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Acta Astronautica

Acta Astronautica 66 (2010) 1534–1543

0094-57

doi:10.1

� Cor

E-m

piergen

(F. Sant

journal homepage: www.elsevier.com/locate/actaastro

A space standards application to university-class microsatellites:The UNISAT experience

Filippo Graziani a,�, Fabrizio Piergentili b, Fabio Santoni a

a Scuola di Ingegneria Aerospaziale, ‘‘Sapienza’’ University of Rome, Italyb Diem, University of Bologna ‘‘Alma Mater Studiorum’’, Italy

a r t i c l e i n f o

Article history:

Received 19 January 2009

Received in revised form

23 November 2009

Accepted 24 November 2009

Keywords:

University satellite

Space standard

Hands-on education

65/$ - see front matter & 2009 Elsevier Ltd. A

016/j.actaastro.2009.11.020

responding author.

ail addresses: Filippo.graziani@uniroma1.it (F.

tili@unibo.it (F. Piergentili), Fabio.Santonii@u

oni).

a b s t r a c t

Hands-on education is recognized as an invaluable tool to improve students’ skills, to

stimulate their enthusiasm and to educate them to teamwork. University class satellite

programs should be developed keeping in mind that education is the main goal and that

university satellites are a unique opportunity to make involved students familiar with

all the phases of space missions. Moreover university budgets for education programs

are much lower than for industrial satellites programs. Therefore two main constraints

must be respected: a time schedule fitting with the student course duration and a low

economic budget. These have an impact on the standard which can be followed in

university class satellite programs.

In this paper university-class satellite standardization is discussed on the basis of

UNISAT program experience, reporting successful project achievements and lessons

learned through unsuccessful experiences. The UNISAT program was established at the

Scuola di Ingegneria Aerospaziale by the Group of Astrodynamics of the University of

Rome ‘‘La Sapienza’’ (GAUSS) as a research and education program in which Ph.D. and

graduate students have the opportunity to gain hands-on experience on small space

missions.

Four university satellites (UNISAT, UNISAT-2, UNISAT-3, UNISAT-4), weighing about

10 kg, have been designed, manufactured, tested and launched every two years since

2000 in the framework of this program

In the paper, after a brief overview of new GAUSS programs, an analysis of the

UNISAT satellites ground test campaign is carried out, identifying the most critical

procedures and requirements to be fulfilled. Moreover a device for low earth orbit low-

cost satellite end-of-life disposal is presented; this system (SIRDARIA) complies with the

international guidelines on space debris.

& 2009 Elsevier Ltd. All rights reserved.

1. Introduction

The Group of Astrodynamics of University of Rome‘‘La Sapienza’’ (GAUSS) established the research and

ll rights reserved.

Graziani), Fabrizio.

niroma1.it

education program UNISAT at the Scuola di IngegneriaAerospaziale, in which Ph.D. and graduate students havethe opportunity to gain hands-on experience on smallspace missions.

Four university satellites (UNISAT, UNISAT-2, UNISAT-3,UNISAT-4), weighing about 10 kg each, have been de-signed, manufactured, tested and launched every two yearssince 2000 in the framework of this program (Ref. [1,2]).

According to the definition of University-class space-craft, the UNISAT satellites were manufactured exploiting

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Fig. 1. UNISAT Satellite Launches.

Fig. 2. Unisat-3 on board DNEPR launch vehicle.

F. Graziani et al. / Acta Astronautica 66 (2010) 1534–1543 1535

terrestrial commercial off-the-shelf (COTS) components,which represent a trade-off between cost and perfor-mance. In the university satellite programs, usuallysubjected to strong economic budget restrictions,expensive space-rated state-of-the-art hardware is notgenerally used. It could lead to severe limitations inrespecting international spacecraft manufacturing stan-dards issued by international organizations such as theInternational Organization for Standardization (ISO),European Cooperation for Space Standardization (ECSS)or by space agencies such as NASA and ESA. However, theexperience achieved through the UNISAT program showsthat few international standards are applicable to themanufacturing of university class microsatellites. In fact, anumber of factors – the short lifetime expected, themanufacturing time schedule which must be withinthe two-year duration of regular university courses, theuse of COTS components which are not always easilyavailable – imply that the manufacturing of the satellitecannot follow previously fixed rigid design but must berealized by trial-and-error techniques and continuousimprovements.

