m2 solar arrays are required. To consider also some cell aging in space, possible smaller solar cell
efficiency, possible peak power requirements and the nonperpendicular Sun incidence angle, a total of 8 m2 has been covered with solar cells. 30% solar cell efficiency, a powerto-mass ratio of 200 W/kg (Bailey and Raffaelle, 2011) (Surampudi, 2011) for the solar arrays and a battery power density of 300 Wh/kg (Surampudi, 2011) have been considered. With these assumptions made the power generation system mass is about 5.7 kg. The battery pack, instead weights almost 1 kg assuming to size this to sustain only the bus needs (no firing in eclipses). The attitude control system weights a total of 5.2 kg where the largest contribution is due to the reaction wheels (3.6 kg). 0.5 kg are considered for the resistojet thrusters and the associated piping. The system electronics is estimated to weight only 0.25 kg (as it mainly acts as data buffer with the OBC and local control unit), the two Sun sensors weight a total of less than 200 g, 0.5 kg the Earth horizon sensor and 0.25 kg each the two star trackers. The telecommunication system weights some 4.9 kg where the largest contribution, 61%, is due to the high gain antenna. The thermal system is estimated for a total mass of 3.2 kg. The MLI used for the bus surfaces, see Section 4, is estimated to weigh about 2 kg and the 1.4 m2 radiator surface is considered weighting 1.2 kg (Wertz et al., 2011). The mass budget is summarized in Table 1. Here 5% of additional contingency is considered on each subsystem mass. The total spacecraft mass, including also the propellant and the contingency is 103 kg and 68 kg is the system dry mass. In Table 1 the mass budget presented is also compared with some state of the art figures of the various subsystems (Wertz et al., 2011). This last comparison confirms that the assumptions and the analyses made are coherent with reference values. A couple of observations are mandatory. The propulsion system for the spacecraft designed is significantly heavier than the reference value; this because the statistical analysis used to derive these state of the art figures is based only on chemical thrusters (Wertz et al., 2011). In addition, in the mass budget of this system also the power management mass is included, thus the power generation system mass (where usually the power management equipment is included in case of chemical spacecraft) is smaller than the reference value The structure considered is slightly lighter than a standard one taking into account the different design philosophy (carbon fibre honeycomb and internal sets) with respect to the spacecraft used for the statistical analysis. A static and dynamic detailed sizing of the structure is however required, and it is left for a more detailed study. The payload includes the two deployers, the two 2U Cubestas (2 kg each) and 0.5 kg of harness and support. This figure is slightly smaller than a standard payload figure for a planetary mission. Nonetheless the deployable payload represents an innovative mission aspect, offering a complete covering of the target, which might be replaced or associated with some other scientific equipment.
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Table 1 Spacecraft mass budget. The estimated mass, the mass including 5% contingency, the initial mass percentage and a comparison with state of the art figures is given for each subsystem. Subsystem
Mass (kg)
Total mass with 5% contingency (kg)
Percentage of dry mass (%)
Reference value (%)
Structure Propulsion system + Tank Xenon propellant Payload Power generation system + energy storage system Attitude and orbit control system Command and data handling system Thermal control system Telecommunication system Launch mass Dry mass
10.0 25.0 35.0 8.0 6.7
10.5 26.2
15.4 38.3
20 13
8.40 7.0
12.3 10.2
15 21
5.2 2.4 3.2 4.9 100.4 65.4
5.4 2.5 3.3 5.1 103.4 68,4
7.9 3.7 4.8 7.5
6 4 6 7
As other payload options several standard instruments can be considered. Maintaining the same payload mass of 8 kg (although also the whole spacecraft structure might be different for accommodating specific payloads) other payload options might be a Laser Imaging Detection and Ranging (LIDAR) (3.7 kg, 17 W) allowing for high resolution surface mapping, geology and geomorphology (Takagi, 2009); a Remote X-ray Fluorescence (XRF) Spectrometer (3 kg, 15 W) to determine the surface composition (possibly also of lighter elements) and processes (Takagi, 2009); an Infrared (IR) camera (3,5 kg, 29 W) for different wavelength analysis (Takagi, 2009) and/or IR-spectrometer (e.g. a Single Ion Recording type microspectrometer for moderate spatial resolutions) to be used for asteroid classification; a close-up camera and panoramic stereo camera (1,5 kg, 11 W, total) for high-resolution imaging of surface materials and stereo mapping/ characterization of possible landing or sampling sites (Nathues et al., 2010); an optical and electron microscope (1 kg, 8 W, total) for the study of geophysical and structural properties of the body and the micron-scale characterization of surface materials together with their elemental/mineralogical analysis (Nathues et al., 2010). Moreover a payload package including several of the SIMONE payload equipment might be considered (13 kg, total). The list is not exhaustive and the specific payload combination, up to the payload mass fraction identified in the scenario, depends on the mission scientific targets. From Table 1, it results that the whole spacecraft launch mass is slightly larger than 100 kg. A refined mission analysis with this update value shows that the same transfer can be realised with almost the same transfer time (a couple of days longer) and with an additional propellant mass requirement that is well below the propellant mass contingency considered. The bus power requirement is estimated to be of the order of 200 W. When operating the two thrusters require 2 470 W, thus the total power load in the electric propulsion transfer mode is of 1140 W. Once at target the main power consumption is represented by the high gain antenna
during the Earth downlink/uplink mode. During the launch and the early operative phases the solar panels are still folded, thus the entire power load is managed by the batteries. After this phase, however, only short eclipses and limited power peak requirements are foreseen. The power required by each subsystem is estimated using some standard figures (Wertz et al., 2011). The thermal control subsystem might require up to 15% of the total incoming power, i.e. 16.5 W that can be delivered when thrusters are not operating. This peak consumption might be considered in a very cold or very hot situation. The transfer ranges between 1 and 1.5 AU (not particularly critical), while the science phase can reach up to 0.5 AU, but in this case the main thrusters are not operating and there is are hundreds of W available. During normal operations only 50 W are estimated to be required to maintain the temperature in the allowable range (and, in particular, to avoid Xenon freezing and for battery heaters). The power management system itself requires almost the 10% of power to operate, corresponding to 11.4 W. The telecommunication subsystem has very high power consumption only during the Earth downlink/uplink mode. This power consumption can reach up the 18% of the total power budget. These 205 W (on which the link budget has been estimated), however, are supposed to be required only for short transfer arcs (the telemetry and the steering laws are mainly autonomous) and once at the target where the thrusters are not continuously operating. The OBC can reach up the 11%, 125 W, and almost the same power can be required by the attitude control system. As a consequence of this rough power budget there are two main situations that can take place. Or the thrusters are operating and the remaining 200 W can be used to run almost all subsystem, even in peak consumption conditions, although not at the same time (if peak consumption is required, but of course all can be run also in parallel in normal operation modes). Once at the target, however, the thrusters are not required to run continuously, thus all the main subsystem, also in peak power mode can be used together.
