Программа 17-го Международного симпозиума по динамике космического полета
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16
June 2003 Monday |
17
June 2003 Tuesday |
18
June 2003 Wednesday |
19
June 2003 Thursday |
20
June 2003 Frisday |
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8:00-09:40 Registration |
8:20-10:00,
Session 4 Low
Thrust Trajectories |
8:20-10:00,
Session 8 Attitude
Dynamics, Estimation, Control |
8:20-10:00,
Session 11 Formation
Flying, Constellation. Maneuvers
Design, Guidance,
Control |
8:20-10:00,
Session 15 Celestial
Mechanics |
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09:40-10:00 OPENNING |
10:00-10:20,
Break |
10:00-10:20,
Break |
10:00-10:20,
Break |
10:00-10:20,
Break |
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10:00-12:00,
Session 1 Flight
Dynamics Operations |
10:20-12:00,
Session 5 Orbit
Design, Mission
Analysis |
10:20-12:00,
Session 9 Flight
Dynamics Software |
10:20-12:00,
Session 12 Orbit
Determination |
10:20-12:00,
Session 16 Low
Thrust Trajectories |
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12:00
– 13:30, Lunch |
12:00
– 13:30, Lunch |
12:00
– 13:30, Lunch |
12:00
– 13:30, Lunch |
12:00
– 13:30, Lunch |
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13:30-15:30, Session 2 Space Debris Dynamics, Space Patrol Systems |
13:30-15:30, Session 6 Deep Space Missions |
13:30-15:30, Session 10 Attitude Dynamics, Estimation, Control |
13:30-15:30, Session 13 Orbit Determination |
Excursion |
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15:30
- 15:50, Break |
15:30
- 15:50, Break |
Excursion |
15:30
- 15:50, Break |
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15:50-17:30, Session 3 Autonomous Navigation, GPS Navigation |
15:50-17:30, Session 7 Atmospheric Entry |
15:50-17:30, Session 14 Attitude Dynamics, Estimation, Control |
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Symposium Dinner |
16 June 2003, Monday
Session 1
Flight Dynamics Operations
10:00-12:00
Etienne DUCROTTE, Jean-Paul BERTHIAS, Beatrice DEGUINE, Francois
DESCLAUX, Jacques FOLIARD, Laurent FRANCILLOUT, Hubert FRAYSSE, Jean-Francois
GOESTER, Marc ROSSI, Stephane ROUSSEAU, Francois LAPORTE
CNES, France
Following
a failed Proton launch, the satellite of telecommunications Astra1K was
released on an orbit of 175 km altitude and 51 degree inclination. This
presentation intends to describe the operations that have come to a controlled
re-entry early on the 10/12/2002. The work began with two maneuvers to raise
the orbit up to an altitude of 290 km. From this time, analyses were undertaken
to allow a decision process that came to decide to de-orbit the spacecraft. The
presentation focuses on the orbit propagation troubles, the available attitudes
to perform the maneuvers, the studied scenarios and the operational sequences.
Frank Dreger
ESA-ESOC (TOS-GFT), Germany
INTEGRAL - ESA's
first gamma-ray observatory - was launched on 17. October 2002. Together with
XMM-Newton it is the second, currently flying observatory mission, which
requires real-time operations support during 24 hours per day and 7 days per
week.
The paper describes the
individual tasks, the user requirements, and the experience gathered during the
first months of INTEGRAL routine operations. The paper also addresses the
limitations experienced when Flight Dynamics tools are provided to non-Flight
Dynamics staff.
Rainer Kresken
EDS, Germany
ESA's INTEGRAL is
equipped with a reaction wheel assembly consisting of four wheels for attitude
control. A thruster-based reaction control system allows momentum bias
maneuvers. The presentation describes the measures to comply to the wheels'
operational constraints with optimum operational efficiency and minimum
interference with scientific observations and gives a comparison with the
operations of ESA's XMM x-ray observatory
Robert A. McCUTCHEON
Hubble Space Telescope Science Institute / Computer
Sciences Corp, USA
This
paper describes a new platform-independent solar-lunar-planetary (SLP)
ephemeris package based on theories and models developed at the Bureau des
longitudes (BDL) Institut de mecanique celeste et de calcul des ephemerides.
For the Moon the new package uses ELP 2000-82B, a semi-analytic theory that
agrees with the Jet Propulsion Laboratory (JPL) LE200 ephemeris to within 0.03
arcseconds. For the Sun and planets the new package uses the BDL Poisson series
method. The coefficients for these series were determined via frequency
analysis of the JPL DE403 ephemerides from 1900 to 2100. The largest
differences between geocentric positions computed via the BDL Poisson series
and positions computed directly from DE403 are on the order of a few
milli-arcseconds. The BDL SLP package also uses the Poisson series method to
provide ephemeris data for the minor planets Ceres and Vesta. All coefficients
are read from flat ASCII files, thereby making the new SLP package usable
without modification on both VMS and Unix platforms. The new package has
replaced the JPL-based SLP package in all Hubble Space Telescope Payload Operations
Control Center Applications Software Support systems.
G. Ziegler
Electronic Data Systems, European Space Operations Centre, Germany
Unlike other space agencies ESA ground
stations are not equipped with special first acquisition antennas which ease
the first acquisition because of their large beam width. To overcome this
disadvantage some standard search strategies are used which essentially scan
the sky spirally around the nominal pointing direction with increasing radii.
These standard strategies are completed by special mission dedicated strategies
sometimes in order to detect the spacecraft as soon as possible. In this
presentation three basic techniques, namely spiral search, waiting point method
and horizon-scan technique, will be described as well as their potentially
expected benefit for the missions Proba and Integral.
B. Santa-Cruz
GMW at ESA/ESOC, Germany
Detlef Sieg
EDS at ESA/ESOC, Germany
It is of high economic interest for ESOC
to know the workload of its globally distributed network of ground stations.
