power system design of esmo

Acta Astronautica 64 (2009) 244–255www.elsevier.com/locate/actaastro

Power system design of ESMO�

Steve Ulricha,∗, Jean-François Veilleuxb, François Landry CorbinbaNGC Aerospace Ltd., 1650 King Ouest, Office 202, Sherbrooke, Québec, Canada J1J 2C3

bUniversité de Sherbrooke, 2500 boul. de l’Université, Sherbrooke, Québec, Canada J1K 2R1

Received 5 December 2007; received in revised form 3 September 2008; accepted 3 September 2008Available online 1 November 2008

Abstract

The European Student Moon Orbiter (ESMO) spacecraft is a student-built mini satellite being designed for a mission to theMoon. Designing and launching mini satellites are becoming a current trend in the space sector since they provide an economicway to perform innovative scientific experiments and in-flight demonstration of novel space technologies. The generation, storage,control, and distribution of the electrical power in a mini satellite represents unique challenges to the power engineer since themass and volume restrictions are very stringent. Regardless of these problems, every subsystem and payload equipment mustbe operated within their specified voltage band whenever they required to be turned on. This paper presents the preliminarydesign of a lightweight, compact, and reliable power system for ESMO that can generate 720W. Some of the key componentsof the EPS include ultra triple-junction (UTJ) GaAs solar cells controlled by maximum power point trackers, and high efficiencyLi-ion secondary batteries recharged in parallel.© 2008 Elsevier Ltd. All rights reserved.

Keywords: Mini satellite; Moon mission; Electrical power subsystem design

1. Introduction

Since the last few decades, there has been a large in-terest in operating mini satellites, mainly because theyare cheaper to build and to launch. These satellites pro-vide space agencies and researchers an ideal testbedfor in-flight demonstration of innovative and high risktechnologies. Used in an educational context, they also

� Presented as paper IAC-07-B4.6.13 at the 58th International As-tronautical Congress, Hyderabad, India, 24–28 September, 2007.

∗Corresponding author. Tel.: +18193489483;fax: +18193489430.

E-mail addresses: [email protected](S. Ulrich), [email protected](J.-F. Veilleux), [email protected](F. Landry Corbin).

0094-5765/$ - see front matter © 2008 Elsevier Ltd. All rights reserved.doi:10.1016/j.actaastro.2008.09.002

prepare students for their future careers in the spacesector by giving them valuable hands-on experience. Infact, the number of student-built mini satellites is rapidlygrowing, especially in Europe and in North America.The MIT’s SPHERES [1] (synchronized position holdengage and reorient experimental satellites) experimentuses the space environment provided by the Space Shut-tle and the International Space Station (ISS) in orderto validate dynamics and control algorithms for minisatellites in formation. Rendezvous and docking of twoSPHERES satellites on-board the ISS has been achievedin 2006.

CanX-1 [2] is the first spacecraft developed withinthe Canadian Advanced Nanospace eXperiment (CanX)program, which was initiated by the University ofToronto in 2001. This technology demonstration space-craft was a single CubeSat and built according to the

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S. Ulrich et al. / Acta Astronautica 64 (2009) 244–255 245

California Polytechnic State University (CalPoly) andStanford University CubeSat standards [3]. Followingthe successful flight of this 10×10×10cm spacecraft, the3.5kg milk carton-sized CanX-2 [4] will be launchedin early 2008 to evaluate technologies that will be usedon the expected formation flying missions, CanX-4/CanX-5 [5].

SSETI Express, which was sponsored by the StudentSpace Exploration and Technology Initiative (SSETI)association, was successfully launched in October 2005.This first SSETI spacecraft was carrying three Cube-Sats, a propulsion module, an imaging camera and anamateur radio transponder [6]. Unfortunately, due to anunrecoverable fault in the electrical power subsystem(EPS), the mission was aborted in 2005. The next SSETIsatellite is the European Student Earth Orbiter (ESEO),which is planned to be launched in 2008 [7]. TheEuropean Student Moon Orbiter (ESMO) is the thirdSSETI spacecraft and was officially approved in 2006for a phase A study by the Education Departmentof the European Space Agency (ESA). Just like hisSSETI predecessors, ESMO will be completely de-signed, built, and operated by European and Canadianstudents in a highly distributed way. This 150kg satel-lite will play a valuable role in the international spacecommunity, namely by: (1) providing new scientificmeasurements relevant to lunar science and the futurehuman exploration of the Moon, (2) by providing flightdemonstration of innovative space technologies devel-oped under university research activities, and (3) byacquiring images of the Moon for public relations andeducation outreach purposes. Moreover, it will stim-ulate space awareness among students and give themtechnical abilities related to space-oriented projects.