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5. Conclusions In this paper a complete mission for the observation of the second Earth moon is presented. The mission is conceived as a survey mission where the main spacecraft together with two nano-satellites can explore the asteroid Cruithne. The paper presents the complete study of such a mission with a first phase of trajectory design and starting from this also a first order spacecraft sizing and design is presented. The main spacecraft is in charge of the Earth-asteroid transfer by means of a couple of ion thrusters. The mission analysis by means of an optimal minimum time approach results in approximately 32 kg of Xenon and less than 11 months of transfer time. The results obtained from this preliminary mission analysis and the subsequent spacecraft sizing are however to be intended as reference values to demonstrate the feasibility of such a mission. Also slightly changing the system mass values and/or the mission analysis approach a 100 kg class spacecraft can still be able to accomplish the mission proposed. The mission analysis output, in particular, can be also scaled to heavier/lighter spacecraft where the thruster specific impulse, the Earth departure C3 and the electric thruster acceleration maintain almost the same value. In situ observation of small bodies, like asteroids, has a very high scientific outcome as their small size did not allow for the modification of the early solar system chemistry by high internal pressures and temperatures. The mission outlined in the paper plans for the multiple body observation by means of two 2U Cubesats (additional observations/analyses are also possible from the main spacecraft). This approach solves several typical issues of the robotic asteroid observation, for instance the typical communication delay and low bit rate from Earth base stations is solved by using the main spacecraft for the telemetry. The two nano-probes can also operate on different orbits and with different payloads. These Cubesats are required to have a bi-directional telemetry with the mother spacecraft. The specific sizing of such nano-satellites is highly dependent of the required analysis (and the constant improvement in the miniaturization of the Cubesat equipment), however, as rough estimation, almost 30% of the Cubesat mass might be allocated for the payload. Multispectral imaging devices, accelerometers, plasma probes and spectroscopy reflector are some of the typical payloads to consider. In addition, the possibility to exploit multiple sampling locations at the same time might also help to design a dynamical picture of the asteroid environment. In nominal conditions in a Low Earth Orbit (LEO) a typical single 2U Cubesat solar panel can generate up to 5 W; this reduces to approximately 2 W considering the Cruithne aphelium and some aging effects. 2 W are slightly less than the same power generated by a single 1U solar panel in LEO. This means that the on-board systems that might be operated on one of such nano platforms can have
power requirements similar to a state of the art 1U Cubesat in LEO. In addition, as shown in Fig. 5, a valuable option to increase the available power of these 2U platforms might be the usage of two solar array wings (e.g. like the ones already used in Defi-C3 and QbX). In this scenario the available power would be some 4 W for each face. The mission conceived is a valuable approach for low cost, fast analyses precursor missions. The approach, indeed, might act as pathfinder for subsequent heavy robotic exploration/landing missions and in a longer run also for human explorations and asteroid mining missions. References Accomazzo, A., Wirth, K.R., Lodiot, S., et al. The flyby of Rosetta at asteroid Sˇteins – mission and science operations. Planetary and Space Science 58 (9), 1058–1065, 2010. Bailey, S., Raffaelle, R. Space Solar Cells and Arrays, in: Luque, A., Hegedus, S. (Eds.), Handbook of Photovoltaic Science and Engineering. John Wiley & Sons, Chichester, West Sussex, UK, pp. 365–401, 2011. Belton, M.J.S., Veverka, J., Thomas, P., et al. Galileo encounter with 951 Gaspra: First pictures of an asteroid. Science 257 (5077), 1647–1652, 1992. Betts, J. Survey of Numerical Methods for Trajectory Optimization. Journal of Guidance, Control and Dynamics 21 (2), 193–207, 1998. Brasser, R., Innanen, K.A., Connors, M., et al. Transient co-orbital asteroids. Icarus 171 (1), 102–109, 2004. Bryson, A.E., Ho, Y.C. Applied Optimal Control. Hemisphere Publishing Corporation, 1975. Cuzzi, J.N., Colwell, J.E., Esposito, L.W., et al., Saturn’s rings: PreCassini status and mission goals, in: The Cassini-Huygens Mission, Springer, Netherlands, pp. 209–251, 2003. Duxbury, T.C., Newburn, R.L., Acton, C.H., et al. Asteroid 5535 Annefrank size, shape, and orientation: stardust first results. Journal of Geophysical Research: Planets 109 (E2), 1991–2012, 2004. Edwards, C.H., Wallace, N.C., Tato, C., et al., The T5 Ion Propulsion Assembly for Drag Compensation on GOCE, in: Second International GOCE User Workshop, Frascati, Italy, 2004. Giovannetti, S., Casaregola, C., Pergola, P., et al., Solar Electric LowThrust Asteroid Sample Return Mission, in: Seventh IAA Symposium on Realistic Near-Term Advanced Scientific Space Missions, Aosta, Italy, 2011. Goebel, D.M., Katz, I. Flight Ion and Hall Thrusters, in: Fundamentals of Electric Propulsion: Ion and Hall Thrusters. John Wiley & Sons, Inc., Hoboken, NJ, USA, 2008. Lamy, P.L., Toth, I., Weaver, H., et al. The nucleus of Comet 67P/ Churyumov-Gerasimenko, the new target of the Rosetta mission. Bulletin of the American Astronomical Society 35, 2003. McNamara, P., Racca, G. Introduction to LISA Pathfinder, LISA-LPFRP-0002, Issue 1, 2009. Namouni, F., Christou, A.A., Murray, C.D. Coorbital dynamics at large eccentricity and inclination. Physical Review Letters 83 (13), 2506– 2509, 1999. Nathues, A., Boehnhardt, H., Harris, A.W., et al. ASTEX: an in situ exploration mission to two near-Earth asteroids. Advances in Space Research 45, 169–182, 2010. PROBA-IP, Micro/Mini-Satellite Interplanetary Mission, ESA Executive Summary, 2009. Rayman, M.D., Varghese, P., Lehman, D.H., et al. Results from the deep space 1 technology validation mission. Acta Astronautica 47 (2), 475– 487, 2000. Russell, C.T., Capaccioni, F., Coradini, A., et al. Dawn mission to Vesta and Ceres. Earth, Moon, and Planets 101 (1–2), 65–91, 2007.
P. Pergola / Advances in Space Research 52 (2013) 1622–1633 Surampudi, S. Overview of the Space Power Conversion and Energy Storage Technologies, NASA-Jet Propulsion Laboratory internal presentation, Pasadena, CA, 2011. Takagi, Y. Science Instruments on Hayabusa Follow on missions, in: International Symposium Marco Polo and other Small Body Sample Return Missions, Paris, France, 2009. Tato, C., de la Cruz, F. Power Control Unit for Ion Propulsion Assembly in GOCE Program, in: International Electric Propulsion Conference IEPC-2007-295, Florence, Italy, 2007. Taylor, J. Dawn Telecommunications, Design and Performance Summary Series, Article 13, DESCANSO, Jet Propulsion Laboratory, Pasadena, CA, 2009. Veverka, J., Farquhar, B., Robinson, M., et al. The landing of the NEARShoemaker spacecraft on asteroid 433 Eros. Nature 413 (6854), 390– 393, 2001.
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Wallace, N., Jameson, P., Saunders, et al., The GOCE Ion Propulsion Assembly Lessons Learnt from the First 22 Months of Flight Operations, in: International Electric Propulsion Conference IEPC2011-327, Wiesbaden, Germany, 2011. Weaver, H.A., Gibson, W.C., Tapley, M.B., et al. Overview of the new horizons science payload. Space Science Reviews 140 (1–4), 75–91, 2008. Wells, N., Walker, R., Green, A., et al. SIMONE: interplanetary microsatellites for NEO rendezvous missions. Acta Astronautica 59 (8–11), 700–709, 2006. Wertz, J.R., Everett, D.F., Puschell, J.J. Space Mission Engineering: The New SMAD. Microcosm Press, Hawthorne, CA, 2011. Yoshikawa, M., Fujiwara, A., Kawaguchi, J.I. Hayabusa and its adventure around the tiny asteroid Itokawa. Highlights of Astronomy 14, 323–324, 2007.