Design flexibility well also suits educational aimsbecause it stimulates students and young engineers todeal with new solutions and to test innovative ideas. Thisis the main difference between industrial and educationprograms: the primary goal of industries is the spacecraftor the system functions, whereas the goal of education isto give the young engineer experience. There is the‘‘freedom to fail’’, if innovative and non-standardizedsolutions are pursued.

Of course there are standards that must always besatisfied when damage to third parties or interferencewith other satellite operations is possible. These con-straints are imposed by the launcher companies whichenforce their rules, e.g., over satellite geometry, inertialcharacteristics, the capacity to withstand launch vibrationand static loads and EMI compatibility. These constraintsconstitute the real standard rules to be followed byuniversity-class satellites.

In the paper, after a brief description of UNISATsatellites and the future GAUSS group programs????(EDUSAT and UNICubeSAT), an overview of the UNISATground test campaign is given, highlighting the mostcritical test to be executed.

After the test campaign the paper reports a descriptionof a device, developed by GAUSS, for satellite removal atthe end-of-life. This device permits compliance withinternational guidelines on space debris mitigation withrespect to after mission satellite disposal. These rulesstate 25 years as the maximum time of in orbitpermanence for LEO debris.

In conclusion, an overview of lessons learned duringthe UNISAT program is given.

2. UNISAT satellites, EDUSAT and UNIcubeSAT

In the framework of the UNISAT program, four univer-sity satellites, weighing about 10 kg each, were designed,manufactured and launched: UNISAT (September 2000),

UNISAT-2 (December 2002), UNISAT-3 (June 2004) andUNISAT-4 (July 2006) (Fig. 1).

The UNISAT-3 satellite (Fig. 1), launched on June 29th2004 from Baikonour cosmodrome (Kazakhstan), is stilloperative after more than five years from launch and thedata are received on a regularly basis from the GroundStation SPIV (San Pietro in Vincoli) located at the Scuola diIngegneria Aerospaziale of the Sapienza University ofRome. Unfortunately, the satellite UNISAT-4 did not reachthe orbit, due to launcher failure in July 2006 (Fig. 2).

The aim of UNISAT satellites is to test in orbit a numberof technological components, namely terrestrial technol-ogy solar panels (Ref. [3]), the maximum peak powertracking (MPPT) system (Ref. [4]), the commercial off-the-shelf (COTS) magnetometer and the on-board datahandling system, based on a COTS embedded system forterrestrial robotic application.

The attitude control system design and manufacturefollows the UNISAT program philosophy of fast develop-ment (about one and half years from concept to launch)and low cost.

The attitude stabilization system best responding tothe overall program requirements is the passive magnetic

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Fig. 3. Imp distribution for the solar cells used on UNISAT-3 photo-

voltaic system.

F. Graziani et al. / Acta Astronautica 66 (2010) 1534–15431536

attitude stabilization system, based on a permanentmagnet and an energy dissipation system, which, forUNISAT-3, consisted of three magnetic hysteresis rods.The satellite is not equipped with dedicated attitudesensors, except for a COTS three axes magnetoresistivemagnetometer (TAM). The attitude determination isperformed using magnetometer readings and three solarpanels telemetry data (Ref. [5]).

Thus, UNISAT satellites are characterized by extremelysimple system architecture and very fast developmenttimes.

The educational results of about 15 years of activity ofthe UNISAT program, the enthusiasm of students, theimportance of a continuous space program in the Italianuniversity to sustain applied research and the training ofyoung engineers and researchers: these factors all seemto indicate the advisability of continuing this hands-onapproach to teaching with the proposal of new student-built satellites. At the present GAUSS is developingthree new projects: UNISAT-5, UNICubeSAT and EduSAT(Ref. [6]).