The ESMO spacecraft will be launched either on-board Ariane 5’s Arianespace Support for AuxiliaryPayload (ASAP) or on-board Soyuz in 2011 from theGuiana Space Center (GSC) in Kourou. Few thousandsof seconds after the launch, ESMO will be separatedfrom the launcher with a spin rate of 5 rpm and deployedin a GTO orbit defined by an inclination (i) of 7◦, a lon-gitude of ascending node (�) of 10◦ west, an argumentof periapsis (�) of 178◦, an eccentricity (e) of 0.716,and a semimajor axis (a) of 24630km. The spacecraftwill stay in this GTO phase for one week and then,continuous tangential thrust will start at the last perigeepassage. No thrusting will occur during eclipses. Thisfirst phase of the mission ends once the apogee reaches200000km. From this point, ESMO will use 21kg ofits xenon propellant for delivering 6.5km/s of total �V,transferring ESMO to its polar lunar orbit in 1 year. Thedesired stable operational lunar orbit (SOLO) is defined

by an inclination (iL) of 90◦, an argument of perisele-nium (�L) of 295◦, an eccentricity (eL) of 0.481, and asemimajor axis (aL) of 3500km. Then, the 2kg primaryimager payload will start transmitting data to Earth fora period of approximately 6 months.

Through the last ESA’s call for proposals which tookplace in July 2006, several teams have officially beenselected to conduct a phase A study for the ESMO mis-sion. These teams and their responsibility in term ofsubsystems are defined as follows:

Power

• Université de Sherbrooke, Canada

Payload

• University of Liege, Belgium• University of Toronto, Canada

Propulsion

• University of Southampton, UK• University of Warwick, UK• Politecnico di Milano, Italy• University di Napoli, Italy

AOCS

• Narvik University College, Norway• University of Lisbon, Portugal• SUPAERO, France• Ryerson University, Canada

On-board data handling

• University of Munich, Germany

Communications

• Wroclaw University of Technology, Poland

Structure

• University of Southampton, UK

Thermal

• Politecnico di Milano, Italy

Mechanisms

• University of Porto, Portugal

246 S. Ulrich et al. / Acta Astronautica 64 (2009) 244–255

Mission analysis

• University of Glasgow, UK

Ground segment

• University of Rome, Italy (ground station Kenya)• University of Munich, Germany (ground stationGermany)

Flight dynamics

• University of Rome, Italy

Simulations

• Warsaw University of Technology, Poland

Contrary to conventional, large satellites that havesignificant power generation and storage capability,mini satellites such as ESMO do not offer the sameadvantages, being of small mass. For this reason, mostmini satellites are being operated in Earth orbits (LEOor GEO) where large �Vs and demanding communi-cation power are typically not required. Sometimes,propulsive maneuvers are not even necessary and smallthrusters are only used for attitude maneuvers. In thatcase, the required power may be less that 100W, whichtranslates a lightweight power system, which is coher-ent with the philosophy behind the small satellite world.As example, ESA’s PROBA-1 [8], a washing machine-size earth observation satellite, operates autonomouslywith only a few watts.

On the other hand, an interplanetary ESMO-likesatellite mission that uses ion engines represents a greatchallenge. Indeed, in order to generate the thrust levelsof interest for the mission, the power need is quite high,of the order of 650–750W at beginning of life (BOL).Therefore, given the strict mass constraint imposed bythe fact that most mini satellites are launched as piggy-back, a state of the art, highly efficient and lightweightEPS becomes mandatory.