ScienceDirect Advances in Space Research 52 (2013) 1622–1633 www.elsevier.com/locate/asr
Small satellite survey mission to the second Earth moon P. Pergola ⇑ Aerospace Engineering Department, University of Pisa, Via G. Caruso 8, 56122 Pisa, Italy Received 15 April 2013; received in revised form 24 July 2013; accepted 29 July 2013 Available online 6 August 2013
Abstract This paper presents an innovative space mission devoted to the survey of the small Earth companion asteroid by means of nano platforms. Also known as the second Earth moon, Cruithne, is the target identified for the mission. Both the trajectory to reach the target and a preliminary spacecraft budget are here detailed. The idea is to exploit high efficient ion thrusters to reduce the propellant mass fraction in such a high total impulse mission (of the order of 1e6 Ns). This approach allows for a 100 kg class spacecraft with a very small Earth escape energy (5 km2/s2) to reach the destination in about 320 days. The 31% propellant mass fraction allows for a payload mass fraction of the order of 8% and this is sufficient to embark on such a small spacecraft a couple of nano-satellites deployed once at the target to carry out a complete survey of the asteroid. Two 2U Cubesats are here considered as representative payload, but also other scientific payloads or different platforms might be considered according with the specific mission needs. The small spacecraft used to transfer these to the target guarantees the manoeuvre capabilities during the interplanetary journey, the protection against radiations along the path and the telecommunication relay functions for the data transmission with Earth stations. The approach outlined in the paper offers reliable solutions to the main issues associated with a deep space nano-satellite mission thus allowing the exploitation of distant targets by means of these tiny spacecraft. The study presents an innovative general strategy for the NEO observation and Cruithne is chosen as test bench. This target, however, mainly for its relevant inclination, requires a relatively large propellant mass fraction that can be reduced if low inclination asteroids are of interest. This might increase the payload mass fraction (e.g. additional Cubesats and/or additional scientific payloads on the main bus) for the same 100 kg class mission. Ó 2013 COSPAR. Published by Elsevier Ltd. All rights reserved. Keywords: Low thrust; Interplanetary missions; Cruithne asteroid; Cubesat; Small satellites
1. A small satellite observation mission During the last 25 years a dozen of space missions visited small bodies in our solar system. During its journey towards Jupiter, in October 1991, the Galileo spacecraft passed at 1.600 km from 951 Gaspra at a relative speed of about 8 km/s, making this asteroid the first one ever observed by a spacecraft (Belton et al., 1992). Seven years later, in 1998, 433 Eros was visited by the NEAR Shoemaker probe (Veverka et al., 2001). This probe first had a flyby with the asteroid and 2 years later, in 2000, started orbiting the same body, making 433 Eros the first asteroid ever orbited by a spacecraft. The same probe visited also ⇑ Tel.: +39 050967211.
E-mail address: [email protected].
another asteroid 253 Mathilde in the 1997 and the Galileo spacecraft itself carried out another asteroid flyby during its path with the 243 Ida in 1993. In 1991 Deep Space 1 was the first probe to test in flight an ion thruster as main propulsive system. This spacecraft reached the asteroid 9969 Braille in 1999 and the mission was extended twice to include also the encounter with Comet Borrelly in 2001 (Rayman et al., 2000). Several interplanetary missions visited, typically with some distant fly-by, an asteroid as secondary scientific target. Beside the already mentioned Galileo, also Cassini, that visited the asteroid 2685 Masursky in 2000 (Cuzzi et al., 2003), and New Horizon, that visited the asteroid 132524 APL in 2006 (Weaver et al., 2008), are worth to be mentioned. The Stardust mission, where the primary targets were to collect samples of cosmic dust and of the dust of the Wild
0273-1177/$36.00 Ó 2013 COSPAR. Published by Elsevier Ltd. All rights reserved. http://dx.doi.org/10.1016/j.asr.2013.07.043
P. Pergola / Advances in Space Research 52 (2013) 1622–1633
2 comet coma, also visited the asteroid 5535 Annefrank on November 2002 at a distance of 3079 km (Duxbury et al., 2004). The first attempt to return asteroid samples to Earth was made by the Hayabusa spacecraft. This probe, equipped with ion thrusters, visited the asteroid 25143 Itokawa in 2005. It carried out a soft landing on the body and collected some samples that were brought back to Earth by means of a re-entry capsule (Yoshikawa et al., 2007). On their way toward their targets there are currently two probes that have already visited some asteroids as side scientific targets. The Rosetta spacecraft (Lamy et al., 2003) is on its way to reach in the 2014 the comet 67P/Churyumov– Gerasimenko. In the September 2008, after two Earth and one Mars swing-bys, the probe visited for 7 min the asteroid 2867 Sˇteins at a distance of 800 km and a relatively slow speed of 8.6 km/s (Accomazzo et al., 2010). Two years later, in July 2010, Rosetta visited also the asteroid 21 Lutetia (a large main belt asteroid) after another Earth swing-by and before entering in hibernation mode (Lamy et al., 2003). The Dawn spacecraft, equipped with three Xenon-fed ion thrusters, is another low thrust mission for the exploration of Vesta and Ceres, the two largest bodies of the main asteroid belt (Russell et al., 2007) (Taylor, 2009). Vesta has been already visited by Dawn from July 2011 to September 2012 and Ceres is scheduled to be reached in February 2015 (Russell et al., 2007). All the probes considered so far reached some asteroids, although in many cases these bodies were not the primary mission targets and the asteroid analysis was carried out only by means of short or distant fly-bys. These probes range from 300 kg up to 1200 kg for the dedicated missions (up to 2500 kg if considering also probes dedicated to planetary observations, like Galileo and Cassini). The great majority of these missions used some complex interplanetary trajectories, with multiple planetary swing-bys and thus long transfer times, and were aided by very powerful launches. New Horizon, for instance, the first spacecraft launched directly into a solar escape trajectory, was launched with a hyperbolic excess energy (C3) of about 170 km2/s2 (Weaver et al., 2008). The aim of the study is to present another option for the exploration of an asteroid. A dedicated mission is considered where the target is intended to be investigated in great detail. With the aim at reducing the mission cost and complexity the paper aims at demonstrating the feasibility of such a dedicated asteroid orbiting mission with a 100 kg class spacecraft launched with a small hyperbolic excess velocity, without any planetary swing-by along its path (thus reducing the transfer times and maintaining a high launch window flexibility) and powered by an electric thruster. In addition, the mission presented considers also the possibility of a detailed exploration of the target by means of multiple nano-satellites operating once at the target. The asteroid considered is a rather challenging target and accordingly the propellant mass fraction here identified can be significantly reduced in case of different targets
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allowing for an even smaller launch mass or a larger payload mass fraction. In this perspective the mission proposed has to be intended as based on a set of conservative assumptions. Cruithne is chosen as the scientific target of the small survey mission proposed. As remnant brick of the solar system formation, for its near Earth orbit and for its peculiar orbital resonance, this asteroid might provide a significant scientific outcome. It is worth mentioning that also other studies (Wells et al., 2006; Giovannetti et al., 2011; PROBA-IP, 2009) proposed small asteroid survey missions, but these are different from the one detailed in the paper for the specific target chosen, the kind of payload considered and thus the exploration phase, the electric propulsion approach and the order of magnitude of the envisaged spacecraft mass. By exploiting the high thrust efficiency of electric thrusters and the current miniaturization trend of space electronic devices, it is proven that Cruithne can be reached by a 100 kg class spacecraft. Once at the target the main spacecraft bus acts mainly as relay system for telecommunication of the two nano-satellites deployed; however, a few kg of additional scientific payload can be also added to the main probe to complete the asteroid science phase. This would result in a larger launch mass thus in longer transfer times and additional propellant mass as detailed in Section 3.2. The spacecraft deploys a couple of 2U Cubesats to obtain a multiple observation of the asteroid surface and interior. These nano-satellites can operate in different wavelengths and/or be equipped with charged particles probes, mass spectrometers, accelerometers and so onto obtain a complete picture of the whole body. The aim of the paper is to demonstrate the feasibility of this innovative mission concept rather than a specific sizing of these nano-platforms. With such approach the complete asteroid exploration can be carried out by means of a single small scale mission. In the paper the mission design phase (see Section 3) is detailed to estimate the mission Dv and the propellant mass fraction required for the subsequent phase of preliminary spacecraft sizing (see Section 4). All the main subsystems are sized by using COTS and state of the art figures to provide a conservative design. 2. Cruithne Also known with the official designation of asteroid 3753, or 1986 TO, Cruithne is a 5 km wide Near Earth Asteroid (NEO) (Namouni et al., 1999). The body is locked in a 1:1 mean motion resonance with the Earth, thus the two bodies appear to follow each other on their heliocentric paths. For this reason Cruhitne is also called the second Earth moon although it is not directly affected by the Earth gravity field. Despite this orbital resonance, at the closest approach, which happens every 285 years, Cruithne comes only within 12.5e6 km, sufficiently far to be not considered as a possible impactor, but enough close to have its orbit perturbed by the Earth (Brasser et al., 2004).
