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<strong>1st</strong> <strong>IAA</strong> <strong>Conference</strong> on
Dynamics and Control of Space Systems
DyCoSS'2012
The first in the series of events organized by the International Academy of Astronautics (<strong>IAA</strong>)
with the cooperation of the American Astronautical Society (AAS) aiming to join specialists in
dynamics and control of space systems is held in Porto, Portugal in the Hotel HF Ipanema Porto
on March 19-21, 2012.
The <strong>Conference</strong> DyCoSS’2012 is followed by a satellite event, 7th International Workshop and
Advanced School “Spaceflight Dynamics and Control” organized at the Engineering School of the
Porto University on March 22-23, 2012 (see the reverse side of this booklet).
The conference focus on topics related to analysis of space systems, namely:
• Attitude dynamics and control
• Attitude sensors and actuators
• Mission design and optimization
• Optimal control in space flight dynamics
• Orbital dynamics and determination
• Satellite constellations and formation flying
• Space structures and tethers
• Spacecraft guidance, navigation and control
During the conference, the complete manuscripts will be available on the conference web page.
The conference proceedings will be published by Univelt, Inc. in AAS Advances in the
Astronautical Sciences Series.
A selection of the best papers will be recommended for publishing, after peer review, in a
Special Issue of the journal Acta Astronautica.
The working language for the conference is English.
REGISTRATION FEES:
Early Bird (registered and paid before January 30, 2012) 300€
Regular (registered and paid from January 31, 2012 to March 5, 2012) 350€
On-site Registration (only in cash and in euros) 400€
The registration fee includes:
• Access to conference sessions
• CD with <strong>Conference</strong> Proceedings (published by UNIVELT)
• <strong>Conference</strong> Dinner, lunches and coffee breaks as scheduled in the conference
program
Registration fee for Accompanying person: 150€
The registration fee for accompanying person includes:
• <strong>Conference</strong> Dinner, lunches and coffee breaks as scheduled in the conference program
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Program and Organizing Committees
FINAL PROGRAM
Co-chairs:
Peter M. Bainum, USA, Jean-Michel Contant, France, Anna D. Guerman, Portugal
Members:
Kyle Terry Alfriend, USA
Hyochoong Bang, Korea
Franco Bernelli, Italy
James Biggs, UK
Jeng-Shing Chern “Rock”, Taiwan
Bernd Dachwald, Germany
Harijono Djojodihardjo, Malaysia
Antonio Elipe, Spain
Margarida Ferreira, Portugal
David Folta, USA
Ijar Fonseca, Brazil
Gerard Gomes, Spain
Filippo Graziani, Italy
Christopher D. Hall, USA
Kathleen Howell, USA
Junichiro Kawaguchi, Japan
Ivan Kosenko, Russia
Richard Longman, USA
Robert Melton, USA
Arun Misra, Canada
Yasuhiro Morita, Japan
Michael Ovchinnikov, Russia
Maria Rosario Pinho, Portugal
Antonio Bertachini Prado, Brazil
Georgi Smirnov, Portugal
David B. Spencer, USA
Paolo Teofilatto, Italy
Zhang Wei-Hua, China
Local Organizing Commitee
Anna D. Guerman, University of Beira Interior, Portugal, coordinator
Margarida Ferreira, University of Porto, Portugal, coordinator
Maria Rosario Pinho, University of Porto, Portugal
Miguel Ângelo Silvestre, University of Beira Interior, Portugal
Georgi Smirnov, University of Minho, Portugal
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1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
PROGRAM AT A GLANCE
CONFERENCE SCHEDULE
Room Ipanema I Ipanema II Biombo
Monday, 19.03
08:30-10:00
10:00-12:30
12:30-14:00
REGISTRATION
OPENING SESSION - ROOM PORTO
LUNCH
14:00-17:20 ADC 1 ODD 1 SGNC 1
17:30-18:00 FP7 SPACE
19:00-21:30 <strong>Conference</strong> Dinner
Tuesday, 20.03
08:30-12:30 ASA MDO 1 SCFF 1
12:30-14:00 LUNCH
14:00-18:00 ADC 2 ODD 2 SGNC 2
Wednesday, 21.03
08:30-12:30 SST MDO 2 SCFF 2
12:30-14:00 LUNCH
14:00-18:00 ADC 3 OCSFD SGNC 3
ADC • Attitude dynamics and control
ASA • Attitude sensors and actuators
MDO • Mission design and optimization
ODD • Orbital dynamics and determination
OCSFD • Optimal control in space flight dynamics
SCFF • Satellite constellations and formation flying
SGNC • Spacecraft guidance, navigation and control
SST • Space structures and tethers
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FINAL PROGRAM
City of Porto
Porto is the second largest city in Portugal and
one of the major urban areas in the Iberian
Peninsula. Porto is the largest city of Northern
Portugal. The Historic Centre of Porto,
classified as World Heritage by UNESCO, is one
of the most attractive entertainment venues
for visitors, providing a rich variety of
monuments and ordinary dwelling, from
different periods stretching back as far as the
14th century. The city of Porto and its river,
Douro, cannot be dissociated. The Douro valley,
with its lovely terraces of vineyards clinging to
the hillsides, is the kingdom of the famous Port
Wine (IVDP Web site) and offers visitors
spectacular views. Porto is the centre of a culturally and naturally rich region that gathers
together history, arts and nature (sea and mountains); it is a perfect starting point for tourist
excursions.
The City of Porto has one of the richest artistic, cultural and historical heritages in Portugal.
This was officially recognized by UNESCO which considered Porto as "World Heritage".
Porto is also a city where its history can be found in a small street or in a dialogue with its
inhabitants, so we invite you to explore it.
Porto Wine is world-famous and has left its mark in
Porto. Although this wine comes from the schistose
lands of the High Douro region, it is brought to age in
Vila Nova de Gaia, in the semi-darkness of the
cellars, in oak casks or in bottles.
This precious nectar has endured and, as time goes
by, it improves its qualities. Visiting the Porto Wine
Cellars is a must: the various companies offer an unique opportunity to discover this universe of
subtle charms, to learn the secrets of this generous wine and to taste it.
Many optional programs are available to visit Porto and its surroundings. A sample of these
programs is presented.
Porto also offers a wide range of restaurants with traditional Portuguese cuisine or more
modern and international flavors.
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Visiting Porto
In recent years, UNESCO recognized Porto
historic center as a World Heritage Site. Among
the architectural highlights of the city, the
Porto Cathedral is the oldest surviving
structure, together with the small Romanesque
Church of Cedofeita, the Gothic Igreja de São
Francisco (Church of Saint Francis), the
remnants of the city walls and a few 15thcentury
houses. The baroque style is well
represented in the city in the elaborate gilt
work interior decoration of the churches of St.
Francis and St. Claire (Santa Clara), the
churches of Mercy (Misericórida) and of the
Clerics (Igreja dos Clérigos), the Palace of the Archbishop, and others. The neoclassicism and
romanticism of the 19th and 20th centuries also added interesting
monuments to the landscape of the city, like the magnificent Stock
Exchange Palace (Palácio da Bolsa), the Hospital of Saint Anthony,
the Municipality, the buildings in the Avenida dos Aliados, the tileadorned
Sao Bento Train Station and the gardens of the Crystal
Palace (Palácio de Cristal). A guided visit to the Palácio da Bolsa,
and in particular the Arab Room, is a major tourist attraction.
Porto has several museums, concert halls, theaters, cinemas, art
galleries, libraries and book shops. The best-known museums of
Porto are the Soares dos Reis National Museum (Museu Nacional de
Soares dos Reis), which is dedicated especially to the Portuguese
artistic movements from the 16th to the 20th century, and the
Museum of Contemporary Art of the Serralves Foundation (Museu de
Arte Contemporânea). The city has concert halls of a rare beauty
and elegance such as the Coliseu do Porto by the Portuguese
architect Cassiano Branco; an exquisite example of the Portuguese decorative arts, it is seen by
many as one of the best and most elegant concert halls in Europe. Other notable venues include
the Rivoli theatre, the Batalha cinema and the recent Casa da Música. The Guardian elaborated
a list of world's top bookshops, featuring Lello in third.
Taylor’s Port Wine Cellars
Taylors’s Port wine cellars is one of the many
Port Wine Cellars in Vila Nova de Gaia, the city
on the other side of the river Douro in Porto.
Surrounded by leafy quarters, resistant to the
city's approaches, Taylor’s Porto Wine Cellars –
established in 1692 and ones of the few that still
belong to the descendants of the founding
Families - match elegant interiors within
beautiful natural gardens; an ambiance which
lends luxury to grand banquets and intimacy to
family celebrations, not to mention the
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FINAL PROGRAM
appropriate environment for business entertainment and meetings. Its Kitchen allows to delight
proposals of Portuguese and international gastronomy, while we are able to enjoy the one that
is considered the most spectacular sight for the old Port and World- wide Heritage declared
zone for UNESCO.
Douro Cruise
Boarding at Gaia Quay for a Douro Cruise will provide a “different perspective” of the historical
zone of the city of Porto. The cruise takes approximately 3 Hours and will make the passage
between the Freixo Bridge and the zone of the Afurada.
Porto Wine Museum
The Port Wine Museum is settled on the ground floor of
the Cais Novo warehouse, a building from the
seventeenth century which was built to be the wine
deposit of the Companhia Geral da Agricultura das
Vinhas do Alto Douro (General Company of Agriculture
of the Alto Douro Vineyards). This museum is
dedicated to the study of the relevance of Port Wine
and its trade to the development of the city. Porto's
history cannot be dissociated from the history of the
wine that made the city famous worldwide. It conditioned the city's cultural, architectural and
social growth. The Museum is intended to be a memory of all those influences and
transformations, and at the same time, it conveys the history of the three Portos brought
together: Porto the city, Porto (Port) the wine, and Porto the Douro river harbor.
Casa da Música
Casa da Música was conceived to mark 2001, the year
in which Porto was Cultural Capital of Europe, and it
is the first new building in Portugal to be entirely
dedicated to music - to the presentation and public
enjoyment of music, to music education and to the
creation of music.
The project took shape in 1999 after the Rem
Koolhaas & Ellen van Loon - Office for Metropolitan
Architecture won the international architectural
competition. Work began in 1999 on the site of
Porto's former central tram garage on the Rotunda da
Boavista, and Casa da Música opened its doors to the
public on April 15th, 2005.
Casa da Música was planned as a home for all types of
music and it is not only part of the urban
redevelopment of Porto but also part of a network of
cultural facilities, for the city of Porto and for the
wider world. Underlying it is an innovative and wideranging
cultural project, which aims to make an
exciting contribution to the national and international
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1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
music scene, as an arena for all types of musical events - from classical music to jazz, from
fado to electronic music, from great international productions to more experimental projects.
Besides concerts, recitals and other types of performance, Casa da Música also organises events
for musicians and musicologists and invests in research into the origins of Portuguese music. In
addition, it plays a very important role in music education. It is a cultural meeting point
between music and other areas of artistic creation and knowledge, providing a space for all
kinds of audiences and creators.
Located on the top floor of Casa da Música, the
restaurant is the only room with access to the outdoor
terrace. It is a large seven-meter high room with
restful colours and excellent cuisine.
The best way to get to know Casa da Música is to take
a guided tour. These tours, led by guides who describe
and explain the building, its construction, its
architecture and its function last approximately one
hour.
GUIDED TOURS: Daily
Portuguese 11:00 | 16:00; English 16:00
Booked visits (groups) 15-35 people at pre-arranged times
Booking Ph.: 220 120 210
visitasguiadas@casadamusica.com
4€/person (free entry to children up to age 12, once accompanied by an adult with ticket)
<strong>Conference</strong> Venue
Hotel Ipanema Porto 4*
Located at the main entrance to the city (near the
Arrábida bridge), the HF Ipanema Porto is
surrounded by shopping centers, cinemas, theatres
and the Casa da Música music Hall and is very near
to the city`s main sports stadium, the Pavilhão Rosa
Mota. The historic part of the city, São Bento railway
station and the Port wine cellars are only 10 minutes
away. With a longstanding presence in the city, the
HF Ipanema Porto has recently been renovated and
redecorated to provide a more modern, comfortable
and practical atmosphere.
Rua do Campo Alegre, 156/172
4150-169 Porto
Phone: + 351 226 075 059
Fax: + 351 226 063 339
Email: hfipanemaporto@hfhotels.com
Site: www.hfhotels.com/hfipanemaporto
GPS coordinates: 8º 37' 48" O / 41º 09' 09" N
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How to get there
From the airport to the hotel
By metro:
FINAL PROGRAM
In the Metro do Aeroporto metro station, take the line E in towards the Estádio do Dragão metro
station. Exit in Casa da Música (HF Ipanema Porto is 1 Km away)
By bus:
Near the airport, take the bus no. 601 (in direction to Cordoaria), and get off at Casa da Música
(HF Ipanema Porto is 1 Km away).
From downtown to the hotel
By metro:
Leaving from the Avenida dos Aliados, continue by going up towards the Trindade metro station
(behind the City Hall/Câmara Municipal do Porto) and take the line A, B, C, or E, towards Casa
da Música.
By bus:
Leaving from the Avenida dos Aliados, take the bus no. 200 (towards Castelo do Queijo) or the
bus no. 207 (towards Foz-Mercado) at the Praça D. João I (next to the Teatro Rivoli). Get off
the bus at the Massarelos bus stop (the HF Ipanema Porto is located within 50m).
By car:
Leaving from the Avenida dos Aliados, go up the street until the traffic lights, and turn left to
Rua de Ceuta, then entering the Túnel de Ceuta. Once in the tunnel, continue towards Hospital
Sto. António. In the exit of the tunnel take the right, passing through the Palácio de Cristal up
to Praça da Galiza. At the traffic lights turn left to Rua do Campo Alegre (the HF Ipanema Porto
is 500 metres away.)
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1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
At the hotel
Rooms
Ipanema I, Ipanema II and Biombo conference rooms are located in floor -1 of Ipanema Porto
Hotel
Ipanema Porto Floor -1 map:
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FINAL PROGRAM
DyCoSS Program
Monday, March 19, 2012
Opening Session (<strong>Conference</strong> Room "Porto") 10:00-12:30
Opening Ceremony:
Prof. Kathleen Howell, Purdue University, the Editor-in-Chief of the Journal of the
Astronautical Sciences, American Astronautical Society
Dr. António Neto da Silva, President of Proespaço, the Portuguese Association of Space
Industries
"PORTUGUESE NATIONAL STRATEGY FOR THE SPACE INDUSTRY - THE PROESPAÇO
PROPOSAL"
Dr. Jean-Michel Contant, Secretary General, International Academy of Astronautics
"THE INTERNATIONAL ACADEMY OF ASTRONAUTICS AND ASTRODYNAMICS: FIFTY YEARS
OF EXCELLENCE"
Keynote lecture:
Dr Constantinos Stavrinidis, Head of Mechanical Engineering Department, ESA-ESTEC
"DYNAMICS AND CONTROL OF SPACE SYSTEMS"
Lunch 12:30-14:00
Session 01:
Attitude Dynamics and Control I ( Ipanema I)
Chairs: Richard LONGMAN and Yasuhiro MORITA
14:00-17:20
14:00
THERMAL ANALYSIS OF THE PIONEER SPACECRAFT
AND THE ANOMALOUS ACCELERATION
<strong>IAA</strong>-AAS-DyCoSS1-01-01
Orfeu Bertolami, Frederico Francisco,
Paulo Gil, and Jorge Páramos
We present a method to evaluate the on-board thermal effects on the Pioneer spacecraft
and to compute the resulting acceleration. It is shown that anomalous acceleration problem
is likely to be resolved.
14:20
CLOSED FORM INTEGRATION OF THE HITZL-BREAKWELL
PROBLEM IN ACTION-ANGLE VARIABLES
<strong>IAA</strong>-AAS-DyCoSS1-01-02
F. Javier Molero, Sebastián Ferrer, and Martin Lara
As an alternative to recent efforts in giving a complete solution to the attitude propagation
of a tumbling triaxial satellite under gravity-gradient, we reformulate the problem in actionangle
variables. The new solution is computed by the Lie-Deprit approach and is given in
closed form, either for the secular or periodic terms, therefore being valid for any triaxial
satellite. Comparisons with other approaches in the literature using non-action-angle
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1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
variables show the efficiency of the new solution for a variety of test cases.
14:40
<strong>IAA</strong>-AAS-DyCoSS1-01-03
THE DYNAMICS AND CONTROL OF AXIAL
SATELLITE GYROSTAT OF VARIABLE STRUCTURE
Vladimir Aslanov
This paper presents the study of the dynamics and control of an axial variable structure
satellite (asymmetric platform and an axisymmetric rotor). Inertia moments of the rotor
change slowly over time. The dynamics of the satellite is described by using ordinary
differential equations with Andoyer-Deprit canonical variables. For undisturbed motion, the
stationary solutions are found, and their stability is studied. The control law is obtained for
the satellite with variable structure on the basis of the stationary solutions. By means of
computer numerical simulations, we have shown that the motion of the satellite controlled
by founded internal torque is stable.
15:00
STABILITY DIAGRAMS OF THE CYLINDRICAL EQUILIBRIA
CONCISE FORM OF THE DYNAMIC AND KINEMATIC SOLUTIONS
OF THE EULER-POINSOT PROBLEM
<strong>IAA</strong>-AAS-DyCoSS1-01-05
Marcello Romano
Known exact solutions of the Euler-Poinsot problem are presented in an original compact
form, which is appealing for education and engineering practice. Solutions for the dynamics
(angular velocity components) and for the kinematics (elements of the rotation matrix) are
given in two theorems. In particular, the dynamic solution initially proposed by Jacobi is
formulated in a universal way, which does not contain any branching in cases. Furthermore,
the rarely cited kinematic solution of Jacobi is re-written in modern notation. The paper is
intended to contribute to increase the familiarity and practical utilization of the exact
solutions of the Euler-Poinsot problem.
15:20
ON STABILITY OF RELATIVE EQUILIBRIA
FOR THE PENDULUM ATTACHED TO EQUATOR
<strong>IAA</strong>-AAS-DyCoSS1-01-06
Alexander Burov and Ivan Kosenko
An existence and stability of relative equilibria for a pendulum attached to equator are
under investigation. It is known that inverted pendulum of small length is unstable: its
degree of instability is equal to two. From the space elevator dynamics one knows that the
vertical equilibrium of the inverted pendulum is stable if the pendulum free end is located
behind the Earth stationary orbit, and the degree of instability is equal to zero. It turns out
there exists a length of the pendulum where new family of spatial equilibria arises, and the
degree of instability grows from one to two.
Coffee-break 15:40-16:00
16:00
<strong>IAA</strong>-AAS-DyCoSS1-01-07
STABILITY OF SPINNING SATELLITE UNDER AXIAL THRUST
Frank Janssens and Jozef van der Ha
This paper considers a spinning rigid body and a particle with internal motion under axial
thrust. This model studies the coning anomaly that occurred near the end of orbit injections
performed by STAR-48 rocket motors. We investigate the stability of this system by means of
linearized equations about a uniform spin. The resulting stability condition defines a
manifold in the parameter space which provides valuable new insights. For example, the
envelope of the stationary solutions is identical to the stability curve. The application of the
model is demonstrated using the orbit injection data of ESA’s Ulysses satellite in 1990.
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FINAL PROGRAM
16:20
MODES OF MOTION OF SOYUZ ORBITAL STAGE AFTER
PAYLOAD SEPARATION AT CARRYING OUT THE
SHORT-TERM RESEARCH EXPERIMENTS
<strong>IAA</strong>-AAS-DyCoSS1-01-09
Igor Belokonov, Alexandr Storozh, and Ivan Timbay
The uncontrolled motion of Soyuz orbital stage after spacecraft separation is considered.
Two variants of motion are analyzed: using and without using of the jet nozzle for stage
withdrawal from the spacecraft. Stochastic model of initial conditions of angular motion are
generated. The influence of the gravitational and aerodynamic moments on the motion
around orbital stage mass center is considered. Orbital stage movement features shown in
preservation of certain angular orientation during time, sufficient for the successful decision
of navigation-communication problems at carrying out of short-term research experiments on
the Soyuz carrier rocket orbital stage after payload separation are revealed.
16:40
<strong>IAA</strong>-AAS-DyCoSS1-01-10
ON-BOARD AUTONOMY FOR ISRO
GEOSYNCHRONOUS SPACECRAFT
Vinod Kumar, Avinash Kumar Kulakarni, K Parameshwaran,
Ramalingam Pandiyan, and Narinder Kumar Malik
This paper will discuss development of the on-board autonomy i.e. Fault Detection Isolation
and Re-configuration (FDIR) & Emergency Sun Re-acquisition (ESR) for the ISRO
Geosynchronous communication spacecraft. The communication satellites are required to
provide uninterrupted communication services. To ensure continuous availability of these
services and prevent loss due to failure in AOCS equipments has led to development of the
self healing FDIR functions to survive the space environment. The FDIR provides quick
reaction to reconfigure the spacecraft automatically in case any AOCS equipment
malfunctions to prevent any possible attitude loss.
Session 02:
Orbital Dynamics and Determination I (Ipanema II)
Chairs: James BIGGS and Filippo GRAZIANI
14:00-17:20
14:00
ACCELEROMETER DATA HANDLING
FOR THE BEPICOLOMBO ORBIT DETERMINATION
<strong>IAA</strong>-AAS-DyCoSS1-02-01
Elisa Maria Alessi, Stefano Cicalo, and Andrea Milani
In this work, we will present the constrained multi-arc strategy we developed to determine
the orbit of the Mercury Planetary Orbiter of the BepiColombo mission to Mercury and
estimate the gravimetry, rotation and relativity parameters at the level required by the
corresponding radioscience experiments. Due to the severe radiation environment the probe
will experience, we have to include in the `solve for quantities' desaturation maneuvers and
accelerometer digital calibrations. The latter kind of quantity is needed to absorb the
systematic noise associated with the non gravitational accelerations that the Italian Spring
Accelerometer will measure.
14:20
AN ALGORITHM FOR ORBIT DETERMINATION OF NAVIGATION
SATELLITES BASED ON A GROUND BEACON
<strong>IAA</strong>-AAS-DyCoSS1-02-02
Zhiguo Chen, Pei Chen, and Chao Han
A beacon-based orbit determination algorithm for navigation satellites is proposed. In this
strategy, signal is transmitted from a ground beacon station to navigation satellites. An
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1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
extended kalman filter (EKF) is utilized for the orbit determination algorithm. The state
process and the observation model are derived to establish the filter model. Based on one
ground beacon a one-day simulation is designed to respectively determine orbit of three GPS
satellites in different orbital planes. According to results of the simulation errors in estimate
of position for satellites are less than 10m, and errors in estimate of clock bias are less than
10ns.
14:40
<strong>IAA</strong>-AAS-DyCoSS1-02-03
TOWARDS AUTOMATED DETERMINATION
OF ORBIT MANOEUVRES FOR GNSS SATELLITES
Gottlob Gienger and Filipe Lopes Pereira
ESOC's navigation support office routinely performs precise orbit determination for GPS,
GLONASS and Galileo navigation satellites. To process >300 station passes per day requires:
Automated search of station passes for orbit, manoeuvres; Precise location of their start and
end times Automated Delta-V estimation. Manoeuvre models were exploited with
Mathematica. The best performing algorithms were implemented into NAPEOS, the
Navigation Package for Earth Observation Satellites. Preliminary tests conducted on several
GPS orbit manoeuvres show very good agreement with the values published by Center for
Orbit Determination in Europe.
15:00
<strong>IAA</strong>-AAS-DyCoSS1-02-04
AN IMPROVED DRAG MODEL
FOR CBERS ORBIT DETERMINATION AND PROPAGATION
Diogo Merguizo Sanchez and Valdemir Carrara
This paper presents a comparison between the current orbit determination and propagation
procedures of CBERS (China-Brazil Earth Remote Sensing) satellite, based on a fixed drag
coefficient, and a model that uses the kinetic theory of gases to compute the forces and
torques on a satellite with a given geometry. The main goal is to retrieve eventual
discrepancies between models, and to check the orbit determination residues in both cases.
The results have shown that there were some significant improvements in orbit
determination. It is expected that the variable drag coefficient can be used in next CBERS
missions.
15:20
<strong>IAA</strong>-AAS-DyCoSS1-02-05
A NEW METHOD OF ORBIT DETERMINATION
BASED ON MODEL ERROR COMPENSATION
Pan Xiaogang, Wang Jiongqi, and Zhou Haiyin
The precision of orbit determination depends on precision of measurement data and the
precision of dynamic model. In modern times, the degree of measurement can be limited in
millimeter degree, so the key method to improve the precision of orbit determination is to
increase the exactness of dynamic model, or to compensate the model error. In this chapter,
some mathematic method are proposed to compensate the uncertainty model error, all
those method is focus on the mathematics models, at last, a new orbit determination
method based on model error compensation is put forward to deal with directly the dynamic
force.
Coffee-break 15:40-17:20
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16:00
<strong>IAA</strong>-AAS-DyCoSS1-02-07
FINAL PROGRAM
ON THE SEQUENTIAL SOLUTION OF KEPLER EQUATION
Manuel Calvo, Antonio Elipe,
Juan Ignacio Montijano, and Luis Randez
The solution of elliptic Kepler equation is a classical problem of Celestial Mechanics with a
long history. Most of the methods proposed in the literature are one-time methods in the
sense that for a given eccentricity and mean anomaly they attempt to provide an
approximation of the unique solution of Kepler equation with maximum accuracy and
minimum computational cost. The aim of this paper is to propose new sequential solution
methods by using the differential equations satisfied by the eccentric anomaly together with
suitable one-step numerical methods for differential equations. Further, numerical
experiments the existing methods will be presented.
16:20
<strong>IAA</strong>-AAS-DyCoSS1-02-08
ORBITAL MANEUVERS WITH SINGLE-INPUT CONTROL
Anna Guerman and Georgi Smirnov
We study the problem of orbital control under the constraints on the thrust direction for a
non-linear dynamical model. We consider a satellite equipped with a passive attitude control
system providing one-axis stabilization and a propulsion system consisting of one or two
thrusters oriented along the stabilized axis. Different applications, such as formation flying,
station-keeping, and low energy orbital transfer, are examined. We address the problem of
satellite deorbiting and suggest some orbital maneuvers to be used to reduce the orbit
lifetime of satellites equipped with a propulsion system with its thrust axis fixed in the body
of satellite.
16:40
DYNAMICS AND STABILITY OF SPACECRAFT RELATIVE MOTION
WITH INTER-SATELLITE ELECTROMAGNETIC FORCE
<strong>IAA</strong>-AAS-DyCoSS1-02-09 Huan Huang, Leping Yang, Yanwei Zhu, and Yuanwen Zhang
Electromagnetic force differs from traditional thrust in its nature as an internal force. This
paper explores a new modeling approach of spacecraft relative motion which could embody
the characteristics of inter-satellite electromagnetic force based on multibody dynamics
theory. Firstly, a dynamics model based on two-body system in the presence of
electromagnetic force is derived following the method outlined by Kane. Secondly, the
dynamics and stability of the proposed model are analyzed. The universal dynamics model
for N-spacecrafts formation system with electromagnetic force is established at last, and the
stability analysis of dynamic equilibria and the simulation are presented respectively.
17:00
ORBIT-ATTITUDE PERTURBATION OF A CHARGED SPACECRAFT
IN THE GEOMAGNETIC FIELD
<strong>IAA</strong>-AAS-DyCoSS1-02-10
Hani Mohammed, Mosatafa Ahmed, and Ashraf Owis
In this work we investigate the orbit-attitude perturbations of a rigid spacecraft due to the
effect of different forces and torques. The spacecraft is considered to be equipped with a
charged sheet of a cylindrical shape and thus affected upon by the Lorentz torques
generated from the geomagnetic field. The effect of the gravitational field up to J_2 and the
Lorentz force are taken into account when formulating the orbit-attitude Hamiltonian. The
problem is then solved using the Lie-Deprit-Kamel perturbation technique. In this technique
we make two successive transformations to eliminate the short and long periodic terms from
the Hamiltonian.
14