This paper presents the preliminary design of the EPSfor the ESMO mission conducted by the ESMO’s EPSteam. The remainder of this paper is as follows. In Sec-tion 2, salient features of ESMO are given. In Section 3,the EPS functional architecture is outlined and in Sec-tion 4 the power budget is detailed. Sections 5–8 presentthe preliminary design of each element of the EPS (thesolar array, the secondary batteries, the power controlunit (PCU), and the power distribution unit (PDU), re-spectively). Finally, Section 9 provides the conclusion.

Fig. 1. ESMO internal view.

2. Features of ESMO

Some of the features of ESMO are illustrated inFig. 1 and are listed below:

• Dimensions: 700×700×800mm (cuboid).• Mass: 150kg.• Lunar orbit: polar orbit with a periselenium of100km.

• Structure: aluminium alloy honeycomb covered withtwo skins of CFRP.

• Payload: imaging narrow angle camera (NAC).• Electric propulsion: T5 ion engine.• Uplink: 2275MHz, 2kbps.• Downlink: 2275MHz, 2.5kbps.• Antennas: S-band omnidirectional×2.• Transmit power: 5W.• Attitude control: 3-axis actuation capability.• Thermal: passive control/radiating panel.• Mission lifetime: 18 months.

3. Functional architecture

The general layout, or the functional architecture,of the EPS is shown in Fig. 2. In this figure, fourcomponents of the EPS can be identified: (1) the solararray which generates power, (2) the PCU which condi-tions the energy, (3) the secondary batteries which area rechargeable mean of energy storage used for eclipseand peak power demand periods, and (4) the PDUwhich distributes the power to each load, or subsystem,of the spacecraft. All components other than the solararray are located inside the spacecraft body. From theconceptual design presented in a previous work [9], theCanadian team has identified that ESMO will be


Solar Array

SecondaryBatteries

Power ControlUnit

PowerDistribution

Unit

ToLoads

Fig. 2. EPS functional architecture.

powered by ultra triple-junction (UTJ) GaAs solar cellscontrolled by maximum power point trackers (MPPT).For energy storage, high efficiency Lithium (Li)-ionsecondary batteries recharged in parallel will be used. Adistributive load regulation approach is also baselined.

4. Power budget

Throughout the mission phases, different subsystemsrequire different power levels in order to operate in anominal way. Table 1 shows the maximum operatingpower requirement of every ESMO subsystems. Fromthis table, it is obvious that the main driver for the powergeneration system sizing is the power processing unit(PPU) of the T5 ion engine. Fortunately, all the sub-systems/components do not require to be turned on atthe same time during the mission. In order to assessaccurately the power consumption throughout the mis-sion, an analysis of dynamic power budgets must beperformed. Basically, dynamic power budgets are built,based on which subsystems are turned on and off for agiven period of time. Dynamic power budgets were builtfor the three critical mission phases, which were de-fined from an EPS perspective: (1) initial operations inlow Earth orbit phase (LEOP), (2) injection into SOLO,and (3) SOLO phase.

In early LEOP, several check-up and routine opera-tions must be performed, such as de-spin, sun acqui-sition, and solar array deployment. Because all theseearly operations are achieved with unfolded solar pan-els, no energy coming from the Sun can be convertedinto electricity. ESMO will then have to rely entirely onits energy storage system.

The injection into SOLO is a critical phase sincethrusting maneuvers are required to place ESMO intothe desired final lunar orbit. Orbital elements of thisphase are considered to be the same as those of theSOLO phase. During the injection into SOLO, the T5ion engine is expected to operate in its low power modeduring attitude maneuvers (where the hollow cathodethruster are used) and during eclipse periods. The max-imum power demand during this phase is of 663W.

Table 1Operating power requirement values .

Subsystem/component Power (W)

AOCSStar tracker 4.0Gyroscope 1.0Reaction wheels 12.0

CommunicationsReceiver (Rx) and RF distribution unit 6.0Transmitter (Tx) 25.0

On-board data handlingOn-board computer 22.0

PayloadNarrow angle camera 10.0

PropulsionT5 Ion engine PPU—HIGH 600.0T5 Ion engine PPU—LOW 38.0Hollow cathode thruster PPU 80.0

PowerPower control unit 7.0Power distribution unit 3.0Battery heaters 60.0

MechanismSolar array drive mechanism 9.0

Table 2Power requirement for initial LEOP operation phase .