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As seen from Earth, Cruithne follows a bean-shaped horseshoe orbit that lasts 364 days (Namouni et al., 1999). Since its period is slightly shorter than 1 year, the Earth remains behind the bean-shaped orbit a little more each year. As a consequence from our point of view the asteroid orbit is not really closed, but it looks more like a spiral loop moving away from Earth (Namouni et al., 1999). As seen in a heliocentric frame, the asteroid moves in a very elongated elliptic path (crosses the Mars orbit and glazes the Mercury one) with 0.4840 AU perihelion, 1.511 AU aphelion and 19.8° of inclination. This last aspect, in particular, makes the mission design rather challenging and requiring a total impulse significantly larger than any almost equatorial NEO. The asteroid orbit as seen in the inertial heliocentric frame is shown from three perspectives in Fig. 1. To obtain the trajectory shape in a reference frame rotating with Earth, the instantaneous relative position of the two bodies is required (Fig. 2, left). These distance vectors are then translated into the Earth position at a date (Fig. 2, centre) and their envelope finally gives the beanshape orbit in magenta in the rightmost plot of Fig. 2. Changing date, the Earth position changes and the apparent asteroid orbit moves describing a three dimensional horse-shoe orbit covering a torus around Earth’s orbit. The bean-shape orbit completes such a torus each 770 years (Brasser et al., 2004).
3. Trajectory design With the aim at reducing the fuel mass fraction, thus allowing a small spacecraft (100 kg class) to perform an asteroid observation mission, an electric propulsion transfer is considered. The path is designed in the heliocentric two body system minimizing the transfer time. This approach has been preferred with respect to a minimum mass transfer to limit the ground segment and to avoid that an extended permanence in space might jeopardize the two nano-spacecraft (or, in general, the payload). The main challenge associated with the transfer sought is the relatively high inclination of Cruithne. This forces the electric transfer to significantly change both the velocity in the ecliptic plane and the velocity perpendicular to that plane. Still following the idea to present a relatively cheap mission, the spacecraft hyperbolic excess energy has been arbitrarily limited to 15 km2/s2. A 100 kg class spacecraft with such C3 and with reduced dimensions (less than 2 m, see Section 4) is compatible with the great majority of the small/medium launcher fairings, even as piggyback payload. The mission analysis presented in Section 3.2 has been refined after the spacecraft design detailed in Section 4. An initial mass of 100 kg has been used and the spacecraft has been considered equipped with two twin T5-like ion thrusters.1 This thruster is used only for providing some 1 http://earth.esa.int/workshops/goce04/participants/101/ poster_T5_Ion_Propulsion_A0.pdf
reference values and any other thruster can be considered changing the spacecraft characteristics and accordingly the transfer time and propellant mass consumption. This specific ion thruster is actually able to provide a thrust between 0.6 and 20.6 mN with a specific impulse reaching up 3500 s and with a power requirement between 55 and 585 W across the thrust range (Edwards et al., 2004). 3200 s of specific impulse and 18 mN thrust were used as nominal values (Goebel and Katz, 2008) and two thrusters have been assumed firing in parallel for the whole transfer. With the aim at outlining a small and affordable mission, no cold redundancy on the thruster has been considered. This means that if one of the thrusters fails the whole mission might still be accomplished with significantly longer transfer times and a new trajectory design. 3.1. Approach The indirect optimization approach here used is well known to be rather sensitive to the initial conditions supplied (Bryson and Ho, 1975). To limit such sensitiveness a preliminary, more robust gradient optimization scheme has been implemented. This step is not required to be convergent and just few iterations, typically less than 10, are sufficient to obtain an initial guess for the continuous and discrete Lagrange multipliers. Starting from these values an indirect optimization scheme has been implemented. The method integrates together the equations of motion and the adjoint equations (Betts, 1998). To increase the numerical stability of the numerical routine and to avoid any singularity associated with the classical orbital elements, the equinoctial elements have been used. The calculus of variation problem is posed in the Mayer form, where the transfer time is the fnctional to be minimized, and the control vector is composed by the continuous and discrete Lagrange multipliers parameterizing the in-plane and out-of-plane thrust angels. The iterative scheme aims at obtaining a stationary value of the functional (Bryson and Ho, 1975). The boundary conditions to be satisfied derive from imposing that the final state of the propulsive phase belongs to the asteroid orbit (Betts, 1998). Together with these boundary constraints also the transversality condition arising from the indirect problem formulation has to be satisfied (Betts, 1998). The anomaly of the encounter point is not fixed in advance but results from the optimization scheme as well as the total transfer time. Since, to reduce the transfer time, coasting arcs were not allowed, from the transfer time the propellant mass consumption can be directly computed and from this the Tsiolkowsky equations allows for computing the mission total Dv. The initial spacecraft velocity is modelled as the Earth heliocentric velocity augmented of some energy assumed to be delivered by the launcher. This initial Dv is always considered as tangential to the Earth heliocentric velocity. A grid search on the departing energy has been carried out considering escape C3 energies between 1 and 15 km2/s2.
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Fig. 1. 3753 Cruithne orbit in three views in the heliocentric inertial frame.
Fig. 2. Construction of the Cruithne orbit relative to Earth. The bean-shape orbit in the rightmost plot represents the envelope of the relative EarthCruithne position vectors over one revolution centred in the Earth position at a date.