UNISAT-5 is the fifth microsatellite of the UNISATprogram: it will maintain the same shape and the samesystem architecture of the previous satellites, butthe attitude stabilization system will be different, usingthe gravity gradient effect, and the communicationsystem will use S band and not only UHF and VHF band.

UNISAT and UNISAT-2 were stabilized by spin, whileUNISAT-3 and UNISAT-4 were stabilized using a passivemagnetic system. A limitation of the passive magneticattitude stabilization system is that the spacecraft is notalways oriented along the local vertical. Because of thechoice to install simple Earth observation payloads onboard UNISAT-5, the gravity gradient stabilization systemwas preferred as it provides nadir pointing by using anappropriate mass distribution of the satellite. The boomtip mass named NARCISO will be active, provided withsolar cells, small batteries, on-board computer andlow power transmitter. It can communicate with thebigger satellite and it will be used to catch images of theUNISAT-5 mother satellite from about 3 m away.

UNICubeSAT is a cube-shaped nanosatellite, 10 cm perside, weighing 1 kg. The idea to make a cubesat respondedto the challenge to design a scaled-down version, from10 kg to 1 kg, of UNISAT (UNICubeSAT) for hands-oneducation purposes. The design phase is now finished,main and critical subsystems prototypes have beenrealized and are at the test phase. The UNICubeSAT hasbeen installed with a renewed, smaller version of theBroglio drag balance for aeronomy studies providing in-situ measurement of neutral atmospheric density. Thisexperiment is well suited also for a university student-built satellite and recent advances in micro and nanoinstrumentation devices have made it possible to scaledown the original Broglio drag balance design to fit insidethe small volume available in Cubesats.

EduSAT (Educational Satellite), proposed and foundedby the Italian Space Agency, is at present under develop-ment. The satellite has the shape of a parallelepiped, witha square 31.5�31.5 cm2 base and 25 cm in height,weighing about 15 kg. The attitude system is based upon

a passive magnetic stabilization system which provides arestoring torque to align one of the satellite axes with thegeomagnetic field. Different payloads are foreseen: one ormore will be proposed and developed by high schoolstudents. Together with university students and Ph.D.students, they are involved in this project. The purpose ofmoving ‘‘hands-on-education’’ from the University toHigh Schools is to educate young generations of studentsto space related topics, teaching them disciplines that areusually so far from high school subjects and experience, inorder to stimulate their interest in scientific careers.

3. University class microsatellite test campaign

In this paragraph a description of ground test cam-paign tests carried out for UNISAT satellites is given inorder to show which are the most useful and critical testprocedures for university satellites.

3.1. Solar cells

All the solar cells were tested by using a computer-controlled multimeter and a sun-simulator. Since the lampof the solar simulator was not calibrated, it was notpossible to identify the true efficiency of the cells. Thetests were carried out in order to obtain relative valuesamong the cells. The strings were assembled by choosingthe cells on the basis of their Imp. It was mandatory toreduce the negative effect that one single cell with a lowcurrent could produce on the whole string. Below, (Fig. 3)the distribution for UNISAT-3 tested solar cells is sketched.

The distribution is almost Gaussian, moreover it showsa few bad cells which were discarded during theassembling phase.

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Fig. 4. Imp vs. Vmp distribution for the solar cells used on UNISAT-3

photovoltaic system. Imp increasing on the X-axis. Fig. 5. Imp vs. Efficiency distribution for the solar cells used on UNISAT-

3 photovoltaic system. Imp increasing on the X-axis.

F. Graziani et al. / Acta Astronautica 66 (2010) 1534–1543 1537

Remembering that the on orbit true values should becorrected by an unknown factor, it is possible to make afew considerations on the behavior of the cells. In Fig. 4the Vmp of the cells is sketched versus the increasing Imp.A casual and almost constant behavior of the Vmp arises,with Imp increasing. This shows that there is nocorrelation between the two values. The first peak andthe lowest peak indicate two bad cells. The lack ofcorrespondence between Imp and Vmp produces, as amajor consequence, a particular distribution of theefficiency versus the Imp of the solar cells (Fig. 5).It appears clear that with Imp increasing, the effici-ency of the cells increases too, in virtue of the fairlyconstant Vmp.