Subphase Power (W) Duration (h)

Separation and AOS 28.0 5.00Check-up 53.0 1.00De-spin 54.0 0.50Sun acquisition 47.0 0.50Solar array deployment 63.0 0.25

Finally, the SOLO phase is quite similar to the SOLOinjection phase, except that during SOLO, the T5 ionengine is always turned off. Also, the imaging camera isassumed to be used for 30min around the periseleniumfollowed by a 1h of transmit (TX) activity around theaposelenium in order to send back to Earth the storeddata.

The details of the power requirement for each of thesethree mission phases are given in Tables 2–4 and il-lustrated in Figs. 3–5. It should be noted that in thesefigures, some parts of the graphs have been highlightedin gray to represent the eclipse periods. After a closeranalysis of these tables and figures, it can be concludedthe LEOP and the SOLO injection phases representthe sizing cases for the battery and for the solar array,


Table 3Power requirement for SOLO orbital injection phase .

Table 4Power requirement for SOLO phase .

respectively. Considering power losses, array degrada-tion, and the battery charge power with some margins,the required solar array output is estimated to 720W atBOL. The computation of the required battery capacitywill be performed in Section 6.

5. Solar arrays

The design of the solar array presents unique chal-lenges to the power system engineer since the mass and

Fig. 3. Power requirement for initial LEOP operation phase.

Fig. 4. Power requirement for SOLO orbital injection phase.

volume available are all very limited. However, in thelast years, solar cells manufacturers have greatly im-proved cell efficiency and power production. Solar cellswith greater than 28% efficiency are now commerciallyavailable. With multi-junction cells, layers of differentmaterials and doping levels are used to extract energyfrom different portions of the light spectrum, convertingmore of the spectrum into power.

For the ESMO spacecraft, the proposed solar arrayconsists of UTJ GaAs solar cells manufactured by Boe-ing Spectrolab, which have an efficiency of 28.3% atBOL and of 24.3% at EOL [10]. The UTJ GaAs cellconsists of Ge, GaAs, and GaInP2 semiconductors in-terconnected by two tunnel junctions.


Fig. 5. Power requirement for SOLO phase.

Fig. 6. I–V characteristics of the UTJ GaAs cell [10].

5.1. I–V characteristics

The performance of a solar cell at BOL is mainlycharacterized by the output voltage and current at itsterminals, which can be determined by an accuratelymeasured I–V characteristics curve of the solar cell.Fig. 6 illustrates the I–V curve for an UTJ cell, as sup-plied by Spectrolab [10].

On this curve, the two extreme points, the open cir-cuit voltage, Voc and the short circuit current, Isc are ofa great importance. Indeed, the maximum power, Pmaxthat can be generated by the cell is given by the follow-ing equation:

Pmax = CFF × Voc × Isc (1)

where the curve fill factor CFF defines the qualityof the voltage–current characteristics. From Fig. 6,

Voc is 2.665V and Isc is 17.05mA/cm2, and themaximum power point (MPP) is at 2.350V (Vmp) and16.30mA/cm2 (Imp). From Spectrolab’s data sheet[10], the CFF is 0.84. Therefore, from Eq. (1), Pmax is0.03816W/cm2.

5.2. Array design

The solar cells on the array are segmented in manystrings connected in parallel. The following equationsare used to determine the parameters needed to designthe array:

Nseries cells per string = array voltage

Vmp(2)

Nparallel panels = array current

Ipanel(3)

Nstrings per panel =IpanelImp

(4)

With a desired array voltage equal to 28V and usingthe numerical value of Vmp defined previously, Eq. (2)yields a total of 12 cells in series per string. Consideringthat six panels (three per wing) will be used on ESMO,and that the array current must be equal to 25.7 A inorder to generate the required 720W, Eq. (3) gives apanel current (Ipanel) of 4.3 A. Finally, assuming cellsize of 24cm2 (5.5×4.4cm), Eq. (4) gives 11 strings perpanel. The total size of one panel is 765×567mm, whichrespects the maximum allowed size of 800×600mm.Since the effective area of a single panel is 3146cm2,each panel will generate ∼120W for a total of 720Wfor six panels. Both sets of three rigid panels will beinitially stowed with the spacecraft during the launchand will be deployed at the end of the LEOP phase.