Then as many support requests as possible can be committed. For this an
automated process and an appropriate software tool supporting short-term
scheduling and long-term forecast are necessary. The user requirements and the
design of the tool are described and a long-term analysis of some ESA missions
is illustrated.
Session 2
Space Debris Dynamics
13:30-15:30
Carmen Pardini, Luciano Anselmo
ISTI/CNR, Italy
During the 1990s, it became clear that
spacecraft and upper stage breakups contributed to the GEO debris environment
while, recently, optical observations have confirmed the presence of a large
population of decimeter sized particles, probably generated by a number of
undetected explosions. Hoping to provide the optical observers with useful
information and clues in identifying and characterizing the past explosion
events near the geosynchronous region, a fragmentation in GEO of a typical
communication spacecraft has been simulated, while assuming a reasonable range
of the fragments' ejection velocities. Each produced debris cloud has been propagated
for 72 years, saving the results at intermediate time steps. The final outcomes
- mainly represented as snapshots, at a given post-explosion time, in the
orbital elements space - show the long-term evolution of each debris cloud as a
function of debris size and ejection velocity.
A.I. Nazarenko
Rosaviakosmos Space Observation Center, Russia
The problem of estimating the possibility
of collisions of space debris (SD) is relatively new. This paper presents the
results of investigations of key problem of collision hazards evaluation: (a)
current SD environment, (b) technique for collision probability evaluation, (c)
characteristics of the relative flux of SD, (d) account of shape and
orientation of typical spacecraft modules, (e) characteristics of the new
software for the SD impact hazards evaluation. The test penetration probability
calculations are performed; the related results are presented and analysed.
Z.Khutorovskiy, V.Boykov, V.Agapov, N.Sbytov,
A.Samotokhin
Keldysh Insitute of Applied Mathematics, KIA Systems
New geostationary objects (GEO) catalogue
had developed. It is based on complex informational model in the Oracle 8 RDBMS
environment, highly efficient orbital motion and orbital determination models.
Special tasks developed for measurements and orbits identification. Optical
measurements obtained by ground stations of the Russian Academy of Sciences
during period of 1975-2002 had processed as well as TLE archive for GEO
objects. Unique results on large volume of TLE accuracy estimation obtained.
The paper will present general architecture of the Catalogue, description of
motion model and orbit determination model (including initial orbit estimation
based on optical measurements only), results of data processing, general
characteristics of GEO objects environment.
I. E. Molotov, G. Tuccari, B. N. Lipatov, X.Y. Hong
KIAM, Russia
The method allows to measure all movement
parameters of space objects by combining the VLBI and radar techniques. The
system consists of powerful ground radar (Evpatoria), three VLBI radio
telescopes (Noto, Bear Lakes, Shanghai) and the correlator in Noto. The
received telemetry or radar echo signals are transferred from radio telescopes
to processor through Internet for cross-correlation in near real-time. This
system is acceptable for investigations of the dynamics properties of space
debris objects, operating satellites, deep space spacecrafts, asteroids and
planets. This work is supported by INTAS 2001-0669 and RFBR 02-02-17568 grants.
T.M.Eneev, G.B.Efimov, R.Z.Akhmetshin, G.S.Zaslavsky
KIAM, Russia
Conception of space patrol system for
discovering hazard asteroids is considered. The patrol system with several
spacecrafts (with space telescopes onboard) on the Earth orbit allow to
discover during 5-6 years a lot of asteroids (with diameter ? 100 m)
approaching the Earth orbit. Conception may be considered as the following step
to existing programs of optical observation from the Earth of dangerous objects
(with diameter 1 km)
Session 3
Autonomous Navigation, GPS Navigation
15:50-17:30
J.L. Gonnaud, R. Fayard,
ASTRIUM, France
The Automated Transfer Vehicle (ATV) is a
European cargo transfer vehicle designed to carryout ESA in-orbit replenishment
missions to the International Space Station. The ATV first flight is currently
planned in 2004. Due to its very specific rendezvous mission toward a manned
orbital facility, the Navigation function has to fulfil drastic reliability and
availability requirements. These requirements are designed to ensure an
unequalled level of Safety with respect to collision hazards regarding the ISS.
This paper provides a deeper insight into some innovative solutions that have
been implemented in the ATV navigation (attitude and relative position) to
allow the fulfilment of these very stringent GNC requirements. The latest
status of validation of the flight software is presented, including
performances validation through statistical simulation campaigns.
Valcir Orlando, Helio Koiti Kuga,
INPE, Brazil
A survey of the research on autonomous
orbit control systems carried out at INPE is presented. INPE started working in
this area in 1995 when an study on the feasibility of an autonomous control
concept of the orbit longitude phase drift (DL0) was performed, in cooperation
with the French Space Agency (CNES). The use of DIODE (French autonomous navigator)
simulated orbit observations was considered. Thereafter, following a world wide
trend, the research work was re-directed to the investigation on the use of the
GPS (Global Positioning System). At first, only the direct use of the coarse
GPS navigation (geometric) solution was considered. In order to improve the
results of the GPS based autonomous control, a GPS simplified navigator was
developed and included in the control procedure. Samples of the results
obtained in each phase of the research work are presented and commented in the
work.
E.L. Akim, D.A. Tuchin,
Keldysh Institute of Applied Mathematics RAS
The creation of a autonomous spacecraft's
control system with use GPS requires the statistical analysis of GPS error
measurements. This work describes a most GPS receiver error measurements. The
pseudorange of the C/A-Code measurements were processed and analysed to obtain
the statistical performances of the three main sources of GPS error: GPS
satellite errors (ephemeris and satellite clock), the Earth atmosphere errors
(ionosphere and troposphere), user receiver errors (frequency drift,
pseudo-noise sequence phase drift, signal detection time). It allowed to exclude
the mentioned errors from common error budget to construct the statistical
model of proper error of pseudorange measurements and to study the influence of
their separate components on user state vector determination. The comparison of
predicted measurement precision with real pseudorange of the C/A-Code was
conducted out.