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
Session 03:
Spacecraft Guidance, Navigation, and Control I (Biombo)
Chair: Hyochoong BANG
14:00-18:00
14:00
<strong>IAA</strong>-AAS-DyCoSS1-03-01
HYBRID PROPULSION TRANSFERS TO THE MOON
Giorgio Mingotti, Francesco Topputo, and Franco Bernelli-
Zazzera
This work analyzes special Earth-Moon transfers that make use of both chemical and solar
electric propulsion. A first high-thrust, low-Isp impulse is used to place the spacecraft into
an exterior-like low-energy transfer to the Moon, possibly performing a lunar gravity assist.
The subsequent use of low-thrust, high-Isp propulsion makes it possible to perform a lunar
ballistic capture leading to a final, low-altitude orbit about the Moon. Hybrid propulsion
transfers outperform both the chemical transfers (Hohmann, interior, and exterior) and the
fully solar electric propulsion transfers (SMART-1-like) in terms of propellant consumption,
although an assessment at system level is still missing.
14:20
<strong>IAA</strong>-AAS-DyCoSS1-03-02
ATTITUDE CONTROL SYSTEM
FOR THE NERVA LAUNCHER AND THE ADDASAT
Radu Rugescu
The space launcher NERVA is the Romanian research and development program for creating
the cheap micro-orbital launcher under the sponsorship of the Romanian Ministry of
Education, Research and Innovation through national grants. The point reconversion of
military obsolete weapons is developed, with the flight test of the guidance inertial platform
performed at the test range Cape Midia in June 2010 and June 2011. The attitude control
system for the NERVA launcher and the ADDASAT is presented, with emphasize on the first
flight tests of the inertial platform and the laboratory tests of the first MRG-001 microthruster.
14:40
<strong>IAA</strong>-AAS-DyCoSS1-03-03
NONLINEAR CONTROL OF LEADER-FOLLOWER FORMATION
FLYING
Giuseppe Di Mauro, Pierluigi Di Lizia,
Roberto Armellin, and Michèle Lavagna
This paper addresses the problem of relative motion control involved in a leader-follower
formation keeping mission. Particularly, center of mass dynamics of two Earth orbiting
satellite is modeled, including the nonlinearity due to Earth’s oblateness. Next, differential
algebra is exploited to compute a high order Taylor expansion of the State-Dependent Riccati
Equation (SDRE) sub-optimal solution with respect to the spacecraft state. The polynomial
approximation can be directly evaluated for new SDRE solutions or exploited to define
accurate initial guesses for the iterative SDRE algorithm. Thus, the computational cost of the
online ARE solution required by SDRE algorithm is significantly reduced.
15:00
<strong>IAA</strong>-AAS-DyCoSS1-03-04
15
DISTANT ORBITS AROUND THE MERCURY
Xue Ma, Junfeng Li, Hexi Baoyin,
Peng Zhang, Jingyang Li, and Yang Chen
For some low cost space missions to Mercury, distant orbits around the planet are studied.
This research is carried out in the framework of the elliptic restricted three-body problem
(ER3BP), because of the planet's non-negligible eccentricity. Two families of prograde and
retrograde orbits in the CR3BP are chosen for their good performance of Sun-synchronism
and stability, and we extend the study of these orbits to the ER3BP case. The stability of all