6. Secondary batteries

Li-ion battery cells were selected for ESMO mainlybecause they provide the best compact and weight-saving solution to ensure continuous operation of space-craft at times when the solar cells are not in sunlight orwhen peak-power demands exceed the power generatedby the solar array (namely before the solar array deploy-ment) [9]. In comparison to other type of secondary bat-teries, Li-ion batteries have the following advantages:

• Their high density of energy (twice as much as NiH2)provides a significant saving of weight (approxi-mately 30–50%). This reduction in weight leavesmore room for larger payloads.


• Li-ion batteries do not require reconditioning sincethey do not suffer from memory effect.

• Li-ion’s low thermal radiation enables the use ofsmaller battery radiators, which further reduces theEPS weight.

• Li-ion batteries retain their charge for longer pe-riod, thus do not require charging while awaiting thelaunch. This reduction in launch pad operations con-tributes to the mission cost reduction.

As explained by Patel [11], the electrical performancerequirement of the secondary batteries depends on theelectrochemistry and on many other parameters in ahighly nonlinear manner. In fact, this is why the batterydesign is one of the most difficult tasks for the powersystem engineer. However, the required capacity of thesecondary battery, in Wh, can be estimated with a sim-ple relation, as follows:

Cr = 1

nDOD

5∑i=1

Pi Ti (5)

Considering that the LEOP phase is the sizing case forthe battery, the term

∑5i=1Pi Ti represents the ideal ca-

pacity, i.e., the summation of power load (in W) timesthe duration (in h) for each of the five initial LEOP op-erations (check-up, de-spin, etc.). This ideal capacity isthen increased to include the battery-to-load transmis-sion efficiency, n, and the depth-of-discharge constraint,DOD. The DOD is defined as the percentage of the bat-tery capacity discharged during a discharge period. Therequired DOD is typically based on the type of mission.In the case of a rapid cycling LEO satellite, the DODwill be limited in order to not overstress the batteries.For a GEO satellite, it will be higher as the battery isonly used a few days per year. After a discussion withSAFT, and because only a few thousand of cycles willoccur, the maximum DOD has been fixed to 65%. Thetransmission efficiency is considered to be 90%.

Using Eq. (5) with the numerical values given inTable 2, the ideal capacity requirement during the initialLEOP operation phase is 265Wh. With the DOD andefficiency considerations, the required battery capacityis 452Wh, or 16.2 Ah at 28V.

To fulfill these requirements, the proposed batteryconsists in the flight-proven Li-ion MicroSat batterymodule, from SAFT [12]. This standard configurationbattery is assembled from MPS 176065 Li-ion cells[13], which individually operates at 3.6V and delivers5.7Ah. The resulting 8s3p configuration battery deliv-ers 480Wh, or 17Ah at 28V, which is sufficient to meetthe requirement of the ESMO mission. Some of the key

Fig. 7. SAFT Li-ion MicroSat module [12].

Table 5SAFT Li-ion MicroSat module characteristics .

Battery characteristic Value

Capacity, Ah 16.8Capacity, Wh 480.0Voltage, V 3.6Configuration 8s3pWeight, kg 4.5Height, mm 95.0Length, mm 170.0Width, mm 220.0

features of this battery include an autonomous heat-ing circuit, sensors for thermal modeling, over-chargeprotection, and built-in redundant safety protections.The complete battery module is illustrated in Fig. 7 andits characteristics are summarized in Table 5.

7. Power control unit

In order to fully optimize the power/mass ratio of themini satellite, a commercial off-the-shelf (COTS) PCUis not considered suitable for the ESMO mini satellite.Instead, a custom-built PCU tailored to the mission’sconstraints is considered as the baseline solution.