M.Yu.Beliaev,
DN.Rulev, E.S.Medvedev,
RSC
Energia, Russia,
V.V.Sazonov
Keldysh
Institute of Applied Mathematics RAS, Russia
The GPS receivers are installed on the
Russian and American Segments of ISS to realize high-precision navigational
support of its flight. The receiver in the Russian Segment is ASN-2401. Its
trial was begun in 2001. It is intended for using in the structure of the
motion and navigation control system of the Service Module. Our paper presents
some estimations of actual accuracy of the receiver, which were obtained by
smoothing of GPS-measurements on time intervals of various lengths.
17 June 2003, Tuesday
Session 4
Low Thrust Trajectories
08:20-10:00
Per Bodin
Swedish Space Corporation, Sweden
Abstract
SMART-1 is the first in a series of low-cost
scientific missions by the European Space Agency. The mission of SMART-1 is to
demonstrate the use of electric primary propulsion in a low-thrust transfer
orbit from earth orbit into lunar orbit. The mission consists also of several
scientific measurements mainly for use in lunar orbit. The Swedish Space
Corporation is the prime contractor for the spacecraft. The presentation will
give an overview of the Attitude and Orbit Control System on SMART-1. In
particular, the presentation will focus on the constraints and drivers imposed
by the electric propulsion engine. Special attention will also be given to the
elaborate on-board autonomy as well as the extensive use of auto-generated code
in the development of the on-board software.
J.
Fourcade
CNES,
France
Not
available
Gennadiy Fedotov, Mikhail Konstantinov
Moscow Aviation Institute, Russia.
Transport opportunities of delivery of a
spacecraft to Mercury with the use of launcher "Soyuz", chemical
upper stage "Fregat" and solar electric propulsion upper stage
(nominal electric power 6.75 kW) is analysed. The outcomes of research are:
-
At use of stationary plasma thruster
(specific impulse 2000 s) it's possible to deliver into Mercury's vicinity the
spacecraft with mass about 800 kg at transfer duration 785 day;
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Use of ionic thruster (3000 s) allows to
increase this mass on 250 kg at transfer duration 1050 day;
-
Using of Venus swingby gives additional
increment of mass on 150 kg.
Alexander A. Sukhanov,
Space Research Institute, Russia
A low thrust transfer is considered. It
is assumed that the thrust direction is subject to a constraint. This
constraint may be caused by specific features of the spacecraft stabilization
mode and attitude control system and in a general case is a function of time
and the spacecraft state vector. Mathematically the constraint is given by a
manifold, which the thrust direction belongs to. It is shown that the optimal
thrust is directed along the projection of the Lawden's primer vector onto the
manifold. Both limited power and constant exhaust velocity cases are
considered. Some examples of the constraints are given. The transporting
trajectory method is applied to the case of the constrained thrust direction;
this method also gives a sufficient condition of the possibility of transfer
with a specified constraint.
Mikhail Konstantinov
Moscow Aviation Institute, Russia
The method of optimization of
multirevolution trajectories of transfer from elliptical orbit into noncoplanar
circular orbit is developed. The basic idea is use of the some model problem of
optimal control. This model is chosen so, that solution of model optimization
problem does not have any considerable difficulties. This solution is
considered base for determination of the optimal solution of a common problem.
The basic advantages of a developed method are:
An absence of a problem of convergence of
iterative procedures of the solution;
Representation of the solution of problem
of an optimal control as synthesis.
Session 5
ORBIT DESIGN, MISSION ANALYSIS
10:20-12:00
Roberta Mugellesi-Dow, Guy Janin
ESOC, Germany
Natan Eismont
IKI, Russia
Abstract
After having recalled characteristics of
orbits for space astronomical observatories and the particular advantage of
selecting synchronous orbits, this paper highlights the novel orbital aspects
of the INTEGRAL mission from the Flight Dynamics point of view: launch on a 72d
orbit using a Russian Proton rocket equipped with the upper stage Block DM, the
special preparation undertaken in ESOC to counter-act non-nominal performance
of the launcher upper stage, and the orbit manoeuvres sequence dependency on
various constraints. The paper outlines the operational and technical
constraints that had to be respected in the definition of the orbit design and
during the implementation, and the major orbital activities that were performed
during the launch and early orbit phase.
N. Eismont, V. Khrapchenkov,
Space Research Institute, Russia
G. Janin, R. Mugellesi-Dow
European Space Operations Centre, Germany,
Not available
A.A.Boyarchuk, A.V.Bagrov,
Institute of Astronomy of the Russian Academy of Sciences, Russia,
K.M.Pichkhadze, V.K.Sysoev,
Lavochkin Science & Industry Corporation, Russia
A technical design of the OSIRIS
two-based optical space interferometer is based on the carbon-fibre technology
and on the newest Russian achievements in laser metrology and light sensors.
The whole instrument was designed as a payload to the Russian Segment of the
ISS and it consists of two blokes that will be brought to the ISS by transport
spacecraft "Progress" and installed there by astronauts. Short-live
devices of the OSIRIS interferometer - light sensor and metrological laser - are
replaceable for fast repairs in the open space conditions. Due to the
carbon-fibre strongly constructed case and to the absence of reserved devices
the total mass of the OSIRIS interferometer will not exceed 250 kg.
The two-based interferometer OSIRIS consists
of four off-axis telescopes, four adjustable delay lines, and two light sensors
of the MAMA-type and metrological system. Besides that every telescope will
have its own guide for precision pointing to the target stars. The whole
instrument will use its own independent to the ISS pointing and tracking
systems. It will be mounted to the RS ISS by magnetic controlled suspension
device.
The OSIRIS Project was admitted to the
Russian National Space Program as a payload to the RS ISS at the promoted stage,
and it might be launched at 2002-2003 years.