This last consideration implies that strings made ofcells with similar Imp will have similar efficiencies.

Fig. 6. UNISAT-3 transverse vibration test.

3.2. Vibration tests

All the UNISAT satellite vibration tests were per-formed at the Oerlikon–Contraves Italia S.p.A. facilities,in order to validate the structure, the solar panelassembling, the electronic components soldering process,the satellite structure finite elements model (FEM), and toensure that the structure fulfils the DNEPR launcherspecifications.

Vibration tests were performed along the X, Y and Z

satellite axes. Testing along the Z axis (longitudinalvibration test) was performed mounting the satellite onthe shaker using a special support, reproducing theDNEPR launcher interface, as shown in Fig. 6. The sameshaker and support, connected to the horizontal vibration

table, were used for the X and Y axes vibration test(transverse vibration).

The satellite was fixed to the support using a steel boltscrewed on the pyrobolt nut, located in the center of thelowest satellite plate. The support was secured to theshaker, or to the transverse vibration plate, by four steelbolts.

Two triaxial accelerometers were fixed to the satellitestructure to perform the vibration test measurements, asshown in Fig. 7. One is located on the upper satellite plateand aligned with the longitudinal satellite axis. The otherone is located in the center of a lateral panel aligned at anangle a=22.51 with the X axis.

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Fig. 7. Upper plate triaxial accelerometer mounting.

Table 1DNEPR Launcher specifications.

Lateral frequency 410 Hz

Longitudinal frequency 420 Hz

1st launch phase random input 6.5 g (rms) for 35 s

2nd launch phase random input 4.5 g (rms)for 14 min

Table 2Sinusoidal input parameters.

Band, Hz Amplitude, g

20y2000 0.5

Sweep rate 1 oct/min

Table 3Random vibration test acceleration spectral density.

Band, Hz Spectral density, g2/Hz

20y80 0.007

80y160 0.007y0.022

160y320 0.022y0.035

320y640 0.035

640y1280 0.035y0.017

1280y2000 0.017y0.005

rms value s (g) 6.5

Duration (sec) 35

F. Graziani et al. / Acta Astronautica 66 (2010) 1534–15431538

A monoaxial accelerometer (control accelerometer)located on the satellite support was used to check theinput signal. In this way a total of seven acquisitionchannels were used (control, Ax, Ay, Az, Bx, By, Bz).

The satellite was tested along the three axes in order tocheck that the structure fits the DNEPR launcher specifi-cations. These require that the satellite’s natural frequen-cies are above the limits listed in Table 1 and describe theDNEPR launcher vibration environment during tworandom input vibration sequences, also listed in Table 1.

The vibration test sequence for all of the three axeswas set as follows:

1.

Evaluation of the satellite’s natural frequencies usingthe sinusoidal input test.

2.

Random vibration test at the levels required by the 1stDNEPR launch phase.

3.

Sinusoidal input test as in 1) to check the effect ofrandom vibration on the satellite’s natural frequencies.

4.

Random vibration test at the levels required by the 2ndDNEPR launch phase.

5.

Sinusoidal input test as in 1) to check the effect ofrandom vibration on the satellite’s natural frequencies.

The sinusoidal input (sine sweep test) parameters arelisted in Table 2. The random signal input of the 1stDNEPR launch phase is listed in Table 3. The 2nd phasespectral density is obtained scaling the levels in Table 3down, to reach the required 4.5 g (rms).

The most significant results of the vibration test arediscussed here, taking the X transverse vibration test as anexample. Similar considerations are valid for the otheraxes Y and Z. The Ax and Bx channel data for the sinusoidal

input test (item 1 and item 3 of the test sequence) alongthe X axis are shown in Figs. 8 and 9. The effect of randomvibration is to reduce slightly the satellite’s naturalfrequencies, while the frequency response peakamplitudes do not change significantly. The first and thesecond natural frequencies are reduced from about 98 Hzand 119.5 Hz to about 94.5 Hz and 116 Hz, respectively.