The PCU consists of the following components: (1)two maximum power point tracker units (MPPTU), (2) abattery monitor (BM), and (3) a power management unit(PMU). The internal and external interfaces between thecomponents of the PCU are illustrated in Fig. 8. Thetwo MPPTs are used to interface the power from the


From SolarPanel 1

From SolarPanel 2

MPPTU 1

MPPTU 2

PMU

Power Control Unit

BM

From/ToOBDH

From/ToSecondary Batteries

DataUnregulatedPower Bus

~ 28 V

Fig. 8. Internal and external interfaces of the PCU.

two solar arrays to the main ∼28V unregulated powerbus and the Li-ion batteries will be interfaced directlyto the main power bus through the BM, which monitorsthe state-of-charge and to prevent the overcharging ofthe batteries. The PMU basically fulfills the functionsof a telecommand and telemetry module (TC/TM). Inthis section, the design of these three PCU componentsis presented.

7.1. Maximum power point tracker unit

The MPPTU functionality is to monitor the voltageand current output of the solar array at all times, in orderto force the solar array output to operate at the MPP. Todo so, a DC–DC converter in a given configuration isfirst used. Then, with an MPPT algorithm implementedin a dedicated microprocessor, the peak power operatingpoint of the solar array on its I–V curve is detected andtracked.

7.1.1. Maximum power point tracker unitTypical DC–DC converter configurations are the

boost, the buck, and the buck–boost and a series ofhybrid versions of those. An overview of all of theseconfigurations can be found in Walker’s work [14].

As with any other EPS components, redundancyshould be limited to the minimum, in order to reducethe mass as much as possible. Therefore, typical se-rial configurations are considered too risky. Indeed, afailure of a series connected DC–DC converter wouldprevent the current to flow, creating a disconnection be-tween the solar arrays and the main power bus, causinga major EPS failure.

SCBC

Solar ArrayVoltage

Solar ArrayVoltage

Solar ArrayVoltage

SCBC

Standard DC-DC Buck Converter

Standard DC-DC Buck Converter

DC-ACInverter

DC-ACInverterFailure

~ 28 V

Fig. 9. (a) Schematic of a functional SCBC and (b) functionality ina case of a failure.

The series connected boost converter (SCBC), whichsimple principle of operation is explained elsewhere[15], overcome this problem. As illustrated in Fig. 9, ifa converter failure occurs, the current would still flowfrom the solar arrays to the power bus. The solar arrayvoltage would just not be conditioned anymore. Thatsituation can quickly become a serious problem if thesolar arrays and the secondary battery are operated atdifferent voltage levels. However, for ESMO, becausethe operating voltage of the solar arrays has been cho-sen to match the battery voltage, a converter failurewould only cause an uncontrolled charging of the bat-tery, thus slightly reducing their operational lifetime.Another great advantage of this configuration is its greatefficiency; around 95–98% in comparison to 85–90%available with the standard converter topologies [14].This translates directly into mass and costs saving. Fi-nally, the SCBC configuration is considered the base-line solution to provide a simple and reliable way toachieve a high efficiency and a fault tolerant DC–DCconversion for ESMO.

7.1.2. MPPT algorithmsNowadays, several MPPT algorithms are available

and can be considered by the power engineer. However,many of them are still under development. Every singlealgorithm has its own advantages and drawbacks hasdiscussed in Chapman’s survey on MPPT algorithms[16]. In order to ensure robust and risk-free MPPT op-erations, only well known and in-flight demonstratedMPPT algorithms are considered for this mini satellitemission to the Moon. The considered algorithms arethe so-called perturb and observe (P&O) algorithm andthe incremental conductance (IncCond) algorithm.


The fundamental operation principle of a P&OMPPTalgorithm is to generate a perturbation in the operatingvoltage of the solar array, and to observe the resultinggain (or loss) on the output power. This technique isvery similar to the hill-climbing method [17,18] whichgenerates a perturbation in the duty cycle, thus produc-ing a perturbation in the operating voltage of the solararray. The P&O algorithm is one of the easiest algo-rithms to implement, in which both current and volt-age are monitored for computation needs. The speed ofconvergence toward the MPP varies according to theperturbation frequency. Smaller steps between pertur-bations offer a faster tracking algorithm, but yields os-cillations around the MPP once reached, causing lossof power. For better performance, the P&O algorithmcan be implemented in a two-stage algorithm includ-ing a fast tracking stage and then a slow, finer trackingstage, to keep a good convergence speed and to limitthe power loss of fluctuation around the MPP [16].