Sanguk Lee, Jae Hoon Kim, Seong Pal Lee
Communications Satellite Development Center, ETRI, Korea
In this paper, we suggest communications
satellite system placed in three Lagrange points, L3, L4, and L5, of the
restricted three-body problem of Earth-Moon system. Generally, geostationary
satellites are used for communications so far. Recently, LEO satellite
constellation is another choice of communications system. The proposed system
which is alternatives of limited geostationary orbit resources, has some weak
points such as long distance from the Earth, much cost to launch satellite,
long delay time, required more power, and so on. It has good points like less
efforts(fuel) for station keeping, less eclipses, etc. In this paper, some
analyses about the proposed system such as characteristics of orbits, eclipses,
perturbations, link requirements for communications, pointing accuracy, ground
station operation concept, possible services to be provided, and so on will be
presented.
M.Lavagna, A.E.Finzi
Dipartimento di Ingegneria Aerospaziale, Italy
This paper proposes a method to support
decisions to be taken within a concurrent approach for the space system
preliminary design: the defined architecture is based on a Multi-Criteria
Decision Making approach mixed with methodologies coming from the Approximate
Reasoning domain. The method here presented is focused on saving analysts' time
and effort by addressing the decisions they have to make during the preliminary
design process to solve inconsistency and bottlenecks risen from the parallel
design of several subsystems. Moreover, different configurations can be
considered at a time ranked according to a pre-selected criteria vector. From a
theoretical point of view, revisited Genetic Algorithms are applied, within
each single subsystem design domain, in order to obtain a non-dominated
solution set to be considered for solving conflicting design at system level;
the Analytical Hierarchical Process - supported by dedicated blocks implemented
by the Fuzzy Logic approach has been selected as the fittest tool to simulate
the causal relationships between variables and objectives, normally prerogative
of the analysts' experience in the spacecraft design domain, within the system
level point of view. Simulations showed the ability of the algorithm to find
conflicts and suggest a set of subsystem parameters to be tuned to converge -
consistently with a user defined cost functions vectors - to a final spacecraft
configuration; the tool runs in real-time with the on-going space system design
process, in order to support the team leader in making decisions. A comparison
with a completely transparent optimization process, implemented by MOGA,
highlighted the capability of the proposed approach to move towards the final
Pareto front solution.
Session 6
Deep Space Missions
13:30-15:30
Stephanie Delavault, Laurent Francillout, Denis Carbonne,
CNES, France
The challenging phase of a NETLANDER-like
mission is the successive release of several Landers to the Martian soil. From
a navigation point of view, this deployment requires a high accuracy level for
the Netlander entry trajectory with only few days of tracking schedule and a
long free-flying phase. This paper presents covariance analyses over the
complete Earth to Mars cruise and Mars approach phases of such a mission
performed to determine the impact on navigation performance of parameters such
as the addition of DDOR measurements, tracking data schedule, maneuver
execution errors: Conclusions are drawn on navigation needs and an assessment
is made of the robustness of navigation performance for the Netlander
deployment phase.
M.Delpech, J-B.Dubois,
CNES, France,
J.E.Riedel, J.R.Guinn
JPL, USA
The Mars Premier mission that was to be
flown in 2007 by CNES in cooperation with NASA/JPL included a rendezvous
experiment to be performed in Mars orbit to validate key technologies
applicable to a Sample Return mission. The paper presents an overview of this
experiment outlining mission design, system implementation, operational aspects
and validation work achieved before the program was stopped in October 2002.
Tuchin A.G., Akim E.L., Botkin A.B., Stepaniants V.A.
Keldysh Institute of Applied Mathematics, Russia
Ruzskiy E.G.,
Lavochkin Association, Russia
The goal of the Phobos sample return
project is Phobos soil delivery to the Earth. This paper describes the Mars
satellite phase of the mission. The task of this phase is to lead the
spacecraft to specific area located at altitude 40-60 km above the Phobos
surface with the accuracy satisfied to autonomous lending system requires. The
questions of ballistics, navigation and flight control are considered. Two
orbits are used during the Phobos approach phase - an observation orbit and a
quasi-synchronous orbit from which landing to the Phobos surface is performed.
G.S. Zaslavskiy, V.G. Zharov, A.V. Chernov
KIAM, Russia
The optimal spacecraft flight to Mars
satellite orbit with using electric propulsion system is studied. The effect of
the attractions of the Sun, the Moon and the Solar system planets on the
spacecraft motion in during of the all flight time is taken into account.
N.M. Ivanov, Yu.F. Kolyuka
Mission Control Center, Russia
The results of the work for the
construction and justification of the flight schemes for the Martian manned
expedition that was done within the frame of the ISTC grant are given. This
expedition implies round transfer Earth – Mars – Earth as well as staying the
manned vehicle on near-to-Mars orbit and landing of astronauts to the Mars
surface.
The flight schemes for two conceptual
variants of the Expedition – two-vehicles scheme variant (including the flights
of manned and cargo ships) and one-ship scheme variant were developed and
analyzed.
The calculations of the mission schemes
were based on the special methods and software tools that allow considering
both the low-thrust and large-thrust flights. The short description of these
methods is represented.
E.P. Molotov, I. E. Molotov
Russian Institute of Space Device Engineering, Russia
Russian Deep Space Network was based on
three large antennas - RT-70 in Evpatoria and Ussuriysk, and RT-64 in Bear
Lakes. Russian deep space and high-apogee missions were controlled with these
dishes. The antennas were temporary stopped after the loss of the Mars-96. The
operations of Evpatoria and Bear Lakes were renewed for radio astronomy goals.
Both antennas participate in VLBI observations and radar researches of space
debris under LFVN project. The limited financing of Ussuriysk RT-70 was
continued recently under preparing the Phobos sample return mission. The
partners interested in joint usage of these antennas are sought.