FINAL PROGRAM
above orbits around Mercury is analyzed in this paper. Our research indicates that orbits
within a large altitude range may keep certain shapes long enough for space missions.
15:20
SINGLE-AXIS POINTING OF A MAGNETICALLY
ACTUATED SPACECRAFT
<strong>IAA</strong>-AAS-DyCoSS1-03-05
Giulio Avanzini, Emanuele Luigi de Angelis,
and Fabrizio Giulietti
The use of magnetic actuators on Low Earth Orbiting spacecrafts represents an attractive
solution because of its simplicity, reliability and a power-effective smooth modulation of
control actions. Attitude regulation yet proves to be a challenging problem, since magnetic
actuators alone do not allow providing three independent control torque components at each
time instant.
The goal of this paper is to derive a rigorous proof of asymptotic stability for a magnetic
control law that leads the satellite to a desired spin condition around a principal axis of
inertia, pointing the spin axis towards a target direction in the inertial reference frame.
Coffee-break 15:40-16:00
16:00
NONLINEAR FILTERING METHODS FOR SPACECRAFT
NAVIGATION BASED ON DIFFERENTIAL ALGEBRA
<strong>IAA</strong>-AAS-DyCoSS1-03-06
Monica Valli, Roberto Armellin,
Pierluigi Di Lizia, and Michèle Lavagna
The nonlinear filtering problem plays an important role in various space-related applications
and especially in orbit determination and navigation problems. Differential algebraic (DA)
techniques are here proposed as a valuable tool to implement the higher-order numerical
and analytic extended Kalman filter. Working in the DA framework allows to consistently
reduce the required computational effort without losing accuracy. The good performance of
the proposed filter has been tested on different orbit determination problems with nonlinear
measurements and realistic orbit uncertainties. Numerical simulations show the good
performance of the filter in case of both complex dynamics and highly nonlinear
measurement problems.
16:20
THREE AXES ROTATIONAL MANEUVER CONTROL AND VIBRATION
SUPPRESSION OF A SMART FLEXIBLE SATELLITE
<strong>IAA</strong>-AAS-DyCoSS1-03-07
Mohammad Azadi, S. Ahmad Fazelzadeh,
Mohammad Eghtesad, and Emad Azadi
In this paper a satellite with a central hub and two flexible appendages which the
piezoelectric sensors and actuators are attached to them are considered. The satellite
moves in a circular orbit around the earth and maneuvers around three axes. The governing
equations of motion are derived based on Lagrange-Rayleigh-Ritz method and there is no
simplification or linearization in these equations. A Lyapunov based control is applied to the
system to not only control the three axes maneuver of the satellite but also suppress the
vibrations of the flexible appendages.
16:40
<strong>IAA</strong>-AAS-DyCoSS1-03-08
ANALYTIC SOLUTIONS OF QUASI-SATELLITE ORBITS IN THE
ELLIPTIC RESTRICTED THREE-BODY PROBLEM WITH
APPLICATION TO PHOBOS
Paulo J.S. Gil and Francisco S.P. Cabral
In the context of the restricted three-body problem, it is impossible to follow a keplerian
16

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
orbit around the second primary when he is very small because its region of influence and
Hill's sphere are too close or even under its surface. It is however still possible to orbit the
celestial body using quasi-satellite orbits (QSO), a type of distant satellite orbits. In this
paper, some approximate 3D analytic solutions of QSO are obtained for the case of the
elliptic 3BP with small eccentricity. Their features are discussed and an application to
Phobos is presented.
17:00
<strong>IAA</strong>-AAS-DyCoSS1-03-10
ELECTRIC SOLAR WIND SAIL CONTROL AND NAVIGATION
Petri Toivanen, Pekka Janhunen, Jouni Envall, and Sini Merikallio
The electric solar wind sail is a propulsion system that uses long centrifugally spanned and
electrically charged tethers to extract the solar wind momentum for spacecraft thrust. The
sail spatial orientation is controlled by modulating each tether voltage separately to produce
net torque for attitude control and thrust vectoring. Here, we cover the basics of the
electric sail control based on our dynamical models including single tether, rigid body, and
fully dynamical simulations of the entire sail. Solar wind data are used to address the effects
of the density and velocity variations on the sail dynamics, control, and navigation.
FP7 SPACE: OPPORTUNITIES FOR INTERNATIONAL COOPERATIVE
17:30-18:00
PROJECTS IN SPACE
(Ipanema I)
João Romana, NCP Space, Portuguese Foundation for Science and Technologies
The European Union 7th Frammework Programme will be allocating 126 M euro to the next
Space call (July 2012-November 2012). This call will financed Cooperative Projects in GMES
downstream services across all areas, Climate Change, Integration of SATCOM-SATNAV-EO,
Exploitation of Space Science data, Development of Space technologies and Critical
Technologies, Space SME Spin-offs and International cooperation with Ukraine and China.
Consortia of European and non-European entities may submit proposals to FP7 SPACE in all
topics. Participants from Non-European Space Faring Nations are particularly welcome,
namely, Canada, USA, Russia, Japan, China, India, South Africa, Brazil and Ukraine.
This session will allow a preliminary glance at topics which will be covered in the next call
and raise awareness on the international scope and call details.
<strong>Conference</strong> Dinner 19:00-21:00
Tuesday, March 20, 2012
Session 04:
Attitude Sensors and Actuators (Ipanema I)
Chairs: Franco BERNELLI and Michael OVCHINNIKOV
8:30-12:30
8:30
TILTED WHEEL FOR THREE-AXIS ATTITUDE CONTROL
OF A RIGID SATELLITE
<strong>IAA</strong>-AAS-DyCoSS1-04-01
Lawrence Oyedeji Inumoh, Nadjim Horri, and Alex Pechev
This paper proposes a new type of control actuator that generates torques in all three
principal axes of a rigid satellite using only a reaction wheel and a simple tilt mechanism.
17

18
FINAL PROGRAM
The mechanism will rotate the spin axis of the wheel about two additional axes thereby
generating high control torque about the axes orthogonal to the wheel spin axis. Torque will
also be generated about the wheel spin axis through the increase or decrease of the wheel
speed.Fundamental mathematical dynamic model and numerical simulations are used to
demonstrate the agile three-axis attitude control capability of this proposed actuator.
8:50
<strong>IAA</strong>-AAS-DyCoSS1-04-02
ANALYSIS OF AMBIGUITY RESOLUTION METHODS FOR ATTITUDE
DETERMINATION USING GPS CARRIER PHASE MEASUREMENTS
Leandro Baroni and Helio Kuga
If three or more GPS antennas are body-mounted properly and GPS measurements are
collected simultaneously, then body orientation can be calculated. Carrier phase double
differences are used as GPS measurements and the ambiguous integer number of cycles in
the measurements must be determined. For algorithm speedup constraints steaming from
the fixed antennas configuration geometry are imposed. Two algorithms were analyzed
accomplishing real-time attitude determination. The test baselines were separated 1 meter
apart perpendicularly. The results showed that the accuracy of both methods is in the order
of 0.1° to 0.2° or better, but CPU times are quite different.
9:10
ATTITUDE DYNAMICS OF A SMALL-SIZED SATELLITE
EQUIPPED WITH HYSTERESIS DAMPER
<strong>IAA</strong>-AAS-DyCoSS1-04-03
Michael Ovchinnikov
Appearance of small-sized satellites, involvement of spin-off companies and universities
renovated interest to passive attitude control systems (PACS) use. Disturbed attitude motion
has to be damped first. Hysteresis damper leads due to its simple design, low cost,
reliability, magnetic characteristics stability in time. Problems to be solved while a
hysteresis damper for PACS is used are considered. Hysteresis models admitting compromise
between reality and descriptiveness of results obtained are given. Examples how to apply the
technique for gravity-gradient and magnetic PACS development and design are presented.
Samples of nanosatellites with hysteresis damper where the proposed approach was used are
shown.
9:30
MICRO-SUN-SENSOR PERFORMANCE VALIDATION
IN GROUND-REPRODUCED ORBITAL CONDITIONS
<strong>IAA</strong>-AAS-DyCoSS1-04-05
Giancarlo Rufino and Michele Grassi
This paper deals with the digital micro sun sensor under development at the university of
Naples: a novel, two-axis digital sun sensor exploiting a CMOS photodetector. This paper is
focused on the tests to validate the adopted design and algorithms. They are carried out
exploiting a fully-functional hardware model of the sensor in a laboratory facility for
simulation of sun illumination, accounting for variable sun-line as resulting from orbit and
attitude dynamics. This paper will present novel, improved algorithms using adaptive image
processing depending on illumination conditions during operation, and tests to assess sensor
performance by means of hardware-in-the-loop tests.
9:50
<strong>IAA</strong>-AAS-DyCoSS1-04-06
A DYNAMIC FRICTION MODEL FOR REACTION WHEELS
Valdemir Carrara and Addolfo Graciano da Silva
This paper addresses the problem of the bearing friction in a reaction wheel and applies a
dynamic friction model in the current control loop. The dynamic friction model assumes that
there are elastic bristles in the contact surfaces that bends when slipping forces are applied.
The bristle behavior mimics the Stribeck effect without discontinuities. Some experiments
were carried out in order to collect the necessary data to estimate the model parameters.

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
The model was then inserted in a Model Compensation Control in current control mode.
Results shown that there was significant improvement in the control performance, mainly
during zero-speed crossing.
Coffee Break 10:10-10:30
10:30
<strong>IAA</strong>-AAS-DyCoSS1-04-07
ON GROUND CALIBRATION OF TETRAHEDRON GYRO PACKAGE
Helio Koiti Kuga, Rafael Henrique Siqueira, and Valdemir Carrara
This work presents the on ground calibration of a gyro package composed of four fiber optic
gyros in a tetrahedron configuration. In this work the on ground calibration is performed with
a 3-axis turn table covering the work range of the chosen gyros. Analysis covering both the
uncalibrated and calibrated accumulated attitude determination error are shown. An Allan
variance analysis is performed in order to detect the main sources of noise and to improve
the mathematical model. The calibration procedures developed in this work could be used in
tests of an in-house FOG-IMU that will fly in next Brazilian missions.
10:50
LOW COST PROXIMITY NAVIGATION SYSTEM
BASED ON LASER TELEMETER
<strong>IAA</strong>-AAS-DyCoSS1-04-08
Fabrizio Stesina, Sergio Chiesa, and Nicole Viola
The paper deals with the crucial issue of proximity navigation for both space exploration and
terrestrial vehicle’s applications. It describes both the applied methodology and the results
obtained for a low cost system based on a COTS (Component Off The Shelf) laser telemeter
mounted on a rotating support, in order to generate bi-dimensional maps for unknown indoor
and outdoor environments. This system has been designed and tested and its capability of
reproducing bi-dimensional maps starting from unknown initial conditions has been
successfully verified.
11:10
A MEMS-BASED MULTI-SENSOR SYSTEM
FOR SATELLITE HIGH-ACCURACY ATTITUDE DETERMINATION
<strong>IAA</strong>-AAS-DyCoSS1-04-09
Igor Merkuryev
This paper deals with satellite attitude determination based on optimal fusion of
microelectromechanical vibratory gyroscope measurements and star trackers by stochastic
filtering. One of the ways to improve performances of micromechanical vibratory gyroscopes
is to analyze their dynamics and errors in order to find efficient methods of digital signal
processing. New nonlinear dynamic properties of electrostatically actuated microstructures
observed under harmonic excitations are studied. Results of bench tests of the integrated
navigating system of the small satellites are presented.
11:30
MODEL-BASED PID CONTROLLER FOR CUBESAT REACTION
WHEELS USING COTS BRUSHLESS DC MOTORS
<strong>IAA</strong>-AAS-DyCoSS1-04-10 Teun Hoevenaars, Steven Engelen, and Jasper Bouwmeester
This paper presents the design, verification and validation of the PID controller for the
reaction wheel system of the Delfi-n3Xt CubeSat. The Delfi-n3Xt reaction wheel system is
based on COTS Brushless DC motors and controlled using Pulse Width Modulation. Because
control is fully implemented in the digital domain, discretization effects have a significant
impact on the performance of the implemented continuous model controller, particularly at
high rotation rates. Applying some simple but effective modification, the model-based PID
controller is adapted such that it operates over the full range of operating speeds.
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FINAL PROGRAM
Session 05:
Mission Design and Optimization I (Ipanema II)
Chairs: Kathleen HOWELL and Jeng-Shing CHERN Rock
8:30-12:30
8:30
<strong>IAA</strong>-AAS-DyCoSS1-05-01
REFERENCE ATTITUDE DERIVING ALGORITHM FOR MULTI STRIP
IMAGING IN HIGHLY AGILE SPACECRAFTS
Ritu Karidhal, Subbarao K., Kesavaraju V., and Pandiyan R.
The computation of time varying image plane motion and deriving the reference attitude for
driving the spacecraft platform is the premise of high resolution optical imaging . Here, two
types of imaging are considered. Type-I imaging is applicable to the camera system having
linear CCDs where the platform attitude should be controlled such that the projection of the
linear CCD on ground is maintained perpendicular to imaging direction. Type-II imaging is
applicable to camera system with an area array CCD where the earth rotation compensation
becomes a must for all stages to capture signal from same ground region.
8:50
<strong>IAA</strong>-AAS-DyCoSS1-05-03
A RESEARCH ON TETHER-DRAGGING DEORBITING MISSION
OF DEFUNCT GEOSTATIONARY SATELLITES
Hai-tao Liu, Le-ping Yang, Qing-bin Zhang, and Yan-wei Zhu
The RObotic Geostationary orbit Restorer (ROGER) plans to use space claw or net mechanism
to capture and remove the defunct satellites in GEO. Tether-dragging deorbiting is a crucial
part on the removal of defunct satellites after capture. This paper conducts an in-depth
research on tether-dragging deorbiting issues of defunct geostationary satellites. Firstly, a
four-phase tether-dragging deorbiting scheme, which takes risks of relative motion into
consideration, is proposed. Secondly, the dynamics model of the tether-dragging deorbiting
system is established based on Lagrange equation. Finally, the four phases of tether-dragging
deorbiting scheme are simulated and simulation results prove its validity.
9:10
<strong>IAA</strong>-AAS-DyCoSS1-05-05
VENUS TRANSFERS DESIGN COMBINING INTERPLANETARY
SUPERHIGHWAYS AND LOW-THRUST ARCS
Chiara Finocchietti, Pierpaolo Pergola, and Mariano Andrenucci
Various space transfers strategies exploiting Interplanetary Superhighways (IPS) are proposed
and analysed to accomplish a reference Venus mission starting from geocentric orbits
compliant with an International Space Station departing scenario or from periodic orbits
around libration points of the Sun-Earth system. Generally, such transfers are impractical by
themselves due to long transfer times they imply. As a solution to this problem, we propose
the coupling of IPS trajectories and propulsive arcs or manoeuvres. This approach allows
both reducing the transfer times and realizing intersection between dynamical structures not
matching naturally in the phase space, at least for the inner planets.
9:30
DESIGN OF LOW-THRUST TRANSFERS TO LIBRATION POINT
PERIODIC ORBITS EXPLOITING MANIFOLD DYNAMICS
<strong>IAA</strong>-AAS-DyCoSS1-05-06 Pierpaolo Pergola, Chiara Finocchietti, and Mariano Andrenucci
We propose and analyse some mission scenarios based on Electric Propulsion systems to
accomplish missions from a low Earth parking orbit toward periodic orbits close to the
collinear points of both the Sun-Earth and Earth-Moon systems. A transfer strategy that
combines low-thrust arcs and manifold dynamics is considered to minimize the fuel
consumption. During the propulsive phase, the thrust laws are determined by means of
optimization approaches. Further, for the design of the ballistic phase, the most relevant
non-gravitational influences are accounted. Moreover, we present a low-thrust alternative
solution to accomplish the same targets of the Euclid mission.
20