The second algorithm, the IncCond algorithm, isbased on the fact that the slopes of the power–voltagearray power curve is zero at the MPP, positive on theleft of the MPP, and negative on the right, as given by

⎧⎨⎩dP/dV = 0

dP/dV > 0

dP/dV < 0

(6)

Since

dP/dV = I + V dI/dV�I + V�I/�V (7)

Eq. (6) can be rewritten as

⎧⎨⎩

�I/�V = −I/V

�I/�V > − I/V

�I/�V < − I/V

(8)

Hence, the MPP can be tracked by comparing theinstantaneous conductance (I/V) to the incrementalconductance (�I/�V). As with the P&O algorithm,the increment size determines the convergence of theIncCond algorithm speed toward the MPP and the os-cillation around it. A two-stage solution can also beimplemented to overcome oscillation problems. More-over, this two-stage solution ensures that the real MPPis tracked in case of multiple local maxima. Again,this technique must monitor current and voltage, butsince this slightly more complex method gives a betteraccuracy and efficiency compared to the P&O, thistechnique has been selected for the ESMO spacecraft’sPCU.

Fig. 10. Tamura S22P025S05 hall effect current sensor [20].

7.2. Battery monitor

Since the secondary battery is a critical componentin ESMO’s EPS, monitoring its state of health is impor-tant. This is done with the BM. Because the choice ofusing a MPPT led to the choice of an unregulated powerbus [9], the battery is directly connected to it throughthe BM. Although complex techniques such as neuralnetwork and fuzzy logic can be used for this purpose[19], the typical approach consists of using a voltageand a current sensor. This typical BM monitors the bat-tery’s voltage and current drawn (which is also used asa power bus monitor since it is directly connected toit). This type of BM enables the control of the state-of-charges of the batteries by the MPPT and prevents bat-tery overcharging (by adjusting the point of operationneeded for both solar arrays).

For efficiency reasons, the current sensor will be aCOTS hall effect current sensor (HECS). It can eitherbe an open-loop or a closed-loop sensor. Open-loopsensors can be a good choice, but closed-loop currentsensors offer higher accuracy by correcting the linear-ity and gain errors through a negative feedback. Sinceboth versions cost about the same price for a nominalcurrent specification, the decision of using closed-loopHECS for ESMO was made. Moreover, this type of sen-sor represents a simple and efficient way to monitor thecurrent source to or sink from the battery. Tamura Cor-poration produces a great variety of COTS HECS whichare commonly available, thus reducing their price. TheS22P025S05 [20], which is illustrated in Fig. 10, rep-resents an attractive and affordable baseline solution.

7.3. Power management unit

The PMU on-board ESMO will be used as a telecom-mand and telemetry module (TC/TM) to (1) decodecommands from OBDH in order to switch loads on and


UnregulatedPower Bus

~ 28 VData

Power Distribution Unit

To HighCurrentLoads

To MediumCurrentLoads

To LowCurrentLoads

Solid-StateCurrentLimiters



Fig. 11. Internal and external interfaces of the PDU.

off and (2) to send monitoring data back to OBDH.Considering the inherent complexity of this component,a COTS solution is adopted. Moreover, this solutionwill to reduce the risk of a PMU failure which wouldbe critical for the mission. The different COTS TM/TCmodules currently considered are the following: the Al-catel Space TMTC-H5 and TMTC-S3 (both from the9353 modular PCDU) [21] and the TERMA commandand monitoring (CM) module [22].

The Alcatel Space TMTC-S3 offers a low-mass(1.25kg) and compact design (340×190×25.5mm)with all the required features for the ESMO mission.Moreover, between the three considered options, it isthe only one that offers a RS422 bus interface as re-quired by the ESMO OBDH subsystem. For all thesereasons, for its ease of implementation and for its goodreliability, the TMTC-S3 PMU is selected for ESMO.

8. Power distribution unit

The PDU provides switch control and protectionbetween the main bus and the loads connected to it. Asimplified block diagram of the PDU is illustrated inFig. 11. Each load, depending on the current theydraw, are categorized as high (I> 2A), medium(2A � I � 1A), or low (I< 1A) current load. Consid-ering that the main power bus is unregulated at ∼28Vand that a distributive regulation has been adopted,each subsystem must have their own regulators.