Session 7
Atmospheric Entry
15:50-17:30
Yu.G.Sikharulidze, A.N.Korchagin,
Keldysh Institute of Applied Mathematics RAS, Russia
Not available
Andrey
Nazarenko,
Space
Observation Center, Russia
Vasiliy Yurasov,
Space
Research Center "Kosmos", Russia
Perspective direction of an accuracy
increasing of satellite orbit determination and prediction for low earth orbit
(LEO) satellites is the organization of the upper atmosphere monitoring, i.e.
analog of a weather service in the lower atmosphere. Our idea of the upper
atmosphere monitoring is based on the usage of the available satellite
atmospheric drag data on catalogued LEO satellites. These data are obtained as
a result of regular satellite observations and updated some times per day.
The variations of atmospheric density for
GOST-25645.115-84 model were estimated over 9 month time interval. For these
purpose real orbital data for several hundred space objects with perigee
heights below 600 km were accumulated and processed. The effectiveness of
density correction under various atmospheric conditions was estimated.
The monitoring of the upper atmosphere
density variations would allow to increase the prediction accuracy of LEO
satellites motion and to obtain more accurate atmospheric density estimates
without development a new model.
V.A. Udaloy, N.M. Ivanov, S.I. Kudryavtsev, A.A
Savchenko
Mission Control Center
The operational flight dynamics support tasks
for the spacecraft controlled from the Russian Mission Control Center are
considered. A modern set of tasks accomplished in the flight dynamics support
of the Soyuz TM and Progress M transport vehicles reentry is shown. Specific
features of the Soyuz TMA transport vehicle reentry support are considered.
The Mission Control Center interface with
other organizations is shown. Methods
of solution of reentry flight dynamic support tasks are briefly described.
Examples of reentry support activities in nominal and contingency situations
are given.
The examples of the Mission Control Center participation in international
projects, both in the field of operational flight support and designning
investigations, are shown.
Yu. N. Kaluzhskikh
KIAM, KIA Systems, Russia
In this article the guide algorithm for
an atmospheric re-entry vehicle is described. It is considered a phase of a
re-entry trajectory at the altitude range of 100 to 10 km. The guide algorithm
has two parts. The first one is to generate an initial reference bank angle
sequence to get a re-entry vehicle at a prescribed point. The second one is to
correct the initial reference bank angle sequence in order to reduce the
influence of disturbed atmosphere during re-entry. Results of a mathematical
simulation are presented in this article too. We have considered 3DOF
trajectories with simplified description of the bank angle change. Simulation
results confirm that the accuracy of the developed re-entry algorithm is about
0.7 km.
18 June 2003, Wednesday
Session 8
Attitude
Dynamics, Estimation, Control
08:20-10:00
G.Gienger,
ESA, Germany,
J.B.Palmer
LogicaCMG, Germany
Each observatory is equipped with 2x4 20N
thrusters for orbit-manoeuvres and momentum-dumps. Orbit manoeuvres are
performed with all 4 prime thrusters, each canted 12 degree w.r.t. the
satellite x-axis.
Attitude-control attempts to reduce the
cumulative angular momentum change from all 4 thrusters to zero via
off-modulation, hence one can set up a linear system of 4 equations for the 4
thruster calibration factors.
For the largest orbit-manoeuvres, cumulative
angular momentum for each individual thruster was of order (+25000, +110000,
105000) Nms. Since total cumulative angular momentum is zeroed to an accuracy
of 1 Nms, the ratios of the 4 calibration factors can thus be determined to
1:20000. The relative accu-racy of the sum of the 4 calibration factors is
approximately the relative accuracy of the x-component of the Delta-v as
observed by orbit determination.
E.V. Babkin, M.Yu.Beliaev,
RSC Energia, Russia,
N.I.Efimov, D.A.Zavalishin, V.V.Sazonov
Keldysh Institute of Applied Mathematics RAS, Russia
We present the results of determination
of residual accelerations in Russian Segment of ISS. A quasi-steady
acceleration component was determined using telemetry information about the
station attitude motion. Basing on the information, we reconstruct the station
motion and calculate the acceleration at any given point on its board as a
function of time. We tested this approach using measurements of US low
frequency triaxial accelerometer MAMS and obtained a good agreement. A
vibrational acceleration component was measured by several three-axis
accelerometers IMU-128. They measure accelerations with magnitude more then
0.001 m/s in the frequency range up to 20 Hz.
M. J. Tuttlebee
Science Systems Space Ltd, Germany
The Attitude Determination and Control
Sub-System, which is part of the ESOC On-ground Flight Dynamics System,
provides tools to perform attitude determination for on-line mission
re-planning at the Mission Operations Centre and attitude history products for
the Science Operations Centres.
E. Vinterhav
Swedish Space Corporation, Sweden
The Swedish small satellite Odin has an
extremely short response time has been built into its operations cycle. Odin is
a 3-axis stabilized, high pointing accuracy, sub millimeter, space observatory.
The operation cycle is defined as: receiving user demands, planning and
generation of tele-commands to implement coordinated execution of platform and
payload activities, reconstruction of attitude and orbit trajectories, feedback
to users of payload status and delivery of data. A small efficient team
together with easily accessible, script based, software based on COTS
components facilitates in rapidly meeting new demands on the planning and attitude
reconstruction from the users.
Session 9
Flight Dynamics Software
10:20-12:00
09-1. IMPROVEMENTS IN ROUTINE ORBIT
DETERMINATION AND ORBIT PREDICTION ACCURACY FOR ENVISAT
Dirk Kuijper
Logica plc based at the European Space Operation Center (ESOC), Germany
Not available
09-2. QUARTZ ++, ASTRIUM EVOLUTIVE FLIGHT
DYNAMICS SYSTEM
F. Raballand, J.L Gonnaud
ASTRIUM, France
Based on its large experience, ASTRIUM
has developed a new generation Flight Dynamics System:
QUARTZ++. This new package is designed
for station keeping operations, transfer preparation and LEOP operations for
GEO spacecraft.
QUARTZ++ is currently used by INTELSAT
for LEOP operations and for the routine control of the INTELSAT fleet.
QUARTZ++ will be soon implemented in
INMARSAT control centre for the Flight Dynamics command during transfer and on
station on inclined orbit of the three INMARSAT 4 spacecraft, ASTRIUM Eurostar
3000 class satellites.