The iterations have been stopped when a tolerance at least of the order of 1e 6 was reached on the accuracy of the final constraints. The equations of motions have been numerically integrated in non dimensional units using a variable step Runge–Kutta 4–5 order integrator with absolute and relative tolerance of the order of 1e 9.
3.2. Results The resulting trajectory implements a direct transfer that lasts less than one year and reaches the asteroid orbit close to its aphelion (the slowest point). The thrust law is characterised by an in-plane thrust angle aiming at increasing the spacecraft velocity and an out-of-plane angle direct constantly toward the asteroid aphelion (negative z semispace) to modify the orbital inclination. The resulting transfer trajectory is shown in the heliocentric inertial frame in Fig. 3. After the gradient first guess generator the forward shooting approach typically converges in less than 500 iterations with a total CPU time of less than 8 min on a 2.2 GHz laptop with 4 GB ram. The transfer lasts 322 days are requires 31.9 kg of propellant mass. The mission total Dv is about 12.01 km/s and the total impulse is about 1.01e6 Ns. The total impulse
a single T5 is able to deliver is of the order of 1.5e6 Ns2 and up to 8500 on/off cycles can be handled (not stringent limit in the planned mission as coasting arcs are not allowed and only few eclipses have to be considered), thus the couple of thrusters should already offer a safe design for the planned mission. The T5 total impulse, however, is the actual limit to take into account in case of mission re-planning for a single thruster failure. The grid search on the departing energy results in a rather flat behaviour in terms of transfer time and propellant mass consumption. This is mainly due to the high orbital inclination of the target that requires some time to be reached (the initial velocity is always supplied tangential to the Earth velocity and thus in the ecliptic plane). This causes also the transfers characterised by a high C3 to require a longer time (more mass) since the initial velocity in the ecliptic plane does not allow reaching the Cruithne orbit in a convenient point. The minimum time (that is also minimum mass as no coasting arcs are considered) trajectory, see Fig. 3, among the results obtained has a C3 energy of 5 km2/s2.
2 http://earth.esa.int/workshops/goce04/participants/101/ poster_T5_Ion_Propulsion_A0.pdf
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Fig. 3. Electric transfer trajectory from Earth to Cruithne orbit as seen in the heliocentric inertial frame.
As the main aim of the study is to present an innovative mission scenario making use of nano platforms to explore NEO objects, the specific mission might differ from the case here presented in particular for the launch mass. Accordingly a sensitive analysis with respect to a variation of the launch mass, from 100 to 150 kg has been carried out. Fig. 4 shows the variation of the transfer time and propellant mass as a function of the launch mass. All the cases have been optimised with the approach described and the launch C3 has been adapted to allow for the convergence of the numerical procedure. In particular up to 110 kg of launch mass 5 km2/s2 proved to be sufficient; from 110 to 130 kg, instead, 10 km2/s2 were considered and from 130 to 150 kg a C3 of 15 km2/s2 was used. It has been observed in this analysis the encounter point with Cruithne orbit moves toward its perihelion. This is a consequence of the lower acceleration level provided by the thruster that
requires a longer time to match Cruithne orbital parameters. From this launch mass sensitivity analysis it results that the propellant mass fraction remains between the 28% and the 32% of the initial mass, thus leaving in all cases about the 70% for the spacecraft systems and the payload. In addition, up to approximately 120 kg of launch mass, the transfer takes less than one year. To provide an impulsive comparison the same transfer has been assessed by means of a rough grid search over one year of Earth departure dates and one year of transfer times. The Lambert problem has been solved for this grid of data and it results that the same target can be reached with a standard bi-impulsive manoeuvre in approximately 8 months and with 17 km/s of total Dv. This value, in particular, is comparable with the electric transfer Dv as in this case the significant inclination change has to be performed
Fig. 4. Transfer time and propellant mass variation with respect to a variation in the launch mass.
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Fig. 5. The spacecraft during the operative phase with the ion thrusters operating, one Cubesat fully deployed and the other one ready to be delivered.
anyway and we cannot take advantage of the negligible gravity field of the target that might have reduced the magnitude of the second impulse. The 84% of the total chemical Dv has to be provided at the departure and it realises the largest part of the inclination change manoeuvre. This Dv might be provided by the launcher, although it is significantly larger than the 5 km2/s2 C3 used for the low thrust analysis.
4. Spacecraft layout To contain dimensions and to better exploit the internal volume, the spacecraft is conceived as an octagonal prism, like the Lisa Pathfinder bus (McNamara and Racca, 2009). The upper surface of the bus is reserved for solar arrays and the bottom surface for the propulsion system. The lateral surfaces are used for the telecommunication system, the radiators and for the Cubesat deploying system. Fig. 5 presents a conceptual view of the probe during the operative phase with the two ion thrusters active (for instance for orbit maintenance purposes), one Cubesat already fully deployed and the other one still in its bay. The main bus structure assures the mechanical support to the spacecraft subsystem and to the deploying system of the two 2U Cubesats. It is an octagonal prism with an outer diameter of 1.4 m (1.6 m is the diameter of the upper face) and 0.7 m height. The upper face is completely covered by a sunshield panel supporting the body mounted solar array. To produce the 1140 W required to operate the two electric thrusters and to satisfy the spacecraft power needs (200 W assumed, see Section 4.1), two additional solar arrays half the size of the body mounted one are deployed after the Earth escape. This 8 m2 panel produces 1430 W End Of Life (EOL) at 1.511 AU distance from the Sun (Cruithne aphelion). The internal spacecraft structure is composed by a large central cylinder from which a set of transversal shear walls connects the inner cylinder to the outer panels, see Fig. 6. The whole structure is assumed to be constructed of sandwich panels with carbon fibre laminate skins bonded to aluminium honeycomb core. The internal cylinder accommodates the ATK 80194-1 18.8 L Xenon tank. This tank has been chosen after a tradeoff between required propellant and available volume. The tank itself has a mass of 7.67 kg, is of spherical shape
with a radius of 340 mm and operates at a pressure of 250 bar.3 At such pressure, and considering a storage temperature around 310 K, approximately 17.8 L are required to store 35 kg of propellant. Of these almost 32 kg are reserved for the transfer and 3 are considered for the attitude control system and for eventual additional contingency. The two ion thrusters are located on the bottom plate of the bus. These are mounted with a small, 7°, angle to have the thrust direction passing through the spacecraft centre of mass. A cylindrical shell is placed all around the thrusters to limit the plume interaction with the other spacecraft surfaces. Each thruster weights about 1.7 kg and has a cylindrical shape 17 25 mm. The control electronics of the thruster is placed in one of the internal bus bays. It is composed by two main units. The proportional Xenon feed system includes the pressure control unit and the flow control section (bang bang pressure regulation valve and Xenon flow control unit). The module receives Xenon directly from the tank (5–250 bar) and regulates it down to approximately 2.5 bar. The module includes also particle filters, pressure transducers and insulation valves (Wallace et al., 2011). The ion propulsion control unit considered is a highly integrated module. It includes the control electronics (MIL-1553 interfaces and timing synchronisation with the spacecraft OBC), the DC-AC converter with low voltage (LV) and high voltage (HV) output, the converter from spacecraft DC to the HV DC source required for the ion beam, LV and HV output control for internal function and for the power distribution to the other units. The control module runs also on an independent micro processor (like the ERC 324) to implement the thrust control algorithm of the two ion thrusters. This unit is estimated to weight approximately 12 kg, requires up to 20 W and has estimate dimensions of 450 200 200. The thermal control system must guarantee a very stable thermal environment during the whole transfer and at destination. During the path the Sun heating is shielded by the thermal shield and the solar array mounted on the upper surface. Two radiators are mounted on two sides of the 3
http://www.psi-pci.com/Data_Sheets_Library/DS194.pdf http://microelectronics.esa.int/erc32/chipset/Hardware%20and% 20Documentation%20Status%20of%20the%20ERC32%203-Chipset% 20i2r1a.pdf 4
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Fig. 6. A global view of the complete spacecraft with deployed arrays (upper) and two views of the internal spacecraft structures where some internal walls have been hidden to show the internal subsystems.