The results of the Y and Z vibration tests aresummarized in Table 4. The results of the Y axis aresimilar to the X axis. The first two natural frequenciesundergo a comparable reduction from about 94 to 85 Hzfor the first mode and from 116 to 113 Hz for the secondmode. The longitudinal (Z axis) first two naturalfrequencies reduce by about 10 Hz.

Structural tests have confirmed that the constraintsimposed by DNEPR Launcher are satisfied.

The tests also confirmed the effectiveness of the solarcell gluing, the integrity of the electronic devices and thereliability of the soldering process.

3.3. Radiation pattern

Electromagnetic compatibility tests for the launcherare not required for the UNISAT satellite, because thesatellite is switched off when on board the DNEPR launchvehicle.

Radiation pattern identification tests were carried outto ensure that the satellite’s external appendices (threeprobes for measurements of ionospheric plasma charac-teristics and a deorbiting device for UNISAT-4) do notdisturb the RX antenna of the satellite, which works at145 MHz. These tests were carried out in a wide open-airfield far from buildings and metallic structures (at least100 m from the antenna system). The antenna was used astransmitter.

The satellite was placed on a 3.50 m pole (Fig. 10). TheRX measurements were made at the same elevation planeas the satellite’s equatorial plane. Fig. 11 shows the resultsof the measurements obtained in terms of an azimuth

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20 40 60 80 100 200 400 600 800 1000 1500 200010-2

10-1

100

101

Sine Ax pre-post random

Frequenza (HZ)

Acc

eler

azio

ne (g

)

Sine Ax preSine Ax post

Fig. 8. Ax channel data from sine sweeps before and after random vibration.

20 40 60 80 100 200 400 600 800 1000 1500 200010-2

10-1

100

101

Sine Bx pre-post random

Frequenza (HZ)

Acc

eler

azio

ne (g

)

Sine Bx preSine Bx post

Fig. 9. Bx channel data from sine sweeps before and after random vibration.

F. Graziani et al. / Acta Astronautica 66 (2010) 1534–1543 1539

polar diagram with 451 per step. These polar diagramsshow that the antenna transmission lobe is symmetric.The measurements were carried out with the receiver10 m and 3 m from the transmitting antenna.

3.4. Vacuum tests

The UNISAT-3 and UNISAT-4 main trays (Fig. 12),containing batteries and electronic devices, were tested in

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Table 4Pre–post random sinus sweep.

Axis Channel I Mode (Hz) II Mode (Hz)

Sinus pre

random

Sinus post

random

Sinus pre

random

Sinus post

random

X Ax �98.0 �94.5 �120 �116

Bx �97.5 �95 �119.5 �116

Y Ay �94.5 �84.5 �115 �114

By �95.0 �84 �115 �113

Z Az �145.0 �130 �265 �250

Bz �140.0 �135 �265 �255

Fig. 10. Unisat-4 mock-up during electromagnetic compatibility tests.

I serie; d = 10m

0

0.5

10

45

90

135

180

225

270

315

VIserie; d = 10m

0

0.5

10

45

90

135

180

225

270

315

Fig. 11. Polar diagrams.

Fig. 12. UNISAT-3 main tray.

F. Graziani et al. / Acta Astronautica 66 (2010) 1534–15431540

a vacuum chamber (Fig. 13), in order to check thetemperature distribution while the system operates in avacuum. The chamber is able to realize about 10�5 mbarvacuum. The test lasted for 4 h. Four temperature sensorswere installed on the batteries, on the aluminumhoneycomb skin, on the 3 V DC/DC converter and on themicroprocessor. They allowed a real-time monitoring ofthe satellite condition.

A control software was burned into the flash memory,providing a continuous check for eventual overheatconditions. The system automatically switches off, whenthe temperature exceed a preset threshold. A lightemitting diode (LED) indicates this critical condition tothe operators.

During the four hours no overheat condition wasreached.

The vacuum tests were performed mainly to test thecapabilities of COTS electronic components to operate insuch an environment. Thermal tests, although usuallyrequired or suggested by international standards, areusually too much expensive for university programs andthey were not carried out; however, it must be noted thatthe internal satellite temperature, measured by previous

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Fig. 13. Vacuum chamber.