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
9:50
<strong>IAA</strong>-AAS-DyCoSS1-05-07
KINEMATIC ATTITUDE MANEUVERS WITH PATH CONSTRAINTS
FOR ASTROSAT – AN INDIAN ASTRONOMY SATELLITE
Pandiyan Ramalingam, Ramesh A. S., and Sharanappa Sindole
Astronomy satellites carry sensitive instruments onboard for celestial observations. These
instruments should not pass through or stare bright objects such as Sun while undergoing
maneuvers. Indian ASTROnomy SATellite (ASTROSAT) carries 5 instruments. The instruments
should avoid Sun while the spacecraft maneuver from one celestial object to another. In this
paper, a simple maneuver strategy developed based on kinematics to avoid bright Sun while
maneuvering the spacecraft has been described. A pre-conceived definition of the spacecraft
axes avoids Sun during observation. The maneuver profile is tested for various combinations
of the targets in order to ascertain that it works without fail.
Coffee Break 10:10-10:30
10:30
<strong>IAA</strong>-AAS-DyCoSS1-05-08
HALO ORBITS CHARACTERIZATION IN THE EARTH-MOON
SYSTEMFOR SCIENTIFIC APPLICATIONS
Pierpaolo Pergola and Elisa Maria Alessi
Recently, a new effort has been reserved to explore the lunar environment with missions
aimed at orbiting the Moon and collecting great amounts of data in view of a future human
installation. In this context, several authors proposed the exploitation of periodic orbits
around the equilibrium point L 1 in the Earth-Moon system with different objectives. The
purpose of our work is to provide effective criteria that can be applied to choose a nominal
orbit operable to these ends. We consider as key parameters lunar surface coverage, solar
eclipses and orbit maintenance costs.
10:50
<strong>IAA</strong>-AAS-DyCoSS1-05-09
DESIGN OF TRANSFER TRAJECTORIES BETWEEN
THREE-DIMENSIONAL RESONANT ORBITS IN
THE RESTRICTED THREE-BODY PROBLEM
Mar Vaquero and Kathleen Howell
The application of dynamical systems techniques to mission design has demonstrated that
employing invariant manifolds and resonant flybys enables previously unknown trajectory
options and potentially reduces the Delta-V requirements. An analysis of three-dimensional
resonant orbits, as well as the computation and visualization of the associated invariant
manifold structures is explored in this investigation. Three-dimensional maps are used to
explore the relationship between the manifold trajectories associated with multiple resonant
orbits. As a result, three-dimensional transfers between stable and unstable resonant orbits
are identified in the Earth-Moon system.
11:10
EARTH-MOON LIBRATION STATIONKEEPING:
THEORY, MODELING, AND OPERATIONS
<strong>IAA</strong>-AAS-DyCoSS1-05-10 David Folta, Thomas Pavlak, Kathleen Howell, and Mark Woodard
Collinear Earth-Moon libration points have emerged as locations with immediate application.
These libration point orbits are inherently unstable and must be controlled at a rapid
frequency which constrains operations and maneuver locations. Stationkeeping is challenging
due to short time scales of divergence, effects of large orbital eccentricity of the secondary
and third body perturbations. We contrast and compare promising strategies including
Optimal Continuation and Mode Analysis that achieved consistent and reasonable operational
stationkeeping costs for the ARTEMIS mission. Background on theory and models used to
achieve these demonstrated results are discussed along with their mathematical
development.
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FINAL PROGRAM
Session 06:
Satellite Constellations and Formation Flying I (Biombo)
Chair: Gerard GOMEZ
8:30-12:30
8:30
<strong>IAA</strong>-AAS-DyCoSS1-06-01
EFFECTS OF ECCENTRICITY OF THE PARENT BODY
ON MULTI-TETHERED SATELLITE FORMATIONS
Giulio Avanzini and Manrico Fedi
This paper discusses the relevance of eccentric reference orbits on the dynamics of a
tethered formation, when a massive cable model is included in the analysis of a multi–
tethered satellite formation. The formations examined in this study are Hub-And-Spoke (HAS)
and Closed-Hub-And-Spoke (CHAS) configurations for In-Plane and Earth-Facing spin planes. A
detailed study on the effect of eccentricity over the formation behaviour is performed. A
stability analysis will evaluate the effect of eccentricity on tether elongation, relative
position, orientation and shape of the formation. The presence of periodic solutions will be
also studied.
8:50
<strong>IAA</strong>-AAS-DyCoSS1-06-02
HIGH-ORDER PERTURBATION SOLUTION
TO THE RELATIVE MOTION PROBLEM
Pini Gurfil and Martin Lara
A solution to the relative motion problem for satellites is constructed based on perturbation
theory. This solution provides the mean orbital elements, which are commonly used for
detecting bounded relative orbits, as well as the osculating ones that allow for direct
propagation in realistic models. The new solution is carried out up to the second order of J 2,
thus accounting for non-linearities of the model to a high extent, and hence being useful for
checking long-term boundedness without limiting to the case of tightly-controlled
formations.
9:10
<strong>IAA</strong>-AAS-DyCoSS1-06-03
APPLICATIONS OF ANALYTICAL SOLUTION
OF RELATIVE MOTION OF THE FORMATION FLYING
Ashraf Owis, Hany Dwidar, and Hani Mohammed
In the current work an analytic solution of the relative motion of formation flying in an
elliptic orbit will be found. The dynamics of the problem will be linearized around an elliptic
orbit reference. The problem is solved using feedback optimal control approach via
generating function technique. Solutions for both soft and hard constraints will be
formulated. Numerical simulations will be done.
9:30
<strong>IAA</strong>-AAS-DyCoSS1-06-04
A COMPARATIVE STUDY OF DYNAMICS MODELS AND CONTROL
STRATEGIES FOR SATELLITE FORMATION FLYING
M. Navab and M. Barati
A simulation method with a modeling error index is introduced to compare eight widely used
dynamics models of satellite formation flying with a precise propagator. This evaluation is
carried out by analyzing the effects of main error sources such as eccentricity of the
reference orbit, formation size, inclination and gravitational perturbations in low Earth
orbits. In the second part, three control strategies are undertaken to be applied on the
dynamics models. Applicability and effectiveness of these control methods to bring and keep
the deputy and chief satellites to a prescribed proximity are investigated.
22

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
9:50
<strong>IAA</strong>-AAS-DyCoSS1-06-05
MODELING AND PROPERTIES OF SATELLITE FORMATIONS
WITH NONCONTACTING INTERNAL FORCES
Weiwei Tsai, Leping Yang, Yanwei Zhu, and Yuanwen Zhang
Satellite formations with noncontacting internal forces have attractive prospect in future
space missions due to their distinct advantages. The key to the successful application of
these novel formations lies in their dynamics models. This paper derives a new relative
dynamics model that incarnates the unique features of such formations through analytical
mechanics, breaking through the bottleneck of conventional methods. A profound
investigation on the dynamics model is carried out on the basis of physical significance
interpretation, space environment perturbation effect analysis and comparison with other
models. Research in this paper lays a theoretical foundation for future design of such
systems.
Coffee Break 10:10-10:30
10:30
AUTONOMOUS DECENTRALIZED COORDINATION CONTROL
FOR FRACTIONATED SPACECRAFT FORMATION
RECONFIGURATION BASED ON CYCLIC PURSUIT STRATEGY
<strong>IAA</strong>-AAS-DyCoSS1-06-06
Min Hu and Guoqiang Zeng
In this paper, the relative translational dynamics of fractionated spacecraft and the general
formation configuration description method are proposed, and the algebraic graph theory is
introduced. Secondly, three-dimensional finite-time cyclic pursuit algorithms are put forward
for both single-integrator kinematics and double-integrator dynamics, which guarantee the
convergence of tracking errors in finite time rather than in the asymptotic sense. Then,
collision free navigation is achieved using the Artificial Potential Filed method, which is very
popular for terrestrial robotics. Finally, fractionated spacecraft re-gather maneuvering and
system upgrade to accommodating new modules are simulated. The simulation results show
the effectiveness of the proposed controller.
10:50
FORMATION FLYING DYNAMICS ANALYSIS
BY MEANS OF A VIRTUAL MULTIBODY APPROACH
<strong>IAA</strong>-AAS-DyCoSS1-06-08
Leonard Felicetti and Giovanni Palmerini
The paper proposes to consider a spacecraft formation as an orbiting multibody which joints
- located at the spacecraft positions - are connected by virtual links. In such a way, it is
showed that the control actions needed to formationkeeping can be easily computed by
means of the efficient software already developed and validated for applications like space
manipulators. In fact, the reactions at the joints of the virtual multibody equal the control
action requested to acquire or maintain the corresponding configuration. A proof to validate
the approach, as well as the successful results from simulations, are presented and
discussed.
11:10
FUNDAMENTAL-SOLUTION GUIDANCE
FOR SATELLITE RELATIVE MOTION IN ELLIPTIC ORBITS
<strong>IAA</strong>-AAS-DyCoSS1-06-09
Ryan Sherrill, Andrew Sinclair, and Thomas Lovell
The motion of a deputy satellite relative to a chief satellite in elliptic orbit can be modeled
with linearized time-varying equations. The solutions to these equations are linear
combinations six fundamental solutions, with the combinatorial constants representing
integrals of the motion. These constants provide a geometric description of the relative
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motion, showing how it is composed from the six fundamental modes. This paper analyzes
maneuver planning for relative motion using the fundamental-solution description. Impulsive
burns by the deputy satellite in the radial, transverse, and cross-track directions are related
to the changes in amplitude of each fundamental solution.
11:30
SWITCHED STRATEGY AND SLIDING MODE
ALGORITHM-BASED RELATIVE POSITION CONTROL METHOD
OF ELECTROMAGNETIC FORMATION
<strong>IAA</strong>-AAS-DyCoSS1-06-10
Flying Jinxiu Zhang and Zhendong Hou
Considering the strong nonlinearity and coupling of electromagnetic formation, this paper
established the relative orbit dynamics with currents generating EM force as control
variables. An switching strategy and variable parameters’ sliding mode control algorithm was
proposed to deal with the deviation of orbital eccentricity, EM force model’s uncertainty and
control input matrix’s peak problem. Applying knowledge of system uncertainty, we obtained
reasonable control parameters of the sliding mode algorithm, the input matrix’s peak
problem due to free electromagnetic dipoles was solved by switching control strategy,
overcoming the sliding mode control method’s drawback of astonishingly large errors in some
local areas.
11:50
ULTRA-LOW EARTH ORBIT FORMATION FLYING PASSIVE
CONTROL USING DIFFERENTIAL AERODYNAMIC FORCES
<strong>IAA</strong>-AAS-DyCoSS1-06-11
Haiping Chen, Yulin Zhang, and Zhaokui Wang
In the ultra-low Earth orbit, which height is about 120km-180km, the spacecrafts are highly
affected by the aerodynamic forces. This paper analyses the aerodynamic forces of a panel
spacecraft, finding that, the order of magnitude of the lift force is large enough to control
the spacecraft formation flying. Based on the differential aerodynamic forces, a special
passive formation is designed. The non-sphere perturbation of Earth influences the formation
of the ultra-low Earth orbit spacecraft. Therefore, a method for including the J2 effects in
the passive formation design is introduced. The simulation shows that the method is
plausible and effective.
Lunch 12:30-14:00
Session 07:
Attitude Dynamics and Control II (Ipanema I)
Chairs: Antonio ELIPE and Anna GUERMAN
14:00-18:00
14:00
<strong>IAA</strong>-AAS-DyCoSS1-07-01
PARAMETER OPTIMIZATION FOR STABILIZERS
Ana Maria Seabra, Georgi Smirnov, and Anna Guerman
The choice of parameters for stabilizers can be based on various criteria; for example, for
linear controllable systems, the pole assignment theorem guarantees the existence of a
stabilizer with very high damping speed. However, such a stabilizer is practically useless
because of so-called pick-effect. The same phenomenon is observed if one chooses a robust
stabilizer constructed by H∞ or µ techniques. In this work we formulate an optimization
problem that allows us to determine optimal parameters of a stabilizer. We develop an
effective numerical tool oriented to optimization of stabilizer parameters according to
different criteria that appear in practice.
24

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
14:20
<strong>IAA</strong>-AAS-DyCoSS1-07-02
SMALL GAIN STABILITY THEORY
FOR MATCHED BASIS FUNCTION REPETITIVE CONTROL
Yunde Shi, Richard Longman, and Masaki Nagashima
Many spacecraft suffer from jitter produced by periodic vibration sources such as rotating
reaction wheels. Vibration isolation mounts are needed for fine pointing equipment. Active
control methods directly addressing frequencies of interest have the potential to completely
cancel the influence of these disturbances. Matched basis function repetitive control finds
error components at these frequencies, and can converge to zero error, using only frequency
response knowledge at addressed frequencies. A small gain stability theory is developed, and
robustness to model error is established. Controllers can be designed by pole-zero
placement, bypassing the complexity of periodic coefficient equations needed in previous
approaches.
14:40
ACTIVE MAGNETIC ATTITUDE CONTROL SYSTEM FOR SUN-
POINTING OF A SPIN-STABILIZED SATELLITE WITHOUT INITIAL
<strong>IAA</strong>-AAS-DyCoSS1-07-03
DETUMBLING
Dmitry Roldugin and Paride Testani
The angular motion of an axisymmetrical satellite equipped with the active magnetic
attitude control system is considered. Five different algorithms are implemented. Dynamics
of the satellite is analytically studied on the whole control loop using averaging technique.
Two coarse sun-pointing algorithms that do not need for the satellite to be detumbled and
nutation damping are studied. Fine sun-pointing algorithm is implemented last. Two
different algorithms are proposed. Active magnetic attitude control system time-response
with respect to different parameters is analyzed. The results show it is better to have highinclined
orbit. Terminal accuracy is obtained numerically.
15:00
A HYBRID ATTITUDE CONTROLLER COMPRISING OF
ELECTROMAGNETIC TORQUE RODS AND AN ACTIVE FLUID RING
<strong>IAA</strong>-AAS-DyCoSS1-07-04
Nona Abolfathi Nobari and Arun K. Misra
In this paper, a micro-satellite using a hybrid controller consisting of two magnetic coils in
the roll and pitch directions and a fluid ring in the yaw direction is considered. The control
torque, which stabilizes the satellite attitude angles, is divided into two components: one
component produced by the fluid ring is parallel to the Earth's magnetic field; the second
component is perpendicular to this direction, which is produced by the magnetic torque
rods. Then, the effect of the failure of the fluid ring and one magnetic coil on the satellite
attitude dynamics is examined.
15:20
FULL MAGNETIC SATELLITE ATTITUDE
CONTROL USING ASRE METHOD
<strong>IAA</strong>-AAS-DyCoSS1-07-05
Antonio L. Rodriguez-Vazquez,
Maria A. Martin-Prats, and Franco Bernelli-Zazzera
This work introduces a full magnetic control scheme for satellite attitude control. It is
particularized for a nadir pointing spacecraft. A new model which includes the satellite
attitude dynamics and kinematics joined with the magnetic field of the Earth is presented,
where the model is written in the State-Dependent-Coefficient (SDC) form. In addition, the
control problem is formulated by using the Approximating Sequence of Riccati Equation
algorithm and solved as a two point boundary value problem. Simulations results show a
satisfactory behavior under some test conditions.
25