spacecraft bus for a total surface of 1.4 m2 (total emissivity 0.8, solar absorbance 0.21). The whole external structure is folded with a 10 layers Multi Layer Insulator (MLI) (large conductance and emissivity, 0.4 kg/m3 density), while the inner structures are covered with a 20 layer MLI (high radiation protection, 0.6 kg/m3 density). A nominal operation temperature of 290 K is considered with a maximum admitted variation of 10°. The main internal heat sources considered are the two ion thrusters; for these a thrust efficiency of 55% is considered. This means that the 45% of the incoming power is transformed into heat to be dissipated. From experimental tests, however, it is known that this heat flow is mainly radiated away with the plume and only the 40% is the heat directly transferred to the structure (about 170 W in total). Active thermal control system elements are considered for the most critical elements like electronic modules and the tank. As general philosophy, it is assumed to provide the minimum necessary heater power in the cold case scenario so that the lower temperature of each unit is maintained near the bottom of the allowable range. Depending on the science phase duration, the most critical situation (hot case) is when Cruithne is at its perihelion. In this situation rather high temperatures can be reached and it is assumed to maintain the spacecraft attitude to have sun pointing solar panels, which offer also the proper thermal shielding. The spacecraft is a three-axis controlled spacecraft. The reaction wheels are desaturated by means of the two ion
thrusters and a set of 6 small (50 g each) Xenon-fed 10 W resistojets mounted on the structure corners. 3 kg of additional propellant are embarked for this purpose. The vertical wheel is placed on top of the Xenon tank, the horizontal one on one side of the same and the other wheel along the spacecraft principal axis, together with the spare (inclined) one and the associated electronics on a bus bay, see Fig. 6. Each wheel weights about 0.9 kg (e.g. RW 905). The attitude and orbit control system is completed with one Earth horizon sensor, two autonomous star trackers and two digital Sun sensors for the Sun acquisition (McNamara and Racca, 2009). These sensors are mounted on the bottom bus plate with the proper orientation (star trackers are mounted where there is the radiator and below one of the two solar array wings, a spacecraft side that should be constantly pointing the deep space), see Fig. 7. During the transfer the attitude is constrained by the thrust steering law, thus it may happen that the solar panels are not perfectly Sun pointing. Actually an incidence Sun light angle up to 40° can be reached, thus the solar array surface has been properly oversized to take into account this situation. This option has been preferred to a Solar Array Driver Assembly (SADA) system as it is lightweight, cheap and more affordable for a small mission.
5
http://www.astrofein.com/2728/dwnld/admin/Datenblatt_RW90.pdf
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Fig. 7. Internal views of the spacecraft from the top (left) and form the bottom (right) the upper solar panel and the bottom covering surfaces have been hidden to show the interior structure and components.
The power subsystem guarantees a stable regulated bus with a constant voltage of 28 V. Because of its low mass and simple management, a pack of three Li-ion batteries has been chosen as energy storage subsystem. The battery pack provides a total of 300 Wh considered for the early mission phases before the two wings are flipped, for the short eclipses and for the peak power mode requests. This battery pack is placed in a spacecraft bay, see Fig. 7, and it is equipped with the proper heaters. Two of these bays are protected against radiations with a 1 mm aluminium shield. In these bays are placed all the sensitive electronic components and the battery pack. These two bays are also the two directly behind the radiators to assure an easier thermal path to control the electronic component temperature, see Fig. 6 and 7. The On-Board Computer (OBC) is in charge of handling the on-board data, the attitude and orbit control system, the management of the platform and the Cubesat deploying system once at destination. Furthermore the control system performs also monitoring functions of the main systems and provides safe system reconfiguration capabilities. The OBC is composed of two processing modules of 33 MHz each (e.g. RAD60006) and a redundant solid state 100 GB memory device. The OBC interface with the spacecraft is assumed to be based on a standard MIL-bus 1553B, while the internal backplane link uses the Space-Wire standard. Dedicate modules for telecommand, telemetry, reconfiguration and for actuator and sensor interfaces are also considered. The OBC primary computer is estimated to weight about 1.5 kg7 and 0.8 kg are considered for the data recorder unit and cabling. 6
http://atc2.aut.uah.es/mprieto/asignaturas/satelites/pdf/rad6000.pdf http://www.sst-us.com/shop/satellite-subsystems/obdh/radiation-tolerant-flight-computer-obc-695
On a side bay of the spacecraft also the Cubesat deploying structure is placed. The two access gates are directly obtained in the main structure wall. The two deployers are placed on top of each other mounted on a tubular structure. The command electronics and all the security/ check systems to monitor the Cubesat health during the transfer are placed on the back of the same structure. Each deployer weights about 1.75 kg and sizes 300 130 180 mm.8 The communication system operates in X-band with an uplink frequency of 7230 MHz and 8495 MHz downlink. The system provides both for commanding and housekeeping telemetry of the main spacecraft and acts as relay system for the telemetry of the Cubesats. This system, indeed, receives the data from the 2U micro satellites and sends these to the ground station. The communication frequency with the payload is assumed to be in S-band. The spacecraft is equipped with a 1.2 m diameter high gain antenna (slightly smaller than the antenna of other deep space missions, like Hayabusa and Mars Global Surveyor) mounted on one side of the spacecraft bus (Taylor, 2009) and two low gain omni-directional antennas mounted in opposite directions to allow for the communications with the Earth station during the early transfer phases and when the main antenna in not Earth pointing, see Figs. 6 and 7. During the operation phase, instead, only the main antenna is in charge of transmitting data to Earth and the two omni-directional antennas (easier to communicate with a moving spacecraft constantly changing aspect angle) are reserved for the bi-directional communications with the Cubesats. The maximum datarate assumed is of 60 ksymbols/s for the low gain antennas and 120 ksymbols/s for
7
8
http://www.isispace.nl/brochures/ISIS_ISIPOD_Brochure_v.7.11.pdf
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the high gain one (McNamara and Racca, 2009). Two Xband transponders are included in the telecommunication system dedicated electronic (placed in the bus bay behind the main antenna) for redundancy and an additional amplification module is included at the transponder output (Taylor, 2009). The telecommunication system is equipped with a dedicated solid state memory of 60 GB for the science data following a “store and forward” strategy. The high gain antenna is estimated to weight 3 kg, the low gain antenna 0.2 kg each and the telecommunication dedicated electronic (X-band solid state amplifier and solid stage data recorder) 1.3 kg. Based on the results of the trajectory analysis in the scenario considered, almost 122 days of visibility are guaranteed after the probe arrival. This means that about 4 months of direct communications are allowed before that the line-of-sight between the spacecraft and the Earth passes beyond Mercury orbit where the intense Sun radiation might affect the data relay. During this period the Earth-spacecraft distance varies between 2.14 (at the arrival point close to the Cruithne aphelion) and 0.3 AU with an associated communication delay between 2.5 and 17.8 min. The minimum Earth distance is reached 76 days after the rendezvous. To assess the telemetry/telecommand and tracking signal availability and quality during the mission phases, a preliminary link budget has been performed. Assuming to use the Deep Space Network antennas as ground receivers and maintaining a signal/noise margin of at least 2 dB, it follows that at the closest point the whole 120 ksymbols/s data rate can be used (even higher rates are allowed), while in the farthest point the maximum data rate allowed for the downlink is 7 ksymbols/s. As a consequence, assuming 8 h of ground station visibility per day, from 0.2 up to 3.2 GB per day can be downloaded. The remaining data (depending on the specific mission application) have to be stored in the 60 GB solid state memory and forwarded at some closer point. 4.1. Mass and power budgets The structure mass is estimated according with state of the art figures as the 10% of the launch mass. This mass includes the carbon fibre honeycomb bus structure, the support plates for the subsystems and the 1 mm shield for the two bays where the more critical components are placed. The propulsion system mass is about 25 kg including the two thrusters, the tank and the main electronic units. The power condition unit and the associated modules are estimated to weight around 12 kg (Tato and de la Cruz, 2007), 1 kg for the bang-bang pressure regulation unit and 1 kg for the Xenon flow control unit together with the pre-regulation valve. To deliver the required 1140 W EOL at Cruithne aphelion, almost 6.4 m2 solar arrays are required. To consider also some cell aging in space, possible smaller solar cell