Fig. 14. UNISAT-2 separation test (Dniepropetrovsk).

F. Graziani et al. / Acta Astronautica 66 (2010) 1534–1543 1541

UNISAT missions, ranges from between �20 up to 50degrees, thus fully compatible with the COTS industrialoperative temperature.

3.5. Fit check

The last required step to be boarded on the DNEPRlauncher is to pass the fit check. This test, required by thelauncher company, is mandatory. It is carried out bysimulating the whole of the satellite-launcher integrationand separation procedures.

A satellite mock-up exactly reproducing the geometry,the weight, the inertia moments and the separationinterface of the satellite flight unit is delivered to thelauncher company facilities, and the company is respon-sible for the test.

In this phase the integration of the satellite to thelauncher is reproduced in order to discover and to preventproblems that could arise at the launch site. Theseparation of the satellite from the launcher is alsosimulated by exploding the real pyrotechnic bolt andremoving the satellite by means of a system based onropes and counterweights (Fig. 14).

The integration procedure is standard for each launch-er; for a university program, therefore, using the samelauncher for different satellites, it would be advisable tomaintain the geometry, weight and interface of the

satellite bus frozen for every satellite. This facilitates fit-check operations without losing sight of the educationalpurpose since different students could attend differentfit-checks.

4. A device to fulfil international guidelines on end oflife disposal

Since 1957 (the Sputnik Launch), approximately20,000 tons of material have been lifted into orbit;4500 tons are still in orbit today (Ref. [7]).

For a civilization that is strongly dependant upon ourspace infrastructure, the issue of space debris is a seriousconcern. A collision with space debris could damage ordestroy a spacecraft such as a manned mission.

Orbital debris is defined as ‘‘any non-operationalobject in space around earth (Ref. [8])’’. It may be:spacecraft no longer operating, rocket bodies, launchand mission related castoffs, remnants of satellite break-ups, solid rocket exhaust, frayed surface materials. Orbitaldebris also includes paint chips, explosive bolts, springsand even a glove lost by an astronaut during ExtraVehicular Operation.

Currently NORAD track objects with a diameter greaterthan approximately 10 cm in low earth orbit (LEO) and1 m in geosynchronous orbit (GEO). Smaller debris causesa constant unpredictable collision risk.

Nevertheless the space debris concern is not limited tothe debris currently in orbit. One current projectionestimates that there will be 2000 new objects launchedinto space in the next 10 years.

Operational debris and on-orbit explosions are ex-pected to increase the debris population still more.

The ultimate solution to orbital debris lies in removingspacecraft from orbit when their operative lifetimes end.

Some years ago, the Inter-Agency Space Debris Co-ordination Committee (IADC) released guidelines (Ref. [9])with the aim of reducing the increase of space debris inprotected regions (GEO and LEO); these counter-measuresreduce the uncontrolled proliferation of space debris andmaintain a sustainable situation for the space environ-ment, reducing risks for future missions. GEO satellitesthat are no longer operative should be moved into

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Fig. 15. SIRDARIA on-board UNISAT-4

Table 5Lifetime for UNISAT-like satellites in different operative situations.

Orbit Sail effective

surface (m2)

Lifetime (yr)

650�650�651 0.09 30

650�650�651 0.15 13.4

650�650�651 0.3 8.6

650�650�651 0.6 7.48

650�650�651 1.2 6.5

800 km sunsynchronous 0.09 186

800 km sunsynchronous 0.15 117

800 km sunsynchronous 0.3 53.9

800 km sunsynchronous 0.6 30

800 km sunsynchronous 1.2 10

F. Graziani et al. / Acta Astronautica 66 (2010) 1534–15431542

graveyard orbits, LEO satellites’ mechanical lifetimeshould be reduced to less than 25 years.

As regards the LEO region, a few deorbiting systemshave been studied. Larger satellites with orbital controlsystem can use propulsion for end-of-life disposal, or anelectro dynamical tether can also be used for the samepurpose (Ref. [10]).

The paper deals with a device for reducing satellitelifetime by increasing satellite ballistic coefficient. Theproposed system increases aerodynamic resistance bydeploying a sail to enlarge the satellite area to mass ratio.