QUARTZ++, in addition to services for
autonomous on ground orbit control, will soon be improved to cover Low Earth
Orbit missions including formation-flying control.
09-3. COMPUTER-BASED INSTRUCTION AND REFERENCE
DOCUMENTATION SYSTEM FOR THE ORBIT DETERMINATION PROGRAM
Gerald R. Hintz, Mark Ryne, Michael Watkins, Maureen Kenney, David
Overoye
Jet Propulsion Laboratory, USA
The Orbit Determination Program set has
been used at the Jet Propulsion Laboratory for nearly half a century to enable
precision navigation of interplanetary and earth-orbiting missions and to
support a myriad of scientific investigations. Executing this software package
successfully is a challenging task for experienced personnel and a daunting one
for junior navigators. The effort described in this paper provides a
computer-based, web-enabled instructional and reference tool to aid both
experienced and beginning personnel in the art of doing orbit determination
with this software.
09-4. FOCUS: A NEW CONCEPT ON FLIGHT DYNAMICS
OPERATIONS
Miguel Angel Molina Cobos
Flight Engineering Business Unit Director, Spain
GMV has developed an internal project,
called focus, which is aimed at producing a new generation of Flight Dynamics
systems by exploiting all GMV's past and present experience in the domain and
all the current baseline of GMV's Flight Dynamics systems. Project focus came
through successful requirements definition during the first half of 1999 and
the development phase started in August 1999. focus has very ambitious goals in
mind, in particular the development of a truly generic operational FDS for all
type of satellite missions (including GEO, LEOP, LEO, satellite formations,
constellations, etc.) to be commercialised as a COTS product.
Focus is now progressing in a phased
approach and the first member of this family has been oriented to GEO
satellites. In this context, a sub-project named focusGEO has been conceived to
produce a first version of focus including all fundamental features of the
final system (MMI, graphics, data access, on-line help, process manager, events
logging, automation), in order to provide a new generation Flight Dynamics
product for geostationary satellites. This focusGEO version is now available
and operational at HISPASAT and EUTELSAT.
SESSION
10
Attitude Dynamics, Estimation, Control
13:30-15:30
Karlheinz Spindler
Fachhochschule Wiesbaden, Germany
We consider the problem of slewing a
rigid spacecraft (typically a space telescope) from rest to rest between
prescribed attitudes while avoiding a forbidden direction during the slew.
Using methods from differential geometric control theory, we derive a control
law which minimizes a cost functional penalizing both high angular velocities
and proximity to the forbidden direction and which turns out to be very
suitable for implementation in the on-board software.
Marcello Romano, Brij N. Agrawal
Spacecraft Research and Design Center, USA
This paper presents the attitude dynamics
and control of the bifocal relay mirror spacecraft. The spacecraft consists of
single axis gimbaled receiver and transmitter telescopes with 1.64 m diameter
primary mirrors and fast steering mirrors for fine beam control. The
transmitter telescope has a majority of the spacecraft bus subsystems including
attitude control sensors and actuators. The attitude control system consists of
either Reaction Wheels or Variable Speed Control Moment Gyros, star trackers,
gyros, sun sensors, magnetometers and magnetic rods. Feed-forward control and
quaternion formulation are used. Kalman filter is used to update the rate gyros
biases and the attitude parameters.
Manfred Schneller
German Space Operations Center, Germany
BIRD, a Bi-spectral Infra-Red Detector is
a micro-satellite mission for earth observation. It has been launched on Oct.
22, 2001 on a PSLV-C3 from Sriharikota, India. Nominal operations include
attitude maneuvers to switch attitude from sun pointing for battery charging to
target pointing for data-takes and for downlinking of high rate data. Since
launch attitude determination, prediction and attitude maneuver design is
performed on-ground filling in for an incomplete onboard attitude control
system not yet implemented to its full extent. Between launch and Apr. 2, 2003
over 250 targets on ground have been successfully recorded.
Mats Rosengren,
European Space Operations Centre, Germany
The name PROBA stands for PRoject of
OnBoard Autonomy and the main purpose of the PROBA spacecraft is to test and
demonstrate some autonomy concepts. In addition it carries new space equipment
for tests/demonstration. The PROBA spacecraft was brought into a
sun-synchronous orbit with a local time of descending node of 10:30 on
2001/10/22 by an Indian PSLV launch vehicle taking off from the SHAR centre on
the Sriharikota island (south of India). Its orbit is slightly eccentric with
an altitude varying between 570 and 670 km.
Stéphane
BERRIVIN (*), Vincent COSTES
CNES, France
Michel AUVERGNE
DASGAL, France
Patrick LEVACHER
LAM, France
Not available
19 June 2003, Thursday
Session 11
Formation Flying, Constellation, Maneuvers Design, Guidance, Control
08:20-10:00
Arnaud Boutonnet, CNES, France
Andrei Baranov, KIA Systems
Vincent Martinot, Alcatel Space
Benedicte Escudier, ENSAE-Supaero
Joseph Noailles, LIMA-Enseeiht
Basically the formation initialization
task is a rendezvous. We begin by showing it can be replaced by a transfer
task.
Considering an injection impulse from the
launcher whose amplitude is free, we focus on the optimization of this
maneuver, the formation phase angle and the transfer maneuvers. Two methods
which allow to minimize the propellant consumption are presented: a one impulse
simple strategy and a two impulse optimal strategy. We obtain analytical and
optimal guidance laws from a geometrical approach combined with constraints
given by the Primer Vector necessary conditions of optimality.
The collision risk is assessed with an
analytical distances analysis. Then different parameters are optimized with
respect to the minimum distance without changing the propellant consumption.