FINAL PROGRAM
Coffee Break 15:40-16:00
16.00
ATTITUDE CONTROL FOR SMALL SATELLITES
USING ROTATION ANGLES
<strong>IAA</strong>-AAS-DyCoSS1-07-06
Teodor-Viorel Chelaru, Adrian-Mihail Stoica,
Cristian Barbu, and Adrian Chelaru
The paper present some aspects for calculus model of small satellites attitude control. The
satellite nonlinear model presented will be with six degrees of freedom. As novelty, the
rotation angles for describe the kinematical equations will be used. To highlight the
advantage of these parameters, kinematical equations will be described also by using Euler’s
angles and Hamilton's quaternion. Two attitude control cases will be analyzed: the reaction
wheels and micro jet engines. The results will be used in project European Space Moon Orbit
- ESMO founded by European Space Agency in which University POLIEHNICA of Bucharest are
involved.
16:20
ROBUST CONTROL OF THE ADVANCED
SOLID ROCKET LAUNCHER
<strong>IAA</strong>-AAS-DyCoSS1-07-07
Yasuhiro Morita
Japan’s advanced solid rocket launcher, Epsilon launch vehicle, is now under development by
JAXA and its first launch is officially declared to be conducted in 2013. The concept of the
next generation launch system requires simpler launch system and better user friendliness to
provide small satellites with an efficient launch. Such innovation makes the robust control
synthesis absolutely difficult. A preliminary design, which utilizes the H infinity control
algorithm, indicates that the obtained robust stability can be considered well within the
scope of expectation. The paper describes the current design of the attitude control system
of the Epsilon rocket launcher.
16:40
HYBRID METHODS FOR DETERMINING TIME-OPTIMAL,
CONSTRAINED SPACECRAFT REORIENTATION MANEUVERS
<strong>IAA</strong>-AAS-DyCoSS1-07-08
Robert Melton
Time-optimal spacecraft slewing maneuvers with path constraints are difficult to compute
even with direct methods. This paper examines the use of a hybrid, two-stage approach, in
which a particle swarm optimizer provides a feasible path (sometimes approximately
optimal), which serves as the input to a pseudospectral optimizer. Performance is compared
between a particle swarm optimizer and a differential evolution optimizer in the first stage.
17:00
USING THE STATE-DEPENDENT RICCATI EQUATION
AND KALMAN FILTER TECHNIQUES TO DESIGN
A SATELLITE ATTITUDE CONTROL SIMULATOR
<strong>IAA</strong>-AAS-DyCoSS1-07-09
Luiz Carlos Souza and Victor Arena
This paper presents the application of the State-Dependent Riccati Equation (SDRE) method
in conjunction with Kalman filter to design and test an attitude control algorithm for a 3D
satellite simulator. The control strategy is based on gas jets and reaction wheel torques to
perform large angle manoeuvre in three axes. Simulation has shown the performance and
robustness of the SDRE and Kalman filter controller applied for angular velocity reduction
associated with stringent pointing requirement. The investigation served to validate the
numerical model and to verify the functionality of the entire simulator system.
26

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
17:20
<strong>IAA</strong>-AAS-DyCoSS1-07-11
STABLE DESIGN OF ATTITUDE CONTROL FOR A SPACECRAFT
Mostafa Bagheri, Mansour Kabganian, and Reza Nadafi
Designed a three-axis attitude control Based on Lyapunov Stability Criteria for tracking
desired path, in the present paper. Attitude control system contains four reaction wheels
that their rotational axes are inclined to the x-y plane by an angle B. The simulation results
showed that the expected perfect tracking and controller is robust in the presence of
unknown external disturbances and alteration of parameters. Results indicate control robust
against twenty-two fold difference of parameter (moment of inertia) and robust against
disturbance 0.047 N.m that equal 39% rather the maximum torque of reaction wheel can
apply (nominal torque of reaction wheel).
Session 08:
Orbital Dynamics and Determination II (Ipanema II)
Chairs: Robert MELTON and Paolo TEOFILATTO
14:00-18:00
14:00
<strong>IAA</strong>-AAS-DyCoSS1-08-01
KEPLERIZATION OF MOTION IN ANY CENTRAL FORCE FIELD
Vladimir Martinusi and Pini Gurfil
The paper focuses on the closed-form solution of the absolute and the relative motion under
the influence of any central potential. It is proven that, by an adequate change of variable,
the motion in a central force field is reduced to a Keplerian motion. It helps to find explicitly
a Laplace-Runge-Lenz first integral, and also helps to provide a closed-form solution to the
equations of motion. The main application of this approach is to change the nominal Kepler
orbit by a central force field orbit, which incorporates parts of the perturbing potentials.
14.20
<strong>IAA</strong>-AAS-DyCoSS1-08-02
STABLE ARTIFICIAL EQUILIBRIUM POINTS
IN THE MARS-PHOBOS SYSTEM
Claudio Bombardelli
The existance of stable equilibrium regions in the CRTBP under constant thrust acceleration
has been pointed out recently. This article analyses the case of the Mars-Phobos system, in
which stable hovering can be obtained for a spacecraft coorbiting with the Moon at a
distance of about 80 km from its center with less than 15 micro-g of constant acceleration.
The system stability is investigated analytically and verified with a full numerical model
accounting for all relevant perturbations.
14:40
ORBITAL CHARACTERISTICS OF ARTIFICIAL SATELLITES
AROUND PLANETARY SATELLITES
<strong>IAA</strong>-AAS-DyCoSS1-08-03
Rodolpho Vilhena de Moraes, Jean P. S. Carvalho,
and Antonio Prado
Due to the scientific purposes of the missions, special orbits are desirable. In this paper,
conditions to get stability of frozen orbits, polar orbits, low inclination orbits and sunsynchronous
orbits around planetary satellites are analyzed taking into account the nonuniform
distribution of the mass of the central body and the influence of the perturbations
due to a massive body in elliptical orbit. An analytical theory using the averaged model and
the Lie-Hori method is presented. Applications were done by performing numerical
integrations of the analytical equations developed. Artificial satellites orbiting the Moon and
Europa are considered.
27

FINAL PROGRAM
15.00
<strong>IAA</strong>-AAS-DyCoSS1-08-04
PERIODIC SOLUTIONS IN PLANETARY ANNULUS PROBLEM
Eva Tresaco, Antonio Elipe, and Andrés Riaguas
Based on previous analyses where we analyzed the motion of an infinitesimal particle under
the attraction of an isolated planar annulus, we raise the question of the dynamic around a
planetary annulus, composed by a massive planar annulus and a central body. For this body,
we carry out a systematic search of the most relevant solutions: periodic orbits. We describe
the in-plane and out-of-plane motion of by means of the numerical continuation of families
of planar and 3-D periodic orbits; thus obtaining how the dynamic around a planetary ring is
organized.
15:20
ORBITAL LAUNCH WINDOW FOR MOON-TO-EARTH
TRAJECTORIES
<strong>IAA</strong>-AAS-DyCoSS1-08-05
Jingyang Li, Shengping Gong, Hexi Baoyin,
Junfeng Li, and Shing Yik Yim
An analytical design method of transEarth trajectory is developed with a finite sphere of
influence model by the patched-conic technique. Orbital launch window has been
established to study the mission sensitivities to transEarth trip time and energy requirement,
and also to provide the basis for the preparation of an orbital launch timetable compatible
with lunar missions and reentry conditions requirements. Results present here are limited to
a single impulsive maneuver. Difference between the results of the analytical model and high
fidelity model is compared. This difference is relatively small and can be easily eliminated
by a simple differential correction procedure.
Coffee Break 15:40-16:00
16:00
DETECTING INVARIANT MANIFOLDS USING
LAGRANGIAN COHERENT STRUCTURES
<strong>IAA</strong>-AAS-DyCoSS1-08-06
Daniel Pérez, Gerard Gómez, and Josep J. Masdemont
In this paper we compute hyperbolic invariant manifolds of a quasi-bicircular model of four
bodies by means of Lagrangian Coherent Structures (LCS). First, we use the Circular
Restricted Three Body Problem (CRTBP) to see how the LCS can be used to detect
stable/unstable manifolds. Then we apply the knowledge obtained to the quasi-bicircular
model.
16:20
UNIVERSAL FUCTIONS IN THE STUDY
OF THE RELATIVE ORBITAL DYNAMICS
<strong>IAA</strong>-AAS-DyCoSS1-08-07
Daniel Condurache and Vladimir Martinusi
The present work offers an approch to the relative orbital dynamics by using universal
functions. The solution to the relative orbital motion is offered in all posible situation, with
no restrictions imposed on the Keplerian reference and targeted trajectories. A unified view
on the relative orbital motion is sugested, by generalizing the previous approaches. The
exact solution is offered to the nonlinear model of the relative motion and it is expressed in
a coordonate free vectorial closed form.
16:40
<strong>IAA</strong>-AAS-DyCoSS1-08-08
A NOTE ON THE DYNAMICS AROUND THE L1,2 LAGRANGE
POINTS OF THE EARTH–MOON SYSTEM
IN A COMPLETE SOLAR SYSTEM MODEL
Yijun Lian, Gerard Gómez,
Josep J. Masdemont, and Guojian Tang
In this paper we study of the dynamics around the L1,2 libration points of the Earth-Moon
28

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
system in a full Solar System model. The study is based on the analysis of the quasi-periodic
solutions around both equilibrium points, whose computation is done using as initial seed the
Lissajous orbits of CR3BP. The analysis is done using a detailed Fourier analysis method for
the orbits. The results are compared with those obtained in [1] for the dynamics around the
L2 point in the quasi-bicircular model.
17:00
NUMERICAL METHOD FOR COMPUTING QUASI-PERIODIC
ORBITS AND THEIR STABILITY IN THE
RESTRICTED THREE-BODY PROBLEM
<strong>IAA</strong>-AAS-DyCoSS1-08-10
Zubin Olikara and Daniel Scheeres
Invariant manifolds in the restricted three-body problem are a powerful tool for the design
of spacecraft trajectories. This work presents an approach to compute families of quasiperiodic
orbits and their corresponding Floquet matrices. This matrix provides linear stability
information and can be used to generate stable and unstable manifolds for hyperbolic quasiperiodic
orbits. Including these orbits along with periodic orbits in the design space offers
additional low-energy transfer options. Both the circular and elliptic restricted three-body
problems are considered.
Session 09:
Spacecraft Guidance, Navigation, and Control II (Biombo)
Chair: Ijar da FONSECA
14:00-18:00
14.00
MODIFIED ACTIVE DISTURBANCE REJECTION CONTROL
FOR STATION-KEEPING OF ELECTROMAGNETIC
SATELLITE FORMATIONS
<strong>IAA</strong>-AAS-DyCoSS1-09-01
Weiwei Tsai, Leping Yang, and Yanwei Zhu
The high precision control problem for station-keeping of electromagnetic satellite
formations is investigated. Electromagnetic formation flight is a novel technology of stationkeeping
that does not require propellant. Control precision of station-keeping could be
improved since the electromagnetic forces are continuous and reversible. Considering the
internal and external disturbance in the system, a high precision control scheme is designed
based on active disturbance rejection control (ADRC) technology. To improve the capability
of ADRC, ideas of sliding model control and other modification are introduced in ADRC.
Simulation results are presented to validate the feasibility and efficiency of the modified
ADRC method.
14:20
SPACECRAFT ELECTROMAGNETIC DOCKING:
A REVIEW AND NEW RESULTS ON DYNAMICS AND CONTROL
<strong>IAA</strong>-AAS-DyCoSS1-09-02
Zhang Yuan-Wen, Yang Le-Ping,
Zhu Yan-Wei, Huang Huan, and Cai Wei-Wei
This paper presents a review on the field of spacecraft electromagnetic docking, including
the issues of dynamics and control. Four control characteristics are introduced and studied:
inter-satellite force trait, coupling, uncertainty and control capability insufficiency. Some
novel nonlinear control approaches are also derived. Firstly, Extended State Observer
method is applied to estimating the model uncertainties which are compensated by the
feedback of its equivalent estimate. Secondly, a dynamic model with twice electromagnetic
force in the right of Hill’s model and a practical tracking control approach are derived. At
the end, some useful conclusions and suggestions are put forward.
29

FINAL PROGRAM
14.40
<strong>IAA</strong>-AAS-DyCoSS1-09-03
A GROUND FACILITY TO TEST GNC ALGORITHMS AND SENSORS
FOR AUTONOMOUS RENDEZVOUS AND DOCKING
Giorgio Guglieri, Pasquale Pellegrino, and Liliana Torre
The paper presents the concepts and the design issues for a GNC implemented in a ground
facility used as a rendezvous and docking demonstrator together with the related simulator,
developed within the STEPS project (Systems and Technologies for Space Exploration), a
research project co-financed by Piedmont Region within EC program P.O.R - F.E.S.R. 2007-
2013. The GNC and the architecture of the test-bed are detailed, including a performance
analysis based on simulated and measured data.
15.00
HARDWARE-IN-THE-LOOP RENDEZVOUS SIMULATION
INVOLVING AN AUTONOMOUS GUIDANCE,
NAVIGATION AND CONTROL SYSTEM
<strong>IAA</strong>-AAS-DyCoSS1-09-04
Heike Benninghoff, Toralf Boge, and Tristan Tzschichholz
The rendezvous process is a key technology in multi-spacecraft missions like on-orbit
servicing missions. An active spacecraft (chaser) approaches a passive spacecraft (target) in
its orbit by performing controlled orbit and attitude maneuvers. The paper presents an
autonomous guidance, navigation and control system for rendezvous using a monocular
camera as vision-based sensor for relative navigation. Image processing algorithms and
navigation filters are employed to get accurate information about the relative position and
attitude between the two spacecrafts. The rendezvous sensor and the entire GNC system is
tested and verified at DLR’s robotic-based test bed European Proximity Operations Simulator
2.0.
15.20
DIFFERENTIAL DRAG SPACECRAFT RENDEZVOUS
USING ADAPTIVE LYAPUNOV CONTROL STRATEGY
<strong>IAA</strong>-AAS-DyCoSS1-09-05
David Perez and Riccardo Bevilacqua
The control forces required for rendezvous maneuvers at LEO orbits can be generated by
varying the aerodynamic drag affecting each spacecraft. This can be accomplished by
rotating a set of panels. Thus, the relative spacecraft motion can be controlled without using
any propellant or fuel since the motion of the panels can be powered by solar energy. An
Adaptive Lyapunov Controller is designed. A limit for the minimum drag acceleration that
ensures Lyapunov stability is found and used to adapt the Lyapunov controller to enhance its
performance. The method is validated using simulations in STK.
Coffee Break 15:40-16:00
16:00
STATE-DEPENDENT RICATTI EQUATION CONTROL
FOR AUTONOMOUS DOCKING OF SPACECRAFT
WITH COUPLED TRANSLATIONAL AND ROTATIONAL DYNAMICS
<strong>IAA</strong>-AAS-DyCoSS1-09-07
Daero Lee and Hyochoong Bang
Accurate dynamic modeling for the translational and rotational maneuvers between two
spacecraft is essential for the final phase of an autonomous docking. This paper describes
the state-dependent Ricatti equation control to meet the conditions for autonomous docking
of spacecraft with coupled translational and rotational dynamics. The control design is based
on a full, six-degrees-of-freedom, nonlinear coupled dynamic model. The spacecraft is
required to be able to perform translational and rotational maneuvers with sufficient
accuracy. Numerical simulations are used to demonstrate that the coupled translational and
30