For over-current protection, solid-state current lim-iters or fuse-relays can be used. Because of their higherflexibility, ease of testing and rearming capabilities overfuse-relays, solid-states current limiters are baselined

for ESMO. The trade-off analysis between these twoprotection devices can be found elsewhere [23]. ThePDU provides power distribution via two types of solid-state current limiters:

• Fold-back current limiters (FCLs) for powering re-ceivers, transmitters, the on-board computer, thePCU, the PDU, and essential AOCS units that are al-ways powered on (star tracker and reaction wheels).In the event of an overload, these current limitersenter in a fold-back mode.

• Latching current limiters (LCLs) which operate astelecommanded switches with current limitation pro-tection at selectable thresholds. The switch automati-cally trips off if the current limitation period exceedsa maximum time limit.

The PDU must also provide dedicated groups ofheater outlets, each outlet being controlled by a tran-sistor switch (TSW) and protected by an LCL as otherloads. However, solid-state current limiters can be quitecomplex to design (electronically and thermally). Onthe other hand, by using a COTS PDU, such as theAlcatel Space LCL and Heaters distribution module(H3 or P2) [21] or the TERMA power distribution mod-ule (PDM) [24], the design complexity drawback canbe avoided. The Alcatel Space P2 module offers nineLCL up to 10A each, one permanent-on LCL up to 5A,one FCL up to 1A and two heater group switch up to10A (which are all found to be well-suited for ESMOwhich requires 10 LCL/FCL as well as 2×6 heaters upto 3.5A). This module also offers a low-mass (1.25kg)and compact design (340×190×25.5mm) with all themain performances required. Moreover, this solutionoffers the advantage of being easily implemented withthe TMTC-S3 PMU since they are both from the samemanufacturer and are meant to be used together in amodular PCDU.

9. Conclusion

In this paper, the preliminary design of a lightweight,compact, and reliable EPS for the ESMO mini satel-lite has been presented, where the detailed EPS blockdiagram is provided in Annex A. Thanks to key com-ponents such as the ultra triple-junction (UTJ) GaAssolar cells and the Li-ion secondary battery, ESMO’sEPS will be able to provide the required energy to allsubsystems, even during thrusting and eclipse periods.Whenever possible, the EPS design for this mission tothe Moon has been performed by selecting flight-proven


solutions and technologies, increasing as much as pos-sible the reliability of the EPS, and thus of the entirespacecraft and mission.

Acknowledgments

The authors wish to thank each member of the ESMOmission for their contributions and dedication to the de-velopment of the ESMO spacecraft, and Roger Walker,from ESA, for his invaluable mentorship. Special thanksto Yannick Borthomieu, Satellite and Launcher Product

UTJ GaAsSolar Array 1

Power Control Unit

MPPTU 1

SBCSDC-DC Conv.

SBCSDC-DC Conv.

IncCondMPPT

MPPTU 2

IncCondMPPT

PMU

From/ToOBOH

UTJ GaAsSolar Array 2

HECSBM

Li-ion MicroSatBatteries

DataUnregulatedPower Bus

~ 28 V

Power Distribution Unit

T5 Ion EnginePPU

Hollow CathodeThruster PPU

Receiver

Transmitter

On-BoardComputer

Battery Heaters

Star Tracker

Gyroscope

Reaction Wheels

Narrow AngleCamera

Power Control Unit

Power DistributionUnit

Solar Array DriveMechanism

Data

Power

High Current Load

Medium Current Load

Low Current Load

LCL

LCL

FCL

FCL

FCL

FCL

FCL

LCL

FCL

LCL

FCL

FCL

LCL

Fig. 12. Detailed EPS block diagram.

Manager at SAFT, for reviewing the battery design. Fi-nally, the ESMO’s EPS team gratefully acknowledgesthe Canadian Space Agency (CSA) for the funding pro-vided through a Space Awareness and Learning Grant,which allowed the authors to participate in the 58th In-ternational Astronautical Congress in Hyderabad, India.

Appendix A. Detailed EPS block diagram

The detailed block diagram of the EPS is illustratedin Fig. 12.


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