Pawel Rozenfeld, Valcir Orlando, Wilson Yamaguti,
INPE, Brazil
The existing INPE's Environmental Data
Collecting System is composed by two small data collecting satellites (SCD1 and
SCD2), inserted in low altitude, 25 degrees inclination orbits; a remote
sensing satellite (CBERS1), positioned in a low altitude 98.4 degrees
inclination orbit; two Data receiving Stations (Cuiaba and Alcantara); about
600 Data Collecting Platforms - DCP, covering a large range of applications,
scattered over Brazilian territory; and a centralized Data Processing and
Distribution Facility. Since the launch of the first satellite until now, the
Data Collecting System experienced a significant growing in terms of user
community, and in terms of application diversity. Thinking in terms of
replacing the current satellites and of improving the system performance, a new
data collecting constellation is studied in this work. The study is basically
concentrated in the determination of the constellation orbits and number of
satellites, which could assure satisfactory coverage and time regularity
characteristic for a region which contains, at most, all the South America
territory, and, of course, the symmetric northern hemisphere region. The
obtained results are presented and extensively discussed.
Uwe Feucht, Christian Arbinger, Michael Kirschner
German Space Operations Center, Germany
Main drivers for the number and size of
orbit control maneuvers for LEO satellite formations are the payload
requirements and, as learned from GRACE, the impact of the attitude variation
due to the aerodynamic drag.
To fulfill these requirements the orbit
maneuver size and frequency is the main parameter for the compensation for the
aerodynamic perturbations. GRACE experience shows that differences from the
expected orbit decay in altitudes around 500 km in a short-term range are
mainly caused by attitude changes (for mid- and long-term planning also the
solar activity variations must be taken into account).
Laurent Francillout, Pascal Desmazeaux
CNES, France
In December 1998, the ESA council decided
to locate the Automated Transfer vehicle Control Center (ATV-CC) at
CNES/Toulouse. Although ATV is an Automated vehicle, it relies on ground
support in particular during deorbitation phase. So the Flight Dynamics
subsystem is in charge of the ATV re-entry and shall guarantee for each mission
configuration (various range of altitude, and phasing) a boundary impact zone
with an adequate level of confidence. The purpose of this paper is to present
the ATV scenario. The analysis has to estimate the minimum required DV budget
allocation to perform deorbitation in safe conditions and to evaluate the
dispersions around nominal trajectories.
Pierre LABOURDETTE
CNES, France
Andrei BARANOV
KIAM/KIAS, Russia
Several
missions perform rendezvous between near-circular and no coplanar orbits (Mars
Sample Return, Constellations). Maneuvers cost must be optimized taking into
account the natural drift of the orbit plane thanks to the second zonal
harmonic coefficient of the gravity field. Additional fuel-savings can be
achieved thanks to small inclination corrections that contribute to the orbit
plane drift. The optimization problem is difficult to be solved with general
non-linear optimization algorithm: non-convergence problem raises when tuning
at the same time nodal drift and in-plane phase angle correction. This is due
to very long duration dedicated to rendezvous (several hundreds of vehicle
revolutions separate two maneuver cycles). We will demonstrate, in the paper,
how a simple analytical approach copes with these problems. Tuning mainly the
semi-major axis by the mean of simple equations allows estimating the
intermediate drifting orbit in order to catch-up at the same time nodal and
in-plane phase angles. Examples will be described showing advantages of the
method.
Session 12
Orbit
Determination
10:20-12:00
Mina Ogawa, Masao Hirota, Kazuaki Nonaka, Yoshio Morooka, Shigehiro
Mori, Mikio Sawabe
NASDA, Japan
SELENE is a Japanese lunar explorer
scheduled to be launched in FY2005. RSAT mission of SELENE is to perform
satellite-to-satellite Doppler measurement over almost the entire moon surface.
RSAT team is going to improve the lunar gravity model using data of two months
or more. The results of preliminary analysis showed that the orbit
determination accuracy can be improved by the lunar gravity model estimated using
SELENE data of shorter period though the data has little effect on improving
the model. This paper presents the outline of RSAT mission, and the results of
the analysis.
Laurent Chausson
Communication & Systemes, France
Stephanie Delavault
CNES, France
During the interplanetary flight of a
probe, optical measurements of celestial bodies (such as asteroids) make
possible autonomous navigation. This cost-effective technique was tested in
1999 by Deep-Space-One. Asteroids are selected according to various criteria
such as magnitude, distance and star density. An additional criterion is the
spatial configuration that optimizes the orbit determination accuracy. The
resulting error ellipsoids are computed using the CNES covariance analysis tool
"Eperon", taking into account several error sources (data noise,
center finding, asteroids ephemeris, non-gravitational forces, etc), for the
Earth-to-Mars transfer of Mars-Premier-2007. This study is performed jointly by
CS and CNES.
Helio Koiti Kuga, Valcir Orlando,
INPE, Brazil
In a search for cheap, easy, and still
fairly reliable system for satellite control, an economical way of yielding
orbit information is to measure the Doppler shift suffered by the signal
transmitted by the satellite, commonly named one-way Doppler measurements. This
paper gives an analysis of such Doppler based orbit determination which will be
used in the French Brazilian Micro-satellite (FBM) satellite under development.
The ground segment consists basically of a control center and a single tracking
station located at Natal, Northeast of Brazil. Requeriments for orbit accuracy
coming from the scientific community is rather loose, so that the main
requirements are due to operations of tracking and scheduling of the control
center. Initially a covariance analysis is shown, which depicts the accuracy
achievable by the orbit determination based solely on one-way Doppler
measurements from a single tracking station. Afterwards, we use one-way Doppler
measurements taken from the SCD1 Brazilian satellite, a live flying satellite
with similar orbit pattern. These measurements presented problems typical of
the ones expected during the FBM mission. Orbit determinations are performed
using such set of data to show the errors with respect to the reference orbit.
At the end some conclusions and recommendations are drawn.