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
rotational motion models can be effective in achieving the conditions for an autonomous
docking.
16:20
<strong>IAA</strong>-AAS-DyCoSS1-09-09
OBSERVER-BASED FINITE TIME CONTROL
FOR PROXIMITY TO TARGET SPACECRAFT
Shunan Wu, Zhaowei Sun, and Juntian Si
The finite time control for proximity to target spacecraft is investigated in this paper. Two
finite time controllers, based on fast terminal sliding mode technique, are respectively
proposed to perform proximity maneuver, which can drive relative position and velocity to
the equilibrium in finite time rather than in asymptotic sense. In particular, the escape
maneuver of target spacecraft is further considered, and a finite time observer is designed
and included in the modified controller to compensate the escape acceleration. The observer
can estimate the escape acceleration in finite time. Simulations are provided to illustrate
the performance of the proposed controllers.
16:40
<strong>IAA</strong>-AAS-DyCoSS1-09-10
IMAGE BASED CONTROL OF A FREE FLYING PLATFORM
Marco Sabatini and Giovanni Battista Palmerini
A free floating platform is realized with a pneumatic suspension system which enables a bidimensional
test of space operations. The platform is equipped with an IMU and actuated via
cold gas thrusters. A target will be acquired by an on-board camera and the image processed
for evaluating the control actions needed to reach it. A special effort will be devoted to the
reduction of the computational load for the image processing, because of the limited onboard
computational resources. Algorithms based on the prediction of the relative targetplatform
motion will be used to avoid false target identifications and consequent mission
failure.
Wednesday, March 21, 2012
Session 10:
Space Structures and Tethers (Ipanema I)
Chairs: Harijono DJOJODIHARDJO and Ivan KOSENKO
8:30-12:30
8:30
NUMERICAL SIMULATIONS OF AN ELECTRODYNAMIC TETHER
DEPLOYMENT FROM A SPOOL-TYPE REEL USING THRUSTERS
<strong>IAA</strong>-AAS-DyCoSS1-10-01
Kentaro Iki, Satomi Kawamoto, and Yoshiki Morino
The Japan Aerospace Exploration Agency (JAXA) has been investigating an active debris
removal system that employs highly efficient electrodynamic tether (EDT) technology for
orbital transfer. This study investigates the tether deployment from a spool-type reel using
thrusters by means of numerical simulations of an EDT system. The thrusters are used in
order to ensure the deployment of a tether with the length of several kilometers. In the
numerical simulations using a multiple mass tether model, the dynamics of tether
deployment is studied and requirements of thruster needed for the deployment is clarified.
8:50
STABLE AND UNSTABLE RELATIVE EQUILIBRIA
OF TWO TETHERED SATELLITES
<strong>IAA</strong>-AAS-DyCoSS1-10-02
Hassan Asiri
A tethered satellite system consisting of two point masses connected by a massless rigid rod
31

FINAL PROGRAM
is considered. Stability of relative equilibria of the system moving in a circular orbit around
the Earth in two dimensions is studied. Equations of motion of the system are derived and
two solutions for these equations are obtained. These solutions represent two different
orientations of the system in space: radial and tangential. The reduced energy-momentum
method is applied to analyze the stability of these orientations. The system is said to be
stable only at the radial position.
9:10
<strong>IAA</strong>-AAS-DyCoSS1-10-06
CSI DUE TO SLOSHING MOTION ON LEO LSS
Ijar da Fonseca and Peter Bainum
This paper discusses the mathematical modeling approaches to represent the sloshing
dynamics, the effect of sloshing motion on spacecraft system stabilization, the problem of
internal energy dissipation associated with rotating spacecrafts, and existing software tools
for dynamics analysis of the sloshing problem. The sloshing phenomenon is presented and
discussed in the scope of microgravity environment that characterizes the on orbit space
vehicles. The gravity-gradient is considered as the main disturbing effect on the LSS
attitude. The proportional integral derivative (PID) control approach is used to derive the
control law and the actuators are a set of reaction wheels and thrusters.
9:30
QUASI-PERIODIC MOTION OF A GROUND-BASED
TETHERED SUB-SATELLITE WITH ATTITUDE
<strong>IAA</strong>-AAS-DyCoSS1-10-07
Dongping Jin, Xiaoyu Wang, and Hao Wen
The quasi-periodic oscillation of a tethered satellite system with attitude is studied, based
on the dynamic similarity. The dimensionless ordinary differential equations that govern the
dynamics of the in-plane tethered sub-satellite system are obtained according to Kane’s
method. Numerical simulations show that there exists a quasi-periodic oscillation in the
system. To verify the quasi-periodic oscillation, an experimental test on the quasi-periodic
motion is made via a ground-based tethered satellite system, which is featured by using a
combination of air-bearing facilities and on-board thrusts to achieve the simulation of the
gravity gradient field forces as well as Coriolis force.
Coffee Break 10:10-10:30
10:30
EFFECTS OF UNCERTAINTIES AND FLEXIBLE DYNAMIC
CONTRIBUTIONS ON THE CONTROL OF A SPACECRAFT FULL-
<strong>IAA</strong>-AAS-DyCoSS1-10-08
COUPLED MODEL
Paolo Gasbarri, Riccardo Monti, and Marco Sabatini
The aim of this work is to present, by starting from the structural FEM model of flexible very
large space structure, a new spacecraft model more suitable for the design attitude control
laws. In this model the coupling effects between the gravitational forces and the elastic
modes will be considered in order to obtain a full-coupled model. In a second phase the
model will be used to evaluate the uncertainties effects on the inertial properties and on the
modal parameters on the robustness of common attitude control laws such as the PD, PID
and LQR applied to an orbiting spacecraft.
10:50
MODAL PARAMETERS IDENTIFICATION OF A TWO-DIMENSIONAL
SPACE STRUCTURE VIA VISUAL BASED TECHNIQUE
<strong>IAA</strong>-AAS-DyCoSS1-10-09
Riccardo Monti, Paolo Gasbarri,
Marco Sabatini, and Giovanni Palmerini
In this paper, an operational technique for the vibrating condition identification is proposed,
32

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
based on the use of visual based systems coupled with OMA (Operational Modal Analysis).
With respect to other proposed approaches managing flexibility effects, as nets of
accelerometers or even GNSS receivers, visual techniques are by far simpler to implement
and pose less constraints to the structure design. The use, even as an opportunity technique,
of optical sensors to monitor structural behavior in space systems has been the topic of
previous research works.
11:10
GENERAL CONCEPTUAL DESIGN PROBLEMS
OF A PARABOLIC SPACE STRUCTURE
<strong>IAA</strong>-AAS-DyCoSS1-10-10
Pedro Gamboa, José Silva, Anna Guerman,
Filipe Couceiro, and Benjamin Schneuwly
This paper addresses the conceptual design of a parabolic space structure with one major
requirement: the maintenance of a perfect parabolic shape required to maximize the
amount of sunlight reflected to its focal point. The work is divided into two main parts
aiming at assessing distinct approaches to the parabolic configuration. The first one
determines the deformed shape of a flat circular membrane subject to an initially uniformly
distributed load representative of the solar pressure and the second part relies upon the
analysis of the deformed field of a boom/membrane umbrella-like structure where several
design parameters are taken into consideration.
11:30
ANALYSIS OF DESIGN PARAMETER INFLUENCE ON THE DYNAMIC
FREQUENCY RESPONSE OF CFFF HONEYCOMB SANDWICH PLATE
<strong>IAA</strong>-AAS-DyCoSS1-10-11
Mankour Abdeldjelil, Boudjemai Abdelmadjid,
Salem Houria, and Amri Reda
In the aerospace domain, sandwich plates represent an efficient structural element,
providing a high stiffness/weight ratio characteristic. Moreover, when using this structural
element, different design parameters and materials of the core can be adopted in order to
obtain desired properties. Geometric parameters of hexagonal plate is specific to absorb
vibrations, hence the effect of each parameters is crucial to determine the rigidity of plates
under a single-point cyclic loading.
Session 11: Mission Design and Optimization II (Ipanema II) 8:30-12:30
Chair: Antonio PRADO
8:30
MINIMUM FUEL MULTI-IMPULSIVE ORBITAL MANEUVERS
USING GENETIC ALGORITHMS
<strong>IAA</strong>-AAS-DyCoSS1-11-01
Denilson Santos and Antonio Prado
This research has the goal to analyze optimal multi-impulsive rendezvous maneuvers
between two spacecrafts that are in coplanar orbits. The main objective is to optimize these
transfers with respect to the fuel consumption, not considering the time required for the
transfer and using multiple burns during the process. A genetic algorithm is used in order to
solve the optimization problem. The problem will be solved using a different number of
impulses, in order to verify which one has the lowest consumption. Several missions can
benefit from the optimization techniques showed in this work.
8:50
MISSION ANALYSIS AND SYSTEMS DESIGN
OF A NEAR-TERM AND FAR-TERM POLE-SITTER MISSION
<strong>IAA</strong>-AAS-DyCoSS1-11-02
Jeannette Heiligers, Matteo Ceriotti, Colin McInnes,
and James Biggs
The pole-sitter concept involves a spacecraft that is continuously above the North or South
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FINAL PROGRAM
Pole. As such, it can provide continuous and hemispherical coverage of the polar regions. To
provide the required thrust, two propulsion strategies are proposed, resulting in a near-term
and far-term pole-sitter mission using solar electric propulsion (SEP) and hybridised SEP and
solar sailing, respectively. For both cases the mission analysis and systems design will be
provided, including minimum propellant pole-sitter orbits, optimal Earth to pole-sitter
transfers, a detailed mass budget analysis to trade-off mission lifetime and payload mass and
a feedback control system for orbital control.
9:10
<strong>IAA</strong>-AAS-DyCoSS1-11-03
STATION KEEPING OF A SOLAR SAIL AROUND A HALO ORBITS
Ariadna Farrés and Angel Jorba
In this work we will discuss the controllability of a solar sail around a Halo Orbit in the Earth
- Sun system. We use dynamical system tools to understand the phase space behaviour for
changes on the sail orientation. We will then use this information to derive a station keeping
strategy. We will also test their robustness when different sources of errors in position and
attitude and included in the simulations.
9:30
SIMPLE METHOD FOR PERFORMANCE EVALUATION
OF MULTISTAGE ROCKETS
<strong>IAA</strong>-AAS-DyCoSS1-11-05
Paolo Teofilatto and Mauro Pontani
Performance evaluation of multistage rockets is aimed at defining the maximum payload
mass that can be inserted in the desired orbit. This work proposes a simple method for a fast
performance evaluation of multistage rockets. The technique at hand is based on three
steps: (i) the flight path angle at each stage separation is guessed through analytical
formulas, (ii) the spacecraft velocity is maximized at each stage separation, and (iii) coasting
arcs are determined in a semi-analytical way. Comparison of the results with the nominal
trajectory of the VEGA launcher demonstrates the suitability of the present approach.
Coffee Break 10:10-10:30
10:30
ASSESSMENT OF THE PROBABILITY OF COLLISION AMONG SPACE
OBJECTS, 2009 IRIDIUM-COSMOS COLLISION CASE STUDY
<strong>IAA</strong>-AAS-DyCoSS1-11-06
Navabi and R. Hamrah
This paper deals with the assessment of the collision probability related to some predicted
close approaches between two orbiting space objects. This follows that if a critical distance
is going to be happened and the probability of its occurrence is exceeding a safe threshold,
the time available for collision avoidance maneuver will be obtained. Simplifying
assumptions such as a linear relative motion and normally distributed position uncertainties
at predicted closest approach time are applied to estimate the probability of collision using a
formulation that takes into account the object sizes, covariance and the relative distance at
point of closest approach.
10:50
DEFINITION AND GENERATION
OF MULTI-SEGMENT LUNAR FREE-RETURN TRAJECTORIES
<strong>IAA</strong>-AAS-DyCoSS1-11-07 Jingyang Li, Shengping Gong, Hexi Baoyin, Junfeng Li, and Xue Ma
A multi-segment lunar free return trajectory for human missions is proposed. Differing from
hybrid returns this transfer trajectory consist of free return sections only, while retaining the
advantage of hybrid returns. The translunar injection is made onto a free-return trajectory
with a great perilune altitude, and then the spacecraft performs a transfer maneuver onto a
new free-return translunar trajectory from which the lunar orbit insertion can be performed.
34

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
An analytical design method of the multi-segment free return trajectories is developed.
Difference between analytical and high fidelity models is relatively small and can be easily
eliminated by a differential correction process.
11:10
APPLICATIONS OF GRAVITY ASSIST FLIGHT IN SEVERAL
WELL KNOWN MISSIONS IN THE PAST
THREE DECADES AND ITS PROSPECT
<strong>IAA</strong>-AAS-DyCoSS1-11-08
Simei Ji
A decade ago, G.A. Flandro published a paper tracing the true origins of gravity assist and its
application in spaceflight. Ten years past, new papers appeared, graphical design methods
become popular, example calculations being carried out and software developed for
automatic design, more importantly, dozens of missions have been successful operated based
on it. This paper focuses on the currently widespread methods of multi-gravity assist design,
emphasized on the application of the theory in several well-done missions, analyzed and
compared the utility of gravity assist transfers in these missions, summarized several
important aspects for the further application in future spaceflight.
11:30
MISSION ANALYSIS FOR OPERATIONAL ORBIT
AND TRANSFER STRATEGY OF ESA PROBA-3 MISSION
<strong>IAA</strong>-AAS-DyCoSS1-11-09
Gabriele Bellei and Juan L. Cano
Proba-3 is the third in ESA’s series of missions for validating developments in space systems
and will demonstrate the technologies required for formation flying of two spacecraft. This
work presents all the mission analysis activities carried out for the design of the operational
orbit and the ascent trajectory from launch to the operational orbit. The operational orbit is
designed to be a HEO with period of 19.8 hours and synchronicity of 14/17 which allows
operating nominally for 2 years and an extension of another 2 and naturally re-enters Earth
before 25 years from launch.
11:50
OPTIMIZATION OF TARGET FUNCTIONING PLANS AND
CONSTELLATIONS OF SATELLITE OBSERVATION
AND COMMUNICATION SYSTEMS
<strong>IAA</strong>-AAS-DyCoSS1-11-10
Valeriy Darnopykh and Veniamin Malyshev
An important control aspect of modern satellite observation and communication systems is
the control of functional processes. Systems constellations can include any number of
spacecrafts (SC) in different planes. Constellations optimization is necessary for increase
efficiency of multi-satellite systems target functioning. This is polynomial complexity
fundamental problem. The paper's purpose is presentation of the methodical approach for
problem solution. It realizes several stages based on the repeated solution of operative plan
optimization. Several examples are considered and illustrated for constellations: of
observation SC with parameters of "RapidEye" and "CBERS" systems, of “Iridium” (66 SC) and
"Teledesic" (288 SC) global communication systems.
Session 12:
Satellite Constellations and Formation Flying II (Biombo)
Chairs: David FOLTA and Georgi SMIRNOV
8:30-12:30
8:30
ANALYSIS OF SATELLITE CONSTELLATIONS ON PSEUDO-SUN-
SYNCHRONOUS, HIGHLY ELLIPTICAL ORBITS
<strong>IAA</strong>-AAS-DyCoSS1-12-01
Yuri Ulybyshev
For almost continuous coverage of lighting Earth’s areas can be used new pseudo-sun-
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synchronous highly elliptic orbits using the critical inclination and orbit apogee in the Earth’s
hemisphere with the coverage areas.. In opposite to the classical Molniya-type orbits, the
new orbits are synchronized with the solar time (or the solar day) that the apogee time is
coincident with the local noon of a coverage area. Qualitative analysis and simulation results
for the constellations with 2-4 satellites at the orbits are presented. For such coverage, the
constellations required less satellites than highly elliptical, Molniya-type constellations for
continuous zonal or regional coverage.
8:50
<strong>IAA</strong>-AAS-DyCoSS1-12-02
CONSTELLATION DESIGN FOR RESPONSIVE REVISITING
Ming Xu and Colin McInnes
Responsive communications or reconnaissance missions typically require frequent revisit to
any point of the Earth’s surface and employ a constellation with two satellites to improve
coverage performance. Different from traditional constellations, an interesting constellation
holding the interval between the successive revisits on all latitudes is presented in this
paper. Numerical simulation is implemented to illustrate that the revisiting constellation
exists in a large region parameterized by the relative values of right ascension of ascending
node and argument of latitude. Finally the analytic solution to the existence of the
significant constellation is provided.
9:10
APPLICATION OF SDRE TECHNIQUE TO ORBITAL AND ATTITUDE
CONTROL OF SPACECRAFT FORMATION FLYING
<strong>IAA</strong>-AAS-DyCoSS1-12-03
Mauro Massari and Mattia Zamaro
In this paper a control system for orbital and attitude dynamics of Formation Flying is
presented. The control exploits the coupling imposed by the set of thrusters used for the
control of both the relative dynamics. The design is based on the State-Dependent-Riccati-
Equation Technique which has been recently applied to nonlinear system in many field of
engineering. The SDRE algorithm is implemented with a timing strategy for both the
controller and the nonlinear Kalman filter to reduce the computational burden. Numerical
results considering the ESA Proba-3 mission as test-bed will be presented for both orbital and
attitude relative manoeuvre.
9:30
ON HILL’S CLOHESSEY WILTSHIRE EQUATION
FOR ANALYSIS OF RELATIVE MOTION
OF FORMATION FLYING GEO-STATIONARY SATELLITES
<strong>IAA</strong>-AAS-DyCoSS1-12-04
Vinod Kumar, Hari B. Hablani, and Ramalingam Pandiyan
This paper uses HCW equations to study the natural relative motion of formation flying
stationary satellites. To develop the complete on-board autonomous navigation and control
of the satellites in formation in the presence of disturbance. Keeping this goal in mind we
have used HCW equations to understand the formation dynamics of the satellites.We will
compare the various types of sensors used for the formationkeeping and their accuracy
specifications. While HCW equations are useful to understand the dynamics of the system but
these equations introduce the uncertainty that is more than the imaging sensor thus making
them un-suitable.
9:50
DESIGN OF AN EARTH-OBSERVATION SATELLITE
CONSTELLATION VIA IMAGE RETURN TIME FIGURE OF MERIT
<strong>IAA</strong>-AAS-DyCoSS1-12-05
Burak Akbulut, Ahmet Özkan, Erhan Topal,
Görkem Oktay, Kağan Ataalp, and Erhan Solakoğlu
This paper aims to analyze an optimal configuration for a satellite constellation needed for
successful completion of an imaging request from a predefined target area. A figure of merit