The system was developed at the Scuola di IngegneriaAerospaziale of the University of Rome and was boardedon UNISAT-4.

University satellites are educational in purpose. Thedeorbiting system fits with such aims by teaching theimportance of avoiding dangerous space debris while testinga prototypal innovative solution to mitigate space debris.

The manufactured device is less complex than othersystems and could be easily hosted on board low costmicrosatellites which usually do not have orbital controlsystems. The proposed system (SIRDARIA, SpacecraftIntegrated Re-entry Device Aero Resistant IncreasingArea) consists mainly of a triangular sail deployed bytwo booms. UNISAT-4 will board two prototypal devicesin ‘‘butterfly’’ configuration (Fig. 15).

Fig. 16. SIRDARIA, internal view and folded sail.

5. Analysis of the SIRDARIA effect on lifetime

Numerical simulations using DAS (DEBRIS ASSESMENTSOFTWARE) (Ref. [11]) were performed to estimate theefficiency of the proposed system. A total surface of 1.2 m2

was considered with respect to a UNISAT main surface of0.09 m2 and different heights were analyzed. The softwareused is also able to take solar cycles into account. In eachsimulation a total satellite mass of 10 kg was considered.

If the satellite is stabilized through a passive magneticsystem, half of the whole sail surface should be con-sidered obtaining about 0.6 m2. This is due to the fact thatduring orbit the sail will not always be perpendicular to

wind direction thus reducing its effective resisting sur-face. This case was also considered and the deorbitingsystem is still effective.

The last situation considered deals with the possibilitythat only half the sail deploys. If magnetic stabilizationexists a mean surface of 0.3 m2 should be considered. Ifonly passive aerodynamic stabilization works, an effectivesurface of about 0.15 m2 should be considered

In Table 5 expected lifetimes for various operativeconditions are sketched. The deorbiting system seems tobe truly effective, significantly decreasing the satellitelifetime. The SIRDARIA contribution is still effective evenat the height of 800 km. Situations in which the SIRDARIAcontribution is effective are highlighted.

A prototype was manufactured on the basis ofthe design depicted (Ref. [12]). The final assembledSIRDARIA prototype is shown in Fig. 16. All themechanical parts are realized in aluminum, dimensionsare 146�146�64 mm3 and it weighs about 1 kg. Thispicture shows the stowed booms, the folded sail, and theconnector between the cover and the satellite. Theopening system is in the lower right corner.

Two prototypes were boarded on the UNISAT-4satellite. They are placed on the upper trail close to thereceiving antenna.

The final satellite configuration is visible in Fig. 17where SIRDARIA assembled on board UNISAT-4 is shown.

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Fig. 17. SIRDARIA on board UNISAT-4.

Table 6Standards for university satellites in LEO.

Launch qualification Launcher

compatibility

� Fit check

� Inertial

characteristic

measurements

� Vibration and

shock test

� EMI

compatibility

Orbital phase Environmental test � Solar array

functional test

� Vacuum test

� Radiation

pattern

Post-mission

disposal

Compliance with

international

guidelines on

space debris

Passive

deorbiting

systems (e.g.

SIRDARIA)

F. Graziani et al. / Acta Astronautica 66 (2010) 1534–1543 1543

The Fig. 17 photo was taken in the Baykonour clean rooma few days before the UNISAT-4 launch, during theintegration operation.

6. Conclusions

In conclusion an overview of the standards and teststhat UNISAT satellites fulfil is given.

The standards proposed for university satellites aresummarized in Table 6.

The experience of the GAUSS group with universitysatellites is still increasing although previous missionstaught different lessons.

Lessons learned can be summarized and collected forevery satellite.

UNISAT, the first satellite, definitely demonstrated thatstudents can effectively design, realize and launch into orbit asmall, low-cost university microsatellite, providing us withbasic knowledge on the main requirements and constraintsof the satellite structure, university ground station imple-mentation, solar cell lamination and array assemblingprocess, ground test campaign and launch campaign.