Maki Maeda
NASDA, Japan
NASDA have started study of precise orbit
determination since 1994. In 1996, we succeeded in the estimation of ADEOS
trajectory less than 1-meter accuracy as a research basis. After that, NASDA
started development of precise orbit determination system which treat on-board
GPS, ground GPS and SLR data. About the GPS satellite orbit, it could be
determined with less than 50cm accuracy using ground GPS data. After the launch
of ADEOS-II on Dec.14th, we started to determine orbit of ADEOS-II, using
on-board GPS (only L1) and SLR data. It will be determined with the less than 5
meters as a comparison result of GPS and SLR.
Yu.F.Kolyuka, T.I.Afanasieva, T.A.Gridchina
Mission Control Center, Russia
Abstract
Carrying out and processing of
measurements fulfilled with the help of instrument “Sage-3” equipment
established onboard the spacecraft "Meteor - 3М" which was started in
December 2002, provides maintenance enough exact knowledge of both the S/C
position and time at the moments of these measurements performance. For the solving
of a problem of high precision navigation binding Sage-3 experiment it was
planned to use the onboard GPS/Glonass equipment for satellite navigation.
However, this equipment established on S/C "Meteor - 3М", appeared to
be disabled.
The presented work is devoted to
revealing of special approaches and development of additional ways to achieve
the accuracy what would turn to be of good precision for "Sage-3"
experiment operation and for S/C "Meteor - 3М" orbit prediction in
conditions of absence of onboard GPS/Glonass equipment measurements, but only
on the base of the regular ground-tracking measuring means stipulated by the
given project for the S/C orbit control.
The results of the S/C orbit
determination, that were obtained with the help of the specified approaches and
ways are listed. These results allow to estimate real accuracy of S/C
"Meteor - 3М" orbit knowledge.
Session 13
Orbit
Determination
13:30-15:30
Hiroaki Umehara, Masaaki Takahashi, Kazuhiro Kimura
Kashima Space Research Center, Japan
Many orbits of the respective objects
should be observed and determined quickly and accurately for full utilization
of finite orbital resources such as the geostationary orbit. For the purpose,
optical observation has an important role because it is applicable for objects
with unknown frequencies transmitted. The U.S. and Russia have taken the
initiative in surveillance. In Japan, high-resolution telescopes were built in order
to detect small unknown objects. Much unknown debris will be discovered soon.
An international concern is the operating expense associated with the detection
and the orbit determination.
Our working Kashima Space Research Center
in CRL built an optical observation facility in 1998.
The telescope is an epsilon Newtonian
reflector.
Seiji Katagiri, Mina Ogawa, Mikio Sawabe, Masao Hirota, Ken Nakajima,
Manabu Hotta
NASDA, Japan
The Real Time Trajectory Estimation
Program (RTEP) is NASDA's system for estimating a satellite's trajectory and
thrust acceleration based on the Kalman filter during the apogee engine firing
(AEF). RTEP has been used experimentally during the AEF of, ETS-VI (1994) and
COMETS (1998). On September 2002, RTEP was used operationally for four AEFs of
DRTS (Data Relay Test Satellite). This paper presents the overview of the RTEP,
operation configuration, the detailed results of the operations concerning the
AEF phase of DRTS, and the future plan to improve the performance of the RTEP.
V.N. Zhukov, E.K. Melnikov
MCC, Russia
The main trends and tasks for flight
dynamics and ballistic and navigational support of ISS are covered. Also
represented are: the main tasks for the ISS pointing, being solved at MCC-M in
the course of maintaining the Station flight altitude strategy, specifics of
MCC-M to MCC-H interaction in the process of space debris avoidance tasks.
Session 14
Attitude
Dynamics, Estimation,Control
15:50-17:30
Maria Cecilia F. P. S. Zanardi
UNESP, Brazil,
Isaura Martinez Puentes Quirelli
UNESP, Brazil,
Helio Koiti Kuga
INPE, Brazil
An analytical approach for
spin-stabilized satellites attitude propagation is presented, considering the
influence of the residual magnetic torque and eddy currents torque. It is
assumed the inclined dipole model for the Earth?s magnetic field and the method
of averaging such torques, over each orbital period, is applied to obtain the
components of the torques in the satellite body frame reference system. The
inclusion of these torques on the rotational motion differential equations of
spin stabilized satellites yields the conditions to derive an analytical
solution. The solution shows that the eddy currents torques causes an
exponential decay of the angular velocity magnitude and the coupled effect of
both torques produces a precession on the spin axis. Numerical simulations
performed with data of the Brazilian satellites (SCD1 and SCD2) show the
agreement between the analytical solution and the actual satellite behaviour.
M.Yu.Beliaev, V.M.Stazhkov, Yu.A.Banit
RSC Energia, Russia,
N.I.Efimov, V.V.Sazonov
Keldysh Institute of Applied Mathematics RAS, Russia
We present the method and results of
estimating the inertia tensor of ISS on the base of the telemetry information
about its orientation and the total angular momentum of gyrodines. At first, we
reconstruct the station attitude motion. Then such a reconstruction is used in
the linear differential equations, which describe the variation of the angular
momentum of gyrodines during the motion and depend upon inertia tensor
components. We estimates these quantities from the condition of the best
approximation of angular momentum measurements by sulutions of those equations
using the least squares method.
Sanguk LEE, Kyoung Min ROH, Jae Hoon KIM, Seong Pal
LEE
Communications Satellite Development Center, Korea
KOrean Multi-Purpose
SATellite-1(KOMPSAT-1) which is remote sensing satellite had been launched in
December 1999 and has been being operated normally by Mission Control
Element(MCE), which was developed by Electronics and Telecommunications
Research Institute(ETRI). For LEOP or contingency operations, orbit
determination and attitude determination by using data from onboard
magnetometers and sun sensors are presented in this paper. Firstly, orbit
determination of KOMPSAT-1 using measured data by magnetometers and conical
earth sensors during normal operation phase is carried out and it is verified
by GPS data gathered from onboard. Secondly, attitude determination of
KOMPSAT-1 using measured data by magnetometers and GPS during normal operation
phase is carried out and the result is verified by attitude data from the
satellite. Finally, The orbit determination and the attitude determination are
carried out simultaneously using