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
is defined for comparison of different constellation configurations. Different numbers of
satellites are analyzed in single and multi orbital planes. Preliminary results show that it is
not beneficial to increase number of satellites in a plane after a certain limit and multi plane
approach does not provide significant improvement to results when a target area within a
communication cone of a single ground station is considered.
Coffee-break 10:10-10:30
10:30
<strong>IAA</strong>-AAS-DyCoSS1-12-06
CONSTELLATION ORBITS SELECTION CONSIDERING
THE INTEROPERABILITY BETWEEN DIFFERENT GNSS
Fan Li, Yang Xuerong, Jiang Chao, and Wang Zhaokui
The paper is mainly focused on the orbits selection for the COMPASS global constellation.
Two optimized schemes were proposed, for the interoperability with GLONASS and Galileo.
Because of the high drifting speed of GPS’s orbit planes, it is very difficult to choose a
constellation configuration suits for GPS. The altitude 24126km was chosen for the COMPASS
global constellation, because of the high performance, less perturbation influences, and so
on. The Walker 24/3/2: 51.59°, 24126km constellation’s performances interoperated with
the GPS, GLONASS and Galileo were studied. Great availability improvement was achieved in
35° elevation through the drifting rate design.
10:50
PRELIMINARY RESULTS OF THE VISION BASED RENDEZVOUS
AND FORMATION FLYING EXPERIMENTS
PERFORMED DURING THE FFIORD/PRISMA EXTENDED MISSION
<strong>IAA</strong>-AAS-DyCoSS1-12-07
Michel Delpech, Jean-Claude Berges,
Sophie Djalal, and Pierre-Yves Guidotti
The Formation Flying In Orbit Ranging Demonstration experiment (FFIORD) has been
accommodated by CNES on PRISMA precursor mission. The paper will recall the most
significant results obtained in 2010/2011 during the FFRF sensor characterization phase then
the rendezvous / FF experiments based on the RF navigation which include in particular
proximity control accuracy at the centimetre level. Additional experiments were run in
October 2011 during the extended mission that involve this time visual sensors (VBS) in the
loop to be used in combination with the FFRF sensor. An experiment overview will be given
along with a result presentation.
11:10
INTEGRATED RELATIVE TRANSLATION AND ROTATION
CONTROLOF SPACECRAFT FORMATION
<strong>IAA</strong>-AAS-DyCoSS1-12-08
Iva Mohan Sundara, Pandiyan Ramalingam,
Debasish Ghose, and Seetharama Bhat
Distribution of small spacecrafts in formation provides a way to improve the resolution by
aperture distribution. This requires autonomous control of relative position and relative
attitude. Present work addresses the formation control using a PID controller to maintain
both relative position and relative attitude. To avoid continuous pulsing due to noise, a
dead-band has been provided in the position loop. PID control has been selected to maintain
the formation in presence of unmodeled disturbances. Simulations show that proposed
controller meets the translational and rotational relative motions even in presence of
disturbances.
11:30
<strong>IAA</strong>-AAS-DyCoSS1-12-09
PERFORMANCE MEASURE BASED TRAJECTORY
OPTIMIZATIONFOR LEO TWIN-SPACECRAFTS
FOR NEAR-EQUATORIAL GROUND-TRACK REQUIREMENTS
Harijono Djojodihardjo
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FINAL PROGRAM
Parametric and optimization study is carried out to explore the potential of spacecraft
formation flying as platform for Near-Earth and Near-Equatorial Tropical Resources And
Environment Monitoring And Disaster MitigationMission. The optimization is based on
minimization of the corresponding delta-V budget which represents the energy cost needed
for each maneuver scenarios. The ground track requirements are elaborated in view of
relevant sensors and instruments. Coverage Figure of Merit and Performance measure for
assessing the benefits of relevant ground-track parameters are redefined. Another
optimization scheme based on some defined ground track requirements and mission
objectives is carried out. Both results are discussed.
11:50
<strong>IAA</strong>-AAS-DyCoSS1-12-10
PRECISION CONTROL SYSTEM FOR INNER-FORMATION
GRAVITY FIELD MEASUREMENT SATELLITE SYSTEM
Junhu Xiang
Inner-formation gravity field measurement satellite system consists of an inner satellite and
an outer satellite, and the inner satellite is shielded from the air drag, solar light pressure
and other non-gravitation forces by the cavity of outer satellite. The precision control
system has a hybrid control architecture including feed forward and feedback control, and
base on the model predictive control method and all-propulsion system of the cold gas
thruster, which has realized the highly precision integrated control of attitude and orbit, and
satisfied the performance requirements of high precision gravity field measurement mission
with lower costs and technology requirements.
12:10
STUDY OF ON-ORBIT SEPARATION SCHEMES FOR
CONFIGURATION INITIALIZATION OF FRACTIONATED
SPACECRAFT CLUSTER FLYING
<strong>IAA</strong>-AAS-DyCoSS1-12-11
Jiang Chao, Yang Xuerong, Wang Zhaokui, and Zhang Yulin
Using one Multi-satellite Deployment System to provide each module once velocity
increment, the fractionated spacecraft cluster configuration can be formed without module
maneuver. At first, the dynamics of cluster configuration initialization is analysed, based on
which the constraints of separation initialization schemes design is given. Then, according to
various of separation parameters, series of separation initialization schemes are designed. At
last, the separation error affect and the cluster flying natural evolution are simulated in the
high precision formation flying simulation environment, and the best separation initialization
scheme is selected by comparing of the simulation results.
Lunch 12:30-14:00
Session 13:
Attitude Dynamics and Control III (Ipanema1)
Chairs: Arun K. MISRA and Peter BAINUM
14:00-18:00
14:00
<strong>IAA</strong>-AAS-DyCoSS1-13-02
ATTITUDE CONTROL SOFTWARE SOLUTION
FOR A BIG SOLAR PANEL ON LEO ORBIT
Stefano Rossi
The future of the alternative energy is in the space: the solar power is one of the best
candidates to help human demand of energy. In space, the capture of the solar energy will
have benefits compared to the earth capture. In this paper, a software solution is presented
for a big solar panel attitude control. The panel, placed on LEO orbit, has been designed
38

1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
with reaction wheels putted in the center of the structure, simulating maneuvers, impacts,
sun tracking, and solar pressure destabilization. Some sizing studies as panel mass, wheel
masses and inertia are presented using this software.
14:20
ATTITUDE AND ORBIT CONTROL SYSTEM (AOCS)
DESIGN FOR ASTROSAT
<strong>IAA</strong>-AAS-DyCoSS1-13-03
Jasvinder Singh Khoral, Natrajan P, Pandiyan Ramalingam,
Krishna Mohan Bharadwaj, and Parmeswaran K
Astrosat is a satellite meant for Astronomical observations. Astrosat has five payloads for
multi-wavelength astronomy observations. The payload configuration constrains the sun to
lie in a very narrow region of the roll-yaw plane.In this paper we show how the constraints
are exploited and a simple rest-rest re-orientation maneuver that avoids bright objects is
implemented via the Euler rotation theorem. As far as is known, the use of the Euler
theorem to avoid bright objects has not been reported in the literature and hence this is a
novel application of the same!
14:40
THREE AXES REACTION THRUSTERS ATTITUDE CONTROL
SYSTEM DURING TRAJECTORY MANOEUVRES FOR ESMO
<strong>IAA</strong>-AAS-DyCoSS1-13-04
Claudiu-Lucian Prioroc and Raluca Mihaela Stefanescu
The attitude inertial control system during Lunar orbit insertion for ESMO has been designed.
Four attitude control thrusters and MEMS rate sensors have been chosen as control hardware
of the satellite. The system has been designed to minimize the delta-V error. Modelling of
the spacecraft attitude control hardware, sloshing, main engines, attitude dynamics and
kinematics is presented with simulation results validating the attitude control system.
15:00
DYNAMICAL ATTITUDE MODEL FOR GAIA,
SENSORS ON-BOARD AND PERFORMANCES
<strong>IAA</strong>-AAS-DyCoSS1-13-05
Daniel Risquez, Floor van Leeuwen, and Anthony Brown
The Dynamical Attitude Model (DAM) is a simulation developed to achieve a detailed
understanding of the Gaia spacecraft attitude. This simulation takes into account some
disturbances and considers as well internal hardware components controlling the satellite
like the AOCS (Attitude and Orbit Control System), sensors (a star tracker and an angular rate
sensor attached to the main instrument), and micro-Newton thrusters. This contribution
provides a brief description of DAM and an analysis of some effects on the Gaia's attitude, as
characteristic profiles of micro-meteoroid impacts and clanks (discontinuities in the attitude
while the angular rate is constant).
15:20
AERODYNAMIC ANGLES-ORIENTED ATTITUDE CONTROL OF A
LOW LIFT-TO-DRAG BICONIC ATMOSPHERIC ENTRY CAPSULE
BASED ON THE EMBEDDED MODEL CONTROL METHODOLOGY
<strong>IAA</strong>-AAS-DyCoSS1-13-07 Jose Alejandro Ospina, Enrico Canuto, and Marcello Buonocore
This paper describes the attitude control system of a low lift-to-drag biconic atmospheric
entry capsule. The control structure derives from the development of the attitude dynamics
and kinematics written in terms of the aerodynamic angles, instead of the Euler or
quaternion kinematics. Non linear dynamic inversion and active disturbance rejection
techniques are used to handle gyroscopic torques, parametric errors and to compensate for
the angular variation of the translational velocity. The performance of the control law is
tested on a high-fidelity simulator and the results are presented.
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FINAL PROGRAM
Coffee-break 15:40-16:00
16:00
A “TURN-BY-TURN” COMBINED WHEEL-TORQUER ACTUATED
CONTROL SYSTEM FOR SPACE-CAPSULE
RECOVERY EXPERIMENT (SRE)
<strong>IAA</strong>-AAS-DyCoSS1-13-08
Harish Joglekar, Nivriti Priyadarshini, P Natarajan,
K Parameswaran, and P J Bhat
A combined wheel-magnetic torquer attitude control is proposed for Space Capsule Recovery
Project. The objective is two axis attitude control using torquers and one axis control using
wheel and wheel momentum unloading using magnetic torquers. The Y axis is continuously
controlled using wheel and other two axis and wheel momentum are dumped in "turn by
turn" fashion. Control cycle is divided into three parts and torquer choice is based on the
maximum magnetic field criterion. At a time a single axis is either controlled/dumped using
one torquer only. The advantage is simultaneous control and dumping and useful for coarse
pointing.
16:20
ATTITUDE CONTROL SYSTEM DESIGN
FOR LOW-ORBIT MICRO-SATELLITE
<strong>IAA</strong>-AAS-DyCoSS1-13-09
Peng Wang, Wei Zheng, and Jie Wu
Micro-satellite, which is lightweight, small, short development cycle and low cost, has
become one of the hottest fields of aerospace research. Aerodynamic torque is the primary
environmental torque for low-orbit satellite. In this paper, a new scheme of attitude control
with active magnetic torque and aerodynamic stabilization torque is proposed. The control
law B_dot and sliding mode control method are used in detumbling and attitude acquisition,
respectively. The index approach law and saturation function are used in designing the
sliding mode attitude control system. Simulation studies demonstrate that the proposed
attitude control system is feasible for low-orbit micro-satellite.
16:40
DESIGN OF THE ACTIVE ATTITUDE DETERMINATION
AND CONTROL SYSTEM FOR E-ST@R CUBESAT
<strong>IAA</strong>-AAS-DyCoSS1-13-10
Fabrizio Stesina and Sabrina Corpino
One of the key points to increase the number of applications for pico/nano-satellite is to
enhance their attitude control system capabilities. The paper presents the results of the
design, development, and verification campaign of the A-ADCS (Active - Attitude
Determination and Control System) of the e-st@r cubesat developed at Politecnico di Torino.
In particular, the results obtained with numerical and Hardware-In-The-Loop (HITL)
simulations are discussed and compared. The analysis shows that the results of the HITL
simulations are consistent with those coming from numerical simulations. The complete
validation of the simulator will be possible when orbital data are available.
17:00
ATTITUDE DETERMINATION AND CONTROL OF PRATHAM,
INDIAN INSTITUTE OF TECHNOLOGY BOMBAY’S
FIRST STUDENT SATELLITE
<strong>IAA</strong>-AAS-DyCoSS1-13-11 Sanyam Suhas Mulay, Jaideep Joshi, Yashovardhan Sushil Chati,
Vaibhav Vasant Unhelkar, Saptarshi Bandyopadhyay,
Shashank Tamaskar, Mallesh Bommanahal,
Chaitanya Talnikar, and Avnish Kumar
This paper describes the attitude determination and control of 'Pratham', IIT Bombay's first
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1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
student satellite. The satellite being built is a 26x26x26 cm cube and weighs nearly 10 kg.
The attitude control requirements on the satellite are pointing accuracy of 10 deg. The
sensors on board the satellite are GPS, magnetometer and sunsensors. Magnetorquers are
used as actuators. All simulations are carried out using Matlab and Simulink. The paper
discusses the control strategies used by Pratham during the detumbling and nominal modes.
Robustness is checked using Monte Carlo simulations. Results are successfully validated with
real-time in-loop simulations.
Session 14:
Optimal Control in Space Flight Dynamics (Ipanema II)
Chairs: Bernd DACHWALD and Mª do Rosário de PINHO
14:00-18:00
14:00
<strong>IAA</strong>-AAS-DyCoSS1-14-01
A NEW NUMERICAL APPROACH FOR SOLVING OPTIMAL INNER
CONTROL OF PARABOLIC PDE PROBLEM
Mahmoud Mahmoudi
In this paper, we develop a computational method for solving an optimal control problem,
which is governed by a parabolic partial differential equations (PDEs) . Our approach is to
transform the PDE problem to high dimension non-homogeneous ordinary differential
equation system (ODES). In follow, the homogeneous part of ODES is solved using semigroup
theory. In the next step, the convergence of this approach could be verified by properties of
one-parameter semigroup theory. In the rest of paper, the optimal control problem is solved
by utilizing the solution of homogeneous part. Finally one numerical example is given.
14:20
DYNAMIC CONTROL AND VIBRATION SUPPRESSION
OF THE RIGID-FLEXIBLE COUPLING SPACE
ROBOT DURING CAPTURE TARGET
<strong>IAA</strong>-AAS-DyCoSS1-14-02
Qiu-Huang Dong and Li Chen
The main issue of this paper is about the space robot capture operation. The stabilization
control and vibration suppression control of the rigid-flexible coupling space manipulator
during capture target is presented in this paper. The dynamic model of the space robot and
target combined system is established base on the mechanics theory. The feedback control is
employed for stabilization of the combined, and the Linear Quadrics optimal control is
employed suppress the vibration of the flexible link, the above control methods are
employed at the same time. The computer simulation result verify the feasibility of the
above control method.
14:40
A METHOD TO SOLVE NONLINEAR OPTIMAL
CONTROL PROBLEMS IN ASTRODYNAMICS
<strong>IAA</strong>-AAS-DyCoSS1-14-03
Francesco Topputo and Franco Bernelli-Zazzera
A method to solve nonlinear optimal control problems is proposed in this work. The method
implements an approximating sequence of time-varying linear quadratic regulators that
converge to the solution of the original, nonlinear problem. Each sub-problem is solved by
manipulating the state transition matrix of the state-costate dynamics. Hard, soft, and
mixed boundary conditions are handled. The presented method is a modified version of an
algorithm known as “approximating sequence of Riccati equations”. Three sample problems
(low-thrust orbital transfer, formation flying control, and low-thrust station keeping) are
treated to show the effectiveness of the method, whose limitations are also discussed.
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15:00
<strong>IAA</strong>-AAS-DyCoSS1-14-04
TIME-OPTIMAL ATTITUDE MANEUVER OF A
SPACECRAFT WITH INERTIAL ACTUATORS
Mikhail Levskii
Concrete problem of spacecraft's terminal reorientation is solved. The case when a change of
spacecraft’s angular position is done with use of inertial executive devices is considered.
Interesting variant, when spacecraft angular momentum is limited and turn time is
minimized, is researched. The conditions of optimality is written in analytical form (using
Pontryagin’s maximum principle and quaternion method), and the properties of optimal
spatial turn are studied. Key relations and equations for optimal motion which specify the
variation of rotation parameters are given. Results of mathematical simulation of spacecraft
motion dynamics under the designed control method are presented.
15:20
FUEL OPTIMAL TRAJECTORY DESIGN USING SOLAR-ELECTRIC-
PROPULSION UNDER POWER CONSTRAINTS
AND PERFORMANCE DEGRADATION
<strong>IAA</strong>-AAS-DyCoSS1-14-05
Peng Zhang, Junfeng Li, Hexi Baoyin, and Yang Chen
In this paper the fuel optimal trajectories using solar-electric-propulsion is considered with
power constraints and performance degradation. The problem is solved applying indirect
optimization method derived from the calculus of variations and Pontryagin’s maximum
principle. For solar array power degradation with time, three different models are
hypothesized in the analysis for three typical missions, which are linear, positive and
negative exponential degradations respectively. The significance is obvious considering the
comprehensive power model. It is economical compared with the mission adding an rough
degradation margin to power budget, and increases the reliability compared with those
considering no power degradation with time.
Coffee-break 15:40-16:00
16:00
<strong>IAA</strong>-AAS-DyCoSS1-14-06
POWER-LIMITED SOLAR ELECTRIC PROPULSION
TRAJECTORY OPTIMIZATION
Yang Chen, Hexi Baoyin, Junfeng Li, and Fanghua Jiang
The Solar Electric Propulsion trajectory optimization with power constraints is investigated.
The fuel-optimal control problem is solved via an indirect method. The optimal control laws
of the thruster input power and the thrust direction are derived. The solar array power
varies with the heliocentric distance, and accordingly the Euler-Lagrange equations hold two
different forms. Therefore, the detection of heliocentric distance with fixed step integration
is presented to ensure accuracy. Moreover, The normalization of initial costates and
switching function detection are employed to improve computational efficiency and
accuracy. The trajectory towards Apophis is discussed to substantiate the efficiency of these
techniques.
16:20
<strong>IAA</strong>-AAS-DyCoSS1-14-07
ANALYSIS OF INTERPLANETARY SOLAR SAIL TRAJECTORIES
WITH ATTITUDE DYNAMICS
Andreas Borggräfe, Andreas Ohndorf, and Bernd Dachwald
This study presents a new approach to the problem of optimal control of solar sails for lowthrust
trajectory optimization. The objective was to find the required control torque
magnitudes in order to steer a solar sail in interplanetary space. A new steering strategy,
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1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
controlling the sail with generic torques applied about the spacecraft body axes, was
integrated into the existing low-thrust trajectory optimization software InTrance. A three
rotational degree-of-freedom rigid-body attitude dynamics model was implemented to
represent the solar sail in space. The resulting control torques were investigated, as they
pose primary requirements to a real on-board attitude control system.
16:40
AUTOMATIC CONTROL SYSTEMS FOR ESMO SATELLITE
USING LQR CONTROL STRATEGY
<strong>IAA</strong>-AAS-DyCoSS1-14-08
Mihaela Raluca Stefanescu, Claudiu Lucian Prioroc,
and Constantin Burdusel
This paper studies the attitude control of ESMO (European Student Moon Orbiter) satellite.
ESMO is a European project developed by ESA which has the purpose of initializing young
students with ESA terminology, team work and not least with space research. As well aspects
of attitude representation and kinematics and dynamics in terms of quaternions and angular
rates will be finding in this paper. The control system is based on the Linear Quadratic
Regulator (LQR) theory. This study presents a new approach to the problem of optimal
control of solar sails for low-thrust trajectory optimization.
17:00
HIGH ORDER OPTIMAL FEEDBACK CONTROL OF LOW-THRUST
ORBITAL TRANSFERS WITH SATURATING ACTUATORS
<strong>IAA</strong>-AAS-DyCoSS1-14-09
Pierluigi Di Lizia, Roberto Armellin,
Alessandro Morselli, and Franco Bernelli Zazzera
A novel nonlinear optimal feedback control algorithm for low-thrust space transfers is
presented. A minimum fuel mass optimal control problem is studied, including control
saturation in the problem formulation. Following the reduction of the optimal control
problem to a two-point boundary value problem, differential algebraic techniques are
applied to obtain a high order expansion of the optimal thrust direction and switching times
with respect to either initial or terminal conditions. The computation of optimal feedback
control laws with saturating actuators in relatively large neighbourhoods of the reference
trajectory is then reduced to the fast evaluation of high order polynomials.
17:20
PONTRYAGIN'S MAXIMUM PRINCIPLE FOR IMPULSIVE CONTROL
PROBLEMS WITH MIXED CONSTRAINTS
<strong>IAA</strong>-AAS-DyCoSS1-14-10
Aram Arutyunov, Dmitry Karamzin, and Fernando Pereira
Necessary conditions in the form of Pontryagin's Maximum Principle are derived for impulsive
control problems with mixed constraints. A new mathematical concept of impulsive control is
introduced as a requirement for the consistency of the impulsive framework. Additionally,
this control concept enables the incorporation of the engineering needs to consider
conventional control action while the impulse develops. Ekeland's Variational Principle and
Lebesgue's discontinuous time variable change are used in the proof. The article also
contains an example showing how such impulsive controls could be relevant in actual
applications.
Session 15:
Spacecraft Guidance, Navigation, and Control III (Biombo)
Chair: Mª. Margarida FERREIRA
14:00-18:00
14:00
<strong>IAA</strong>-AAS-DyCoSS1-15-01
WEAK EXPONENTIAL STABILITY VIA AVERAGING
METHOD AND APPLICATIONS
Ricardo Gama and Georgi Smirnov
43