UNISAT-2 was not submitted to the ground testcampaign and it did not function. This showed that anenvironmental test campaign is needed and shouldalways be carried out. Nevertheless the educational aimwas achieved since the satellite design provided us withbasic knowledge on the On Board Data Handling (OBDH)System development, and an Electronic Laboratory wasestablished at the School of Aerospace Engineering of theUniversity of Rome ‘‘La Sapienza’’.

UNISAT-3 is still operative after five years from thelaunch. This shows that low cost university microsatellitelifetime can be longer than one year. The satellite design andrealization provided us with improved knowledge on thedevelopment of OBDH and power management systems.Moreover attitude sensors were boarded on UNISAT-3 whichpermitted determination of the satellite attitude.

UNISAT-4 did not reach the orbital phase because of laun-cher failure but its design and realization provided us withknowledge on the different payloads boarded on the satellite.

Insurance being out of budget for university satellites,it is suggested that wherever possible two satellitesshould be manufactured at the same time in order tohave a back-up flight unit, useful in such situations.

Although the satellite never reached the orbital phaseit should be remembered that: ‘‘getting the studentsatellites to the launch completes 95 percent of theeducational goals of the project’’ (Ref. [13]).

References

[1] F. Graziani, F. Santoni, The UNISAT program for space education,Keynote Invited Paper, 23rd International Symposium on SpaceTechnology and Science, Matsue, Japan, May 26 –June 2, 2002.

[2] F. Santoni, F. Piergentili, F. Bulgarelli, F. Graziani, The UNISATprogram: lessons learned and achieved results, paper IAC-06-E1.1.10, 57th IAC, International Astronautical Congress 2006,Valencia, Spagna, October 2006.

[3] F. Santoni, F. Piergentili, Analysis of the UNISAT-3 solar array in-orbit performance, Journal of Spacecraft and Rockets 45 (1) (2008).

[4] F. Santoni, F. Piergentili, ‘‘Design and test of a maximum powerpoint tracking system for Unisat-3 microsatellite’’, 55th IAC,International Astronautical Congress 2004, Vancouver, Canada,4–10 October 2004, iaf paper n. iac-04-r.2.01.

[5] F. Santoni, F. Piergentili, ‘‘Unisat-3 attitude determination usingsolar panel and magnetometer data’’, 56th IAC, InternationalAstronautical Congress 2005, Fukuoka, Japan, October 2005, iafpaper n. iac-05-C1.2.06.

[6] F. Graziani, M.L. Battagliere, F. Piergentili, F. Santoni, ‘‘UNISAT-5,UNICubeSAT and EDUSAT: Three Italian projects in hands-oneducation’’, International Workshop on small satellites new mis-sions and new technologies, 5–7 June 2008, Istanbul, Turkey.

[7] Technical Report on Space Debris in 1999 by UN/COPUOS/STSC.[8] Report of the Inter-Agency Space Debris Coordination Committee

Activities on IADC Space Debris Mitigation Guidelines & SupportingDocument, Presented to the 42th Session of the Scientific andTechnical Subcommittee United Nations Committee on the PeacefulUses of Outer Space.

[9] ‘‘Support to the IADC Space Debris Mitigation Guidelines’’, Proceedingsof 22th IADC meeting, WG4, Abano Terme (Pd), Italy, 21 April 2004.

[10] C.P.L. Anselmo, T. Hanada, P.H. Krisko, ‘‘Benefits and Risks of usingElectrodynamic Tethers to De-orbit Spacecraft’’ Results of the AI19.1 Study Plan, 23rd rd IADC Meeting, 21–22 April 2005.

[11] /http://oig1.gsfc.nasa.gov/scripts/foxweb.exe/homepage@app01?tdac=NKOB4PDFG9RZXO3RSUM4S.

[12] F. Piergentili, F. Graziani, ‘‘SIRDARIA: A low-cost autonomous deorbit-ing system for microsatellites’’, 57th IAC, International AstronauticalCongress 2006, Valencia, Spain, October 2006, IAC-06-B6.4.07.

[13] Professor Bob Twiggs, CubeSat project codirector, Department ofAeronautics & Astronautics, Stanford University, California, SpaceNews, September 4, 2006.