efficiency, possible peak power requirements and the nonperpendicular Sun incidence angle, a total of 8 m2 has been covered with solar cells. 30% solar cell efficiency, a powerto-mass ratio of 200 W/kg (Bailey and Raffaelle, 2011) (Surampudi, 2011) for the solar arrays and a battery power density of 300 Wh/kg (Surampudi, 2011) have been considered. With these assumptions made the power generation system mass is about 5.7 kg. The battery pack, instead weights almost 1 kg assuming to size this to sustain only the bus needs (no firing in eclipses). The attitude control system weights a total of 5.2 kg where the largest contribution is due to the reaction wheels (3.6 kg). 0.5 kg are considered for the resistojet thrusters and the associated piping. The system electronics is estimated to weight only 0.25 kg (as it mainly acts as data buffer with the OBC and local control unit), the two Sun sensors weight a total of less than 200 g, 0.5 kg the Earth horizon sensor and 0.25 kg each the two star trackers. The telecommunication system weights some 4.9 kg where the largest contribution, 61%, is due to the high gain antenna. The thermal system is estimated for a total mass of 3.2 kg. The MLI used for the bus surfaces, see Section 4, is estimated to weigh about 2 kg and the 1.4 m2 radiator surface is considered weighting 1.2 kg (Wertz et al., 2011). The mass budget is summarized in Table 1. Here 5% of additional contingency is considered on each subsystem mass. The total spacecraft mass, including also the propellant and the contingency is 103 kg and 68 kg is the system dry mass. In Table 1 the mass budget presented is also compared with some state of the art figures of the various subsystems (Wertz et al., 2011). This last comparison confirms that the assumptions and the analyses made are coherent with reference values. A couple of observations are mandatory. The propulsion system for the spacecraft designed is significantly heavier than the reference value; this because the statistical analysis used to derive these state of the art figures is based only on chemical thrusters (Wertz et al., 2011). In addition, in the mass budget of this system also the power management mass is included, thus the power generation system mass (where usually the power management equipment is included in case of chemical spacecraft) is smaller than the reference value The structure considered is slightly lighter than a standard one taking into account the different design philosophy (carbon fibre honeycomb and internal sets) with respect to the spacecraft used for the statistical analysis. A static and dynamic detailed sizing of the structure is however required, and it is left for a more detailed study. The payload includes the two deployers, the two 2U Cubestas (2 kg each) and 0.5 kg of harness and support. This figure is slightly smaller than a standard payload figure for a planetary mission. Nonetheless the deployable payload represents an innovative mission aspect, offering a complete covering of the target, which might be replaced or associated with some other scientific equipment.
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Table 1 Spacecraft mass budget. The estimated mass, the mass including 5% contingency, the initial mass percentage and a comparison with state of the art figures is given for each subsystem. Subsystem
Mass (kg)
Total mass with 5% contingency (kg)
Percentage of dry mass (%)
Reference value (%)
Structure Propulsion system + Tank Xenon propellant Payload Power generation system + energy storage system Attitude and orbit control system Command and data handling system Thermal control system Telecommunication system Launch mass Dry mass
10.0 25.0 35.0 8.0 6.7
10.5 26.2
15.4 38.3
20 13
8.40 7.0
12.3 10.2
15 21
5.2 2.4 3.2 4.9 100.4 65.4
5.4 2.5 3.3 5.1 103.4 68,4
7.9 3.7 4.8 7.5
6 4 6 7
As other payload options several standard instruments can be considered. Maintaining the same payload mass of 8 kg (although also the whole spacecraft structure might be different for accommodating specific payloads) other payload options might be a Laser Imaging Detection and Ranging (LIDAR) (3.7 kg, 17 W) allowing for high resolution surface mapping, geology and geomorphology (Takagi, 2009); a Remote X-ray Fluorescence (XRF) Spectrometer (3 kg, 15 W) to determine the surface composition (possibly also of lighter elements) and processes (Takagi, 2009); an Infrared (IR) camera (3,5 kg, 29 W) for different wavelength analysis (Takagi, 2009) and/or IR-spectrometer (e.g. a Single Ion Recording type microspectrometer for moderate spatial resolutions) to be used for asteroid classification; a close-up camera and panoramic stereo camera (1,5 kg, 11 W, total) for high-resolution imaging of surface materials and stereo mapping/ characterization of possible landing or sampling sites (Nathues et al., 2010); an optical and electron microscope (1 kg, 8 W, total) for the study of geophysical and structural properties of the body and the micron-scale characterization of surface materials together with their elemental/mineralogical analysis (Nathues et al., 2010). Moreover a payload package including several of the SIMONE payload equipment might be considered (13 kg, total). The list is not exhaustive and the specific payload combination, up to the payload mass fraction identified in the scenario, depends on the mission scientific targets. From Table 1, it results that the whole spacecraft launch mass is slightly larger than 100 kg. A refined mission analysis with this update value shows that the same transfer can be realised with almost the same transfer time (a couple of days longer) and with an additional propellant mass requirement that is well below the propellant mass contingency considered. The bus power requirement is estimated to be of the order of 200 W. When operating the two thrusters require 2 470 W, thus the total power load in the electric propulsion transfer mode is of 1140 W. Once at target the main power consumption is represented by the high gain antenna
during the Earth downlink/uplink mode. During the launch and the early operative phases the solar panels are still folded, thus the entire power load is managed by the batteries. After this phase, however, only short eclipses and limited power peak requirements are foreseen. The power required by each subsystem is estimated using some standard figures (Wertz et al., 2011). The thermal control subsystem might require up to 15% of the total incoming power, i.e. 16.5 W that can be delivered when thrusters are not operating. This peak consumption might be considered in a very cold or very hot situation. The transfer ranges between 1 and 1.5 AU (not particularly critical), while the science phase can reach up to 0.5 AU, but in this case the main thrusters are not operating and there is are hundreds of W available. During normal operations only 50 W are estimated to be required to maintain the temperature in the allowable range (and, in particular, to avoid Xenon freezing and for battery heaters). The power management system itself requires almost the 10% of power to operate, corresponding to 11.4 W. The telecommunication subsystem has very high power consumption only during the Earth downlink/uplink mode. This power consumption can reach up the 18% of the total power budget. These 205 W (on which the link budget has been estimated), however, are supposed to be required only for short transfer arcs (the telemetry and the steering laws are mainly autonomous) and once at the target where the thrusters are not continuously operating. The OBC can reach up the 11%, 125 W, and almost the same power can be required by the attitude control system. As a consequence of this rough power budget there are two main situations that can take place. Or the thrusters are operating and the remaining 200 W can be used to run almost all subsystem, even in peak consumption conditions, although not at the same time (if peak consumption is required, but of course all can be run also in parallel in normal operation modes). Once at the target, however, the thrusters are not required to run continuously, thus all the main subsystem, also in peak power mode can be used together.