FINAL PROGRAM
The averaging method is one of the most powerful methods used to analyse differential
equations appearing in the study of nonlinear problems. The Averaging method has also been
successfully applied in control problems. In this work we propose a new approach to
conclude exponential stability of time varying systems applying averaging method. We also
show that this result can be applied to differential inclusions generated by control systems.
Some practical applications examples of the presented results will be also presented,
illustrating it's applicability for solving space dynamics problems.
14:20
GUIDANCE, NAVIGATION AND CONTROL
FOR PROPULSIVE SOFT LANDING ON MARS
<strong>IAA</strong>-AAS-DyCoSS1-15-02
Andres Guillermo Molano, Enrico Canuto,
and Ferdinando Cometto
This paper describes the Guidance, Navigation and Control (GNC) for the final propulsive soft
landing phase of the entry, descent and landing of an interplanetary lander on Mars. The soft
landing GNC uses and extended model of the vehicle attitude and horizontal motion to build
a full state (attitude and translation) reference trajectory from the current state to the
desired (possibly varying) landing site. Once the reference trajectory is computed, a tracking
algorithm estimates and compensates the external disturbances in real time. The principles
of the algorithms are provided together with the outcome of realistic Montecarlo
simulations.
14:40
DEVELOPMENT AND VALIDATION IN REAL TIME ENVIRONMENT
OF A GNC SYSTEM FOR PLANETARY RV
<strong>IAA</strong>-AAS-DyCoSS1-15-03
Juan C. Bastante, José Vasconcelos, João Monteiro,
Gabriele Bellei, Andrea Mafficini, João Dinis, and Sohrab Salehi
The proposed paper shows the development, up to TRL5 (‘Component and/or breadboard
validation in relevant environment), of a GNC system for RV relying mainly on camera
measurements. The laboratory environment used for this TRL upgrade, a complete Real Time
Test Bench with Hardware in the Loop, is installed in the Laboratory of Optics and Laser of
the Faculty of Sciences of the University of Lisbon. The main conclusions derived from the
test campaign executed last summer are presented in full detail
15:00
CONTROL SYSTEM DESIGN OF KOREA
LUNAR LANDER DEMONSTRATOR
<strong>IAA</strong>-AAS-DyCoSS1-15-04
Dong-Young Rew, Sang-Wook Kang,
Gwanghyeok Ju, and Sang-Ryool Lee
A series of lunar exploration program consists of lunar orbiter and lunar lander missions are
planned in Korea during 2020s. Korea Aerospace Research Institute (KARI) is performing
conceptual design phase researches for the lunar explorers. A lunar lander which will be used
for demonstrating lunar landing technologies and will be used for testbed for new technology
is being developed. In this paper, design concept of lunar lander demonstrator is introduced.
Trade-off details in configuring thrusters for descending and attitude control are discussed.
Guidance and control logic for ascending, hovering and horizontally moving, descending, soft
landing are presented with simulation results.
15:20
PAYLOAD MAXIMIZATION OF LTGL WITH
NORMAL FORCE DISTURBANCE
<strong>IAA</strong>-AAS-DyCoSS1-15-05
Rock Jeng-Shing Chern and Chie-Ming Huang
This paper presents the research in maximizing the payload of the linear-tangent guidance
law (LTGL) under the normal force disturbance by using the parameter optimization method.
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1 ST <strong>IAA</strong> <strong>Conference</strong> on Dynamics and Control of Space Systems – DyCoSS’2012
It is intended to be developed for the Taiwan small launch vehicle. A three-stage launch
vehicle is considered. It is capable of inserting a 1,000 kg satellite into a 647 km altitude
circular orbit at the nominal condition. The study is performed for the payload maximization
of the LTGL by following the parameter optimization method with the normal force
disturbance included. Precise orbit insertion with payload maximized under nornal force
distrubance can be fulfilled.
Coffee-break 15:40-16:00
16:00
NONLINEAR HIERARCHY-STRUCTURED PREDICTIVE CONTROLLER
DESIGN FOR REENTRY VEHICLE
<strong>IAA</strong>-AAS-DyCoSS1-15-06
Peng Wang, Luhua Liu, and Jie Wu
This paper presents the research in maximizing the payload of the linear-tangent guidance
law (LTGL) under the normal force disturbance by using the parameter optimization method.
It is intended to be developed for the Taiwan small launch vehicle. A three-stage launch
vehicle is considered. It is capable of inserting a 1,000 kg satellite into a 647 km altitude
circular orbit at the nominal condition. The study is performed for the payload maximization
of the LTGL by following the parameter optimization method with the normal force
disturbance included. Precise orbit insertion with payload maximized under nornal force
distrubance can be fulfilled.
16:20
IMPROVING NAVIGATION SYSTEM OF A LAUNCHER THROUGH
THE USE OF ON-BOARD AUTONOMOUS STATE ESTIMATION
<strong>IAA</strong>-AAS-DyCoSS1-15-07
Simone Battistini
Traditional navigation systems of launch vehicles are mainly based on inertial navigation,
referring to inertial measurements by an IMU. This considerably affects the accuracy of state
estimation and it has also relevant implications on costs. The idea for improving modern
systems is to rely also on external measurements, such as GPS. This implies to have an
accurate algorithm for state estimation, which optimizes the fusion between different
measurements. Unscented Kalman Filter (UKF) is a very interesting tool for for non linear
state estimation. In this work the implementation of this algorithm applied to the case of a
rocket is presented.
16:40
DRAG DERIVED ALTITUDE FOR IMPROVED NAVIGATION
ACCURACY IN MARS ENTRY DESCENT AND LANDING SYSTEM
Emanuele Di Sotto, João Branco, Luis Guerreiro,
<strong>IAA</strong>-AAS-DyCoSS1-15-08
Nuno Paulino, and Jean François Hamel
Precise on-board navigation accuracy is a key issue for Mars Entry Descend and Landing
Systems. During the entry phase, a probe can count only on the accuracy of a proper Inertial
Navigation System, whose performances are limited by: sensors accuracy, initialization at
Mars Entry Interface, platform alignment, instability of vertical channel. The vertical
channel stabilization is particular important and drives the altitude navigation accuracy up
to the main parachute deployment. This paper shows how an on board stored aerodynamics
database, along with an estimation of the Mars density profile, allow stabilizing the vertical
channel through the Drag Derived Altitude Technique.
17:00
<strong>IAA</strong>-AAS-DyCoSS1-15-09
GUIDANCE AND CONTROL FOR THE ATMOSPHERIC GUIDED
ENTRY OF LOW LIFT-TO-DRAG ENTRY VEHICLE BASED
ON THE EMBEDDED MODEL CONTROL
Jose Alejandro Ospina, Enrico Canuto, and Buonocore Marcello
45

FINAL PROGRAM
The translational control law of a biconic atmospheric entry capsule using the bank angle as
the command are described. The algorithm is separated into a path-planning and a
reference-path-tracking algorithm. The path-planning algorithm computes the entry
trajectory starting from the navigated state at the entry point and reaching the desired
parachute deployment point. The atmospheric and aerodynamic dispersion are compensated
in real-time using the Embedded Model Architecture in which external disturbances are
estimated and compensated in real time. The main concepts for both algorithms are
presented and the results of a realistic Montecarlo simulation campaign are provided.
17:20
A NEW REENTRY TRAJECTORY DESIGN APPROACH FOR LIFT-
CONTROL FLIGHT VEHICLE BASED ON CRUISING
PARAMETERS DESIGN CONCEPT
<strong>IAA</strong>-AAS-DyCoSS1-15-10
Zhengchun He, Fengqi Zheng, Weiqi Qian, and Kaifeng He
In this paper, a new trajectory design approach for lift-control reentry flight vehicle based
on cruising parameters design concept is proposed. The cruising parameter could be
aerodynamic overload, dynamic pressure, or aeroheating rate at stagnation point. For a high
speed reentry flight vehicle, based on this trajectory design approach, it could firstly cruise
at an appropriate aeroheating rate, then at an appropriate aerodynamic overload, and finally
at an appropriate dynamic pressure. As an example, the entry trajectories designed for a
certain lift-control reentry vehicle based on the new approach are presented in this paper,
too.
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