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5. Conclusions In this paper a complete mission for the observation of the second Earth moon is presented. The mission is conceived as a survey mission where the main spacecraft together with two nano-satellites can explore the asteroid Cruithne. The paper presents the complete study of such a mission with a first phase of trajectory design and starting from this also a first order spacecraft sizing and design is presented. The main spacecraft is in charge of the Earth-asteroid transfer by means of a couple of ion thrusters. The mission analysis by means of an optimal minimum time approach results in approximately 32 kg of Xenon and less than 11 months of transfer time. The results obtained from this preliminary mission analysis and the subsequent spacecraft sizing are however to be intended as reference values to demonstrate the feasibility of such a mission. Also slightly changing the system mass values and/or the mission analysis approach a 100 kg class spacecraft can still be able to accomplish the mission proposed. The mission analysis output, in particular, can be also scaled to heavier/lighter spacecraft where the thruster specific impulse, the Earth departure C3 and the electric thruster acceleration maintain almost the same value. In situ observation of small bodies, like asteroids, has a very high scientific outcome as their small size did not allow for the modification of the early solar system chemistry by high internal pressures and temperatures. The mission outlined in the paper plans for the multiple body observation by means of two 2U Cubesats (additional observations/analyses are also possible from the main spacecraft). This approach solves several typical issues of the robotic asteroid observation, for instance the typical communication delay and low bit rate from Earth base stations is solved by using the main spacecraft for the telemetry. The two nano-probes can also operate on different orbits and with different payloads. These Cubesats are required to have a bi-directional telemetry with the mother spacecraft. The specific sizing of such nano-satellites is highly dependent of the required analysis (and the constant improvement in the miniaturization of the Cubesat equipment), however, as rough estimation, almost 30% of the Cubesat mass might be allocated for the payload. Multispectral imaging devices, accelerometers, plasma probes and spectroscopy reflector are some of the typical payloads to consider. In addition, the possibility to exploit multiple sampling locations at the same time might also help to design a dynamical picture of the asteroid environment. In nominal conditions in a Low Earth Orbit (LEO) a typical single 2U Cubesat solar panel can generate up to 5 W; this reduces to approximately 2 W considering the Cruithne aphelium and some aging effects. 2 W are slightly less than the same power generated by a single 1U solar panel in LEO. This means that the on-board systems that might be operated on one of such nano platforms can have
power requirements similar to a state of the art 1U Cubesat in LEO. In addition, as shown in Fig. 5, a valuable option to increase the available power of these 2U platforms might be the usage of two solar array wings (e.g. like the ones already used in Defi-C3 and QbX). In this scenario the available power would be some 4 W for each face. The mission conceived is a valuable approach for low cost, fast analyses precursor missions. The approach, indeed, might act as pathfinder for subsequent heavy robotic exploration/landing missions and in a longer run also for human explorations and asteroid mining missions. References Accomazzo, A., Wirth, K.R., Lodiot, S., et al. The flyby of Rosetta at asteroid Sˇteins – mission and science operations. Planetary and Space Science 58 (9), 1058–1065, 2010. Bailey, S., Raffaelle, R. Space Solar Cells and Arrays, in: Luque, A., Hegedus, S. (Eds.), Handbook of Photovoltaic Science and Engineering. John Wiley & Sons, Chichester, West Sussex, UK, pp. 365–401, 2011. Belton, M.J.S., Veverka, J., Thomas, P., et al. Galileo encounter with 951 Gaspra: First pictures of an asteroid. Science 257 (5077), 1647–1652, 1992. Betts, J. Survey of Numerical Methods for Trajectory Optimization. Journal of Guidance, Control and Dynamics 21 (2), 193–207, 1998. Brasser, R., Innanen, K.A., Connors, M., et al. Transient co-orbital asteroids. Icarus 171 (1), 102–109, 2004. Bryson, A.E., Ho, Y.C. Applied Optimal Control. Hemisphere Publishing Corporation, 1975. Cuzzi, J.N., Colwell, J.E., Esposito, L.W., et al., Saturn’s rings: PreCassini status and mission goals, in: The Cassini-Huygens Mission, Springer, Netherlands, pp. 209–251, 2003. Duxbury, T.C., Newburn, R.L., Acton, C.H., et al. Asteroid 5535 Annefrank size, shape, and orientation: stardust first results. Journal of Geophysical Research: Planets 109 (E2), 1991–2012, 2004. Edwards, C.H., Wallace, N.C., Tato, C., et al., The T5 Ion Propulsion Assembly for Drag Compensation on GOCE, in: Second International GOCE User Workshop, Frascati, Italy, 2004. Giovannetti, S., Casaregola, C., Pergola, P., et al., Solar Electric LowThrust Asteroid Sample Return Mission, in: Seventh IAA Symposium on Realistic Near-Term Advanced Scientific Space Missions, Aosta, Italy, 2011. Goebel, D.M., Katz, I. Flight Ion and Hall Thrusters, in: Fundamentals of Electric Propulsion: Ion and Hall Thrusters. John Wiley & Sons, Inc., Hoboken, NJ, USA, 2008. Lamy, P.L., Toth, I., Weaver, H., et al. The nucleus of Comet 67P/ Churyumov-Gerasimenko, the new target of the Rosetta mission. Bulletin of the American Astronomical Society 35, 2003. McNamara, P., Racca, G. Introduction to LISA Pathfinder, LISA-LPFRP-0002, Issue 1, 2009. Namouni, F., Christou, A.A., Murray, C.D. Coorbital dynamics at large eccentricity and inclination. Physical Review Letters 83 (13), 2506– 2509, 1999. Nathues, A., Boehnhardt, H., Harris, A.W., et al. ASTEX: an in situ exploration mission to two near-Earth asteroids. Advances in Space Research 45, 169–182, 2010. PROBA-IP, Micro/Mini-Satellite Interplanetary Mission, ESA Executive Summary, 2009. Rayman, M.D., Varghese, P., Lehman, D.H., et al. Results from the deep space 1 technology validation mission. Acta Astronautica 47 (2), 475– 487, 2000. Russell, C.T., Capaccioni, F., Coradini, A., et al. Dawn mission to Vesta and Ceres. Earth, Moon, and Planets 101 (1–2), 65–91, 2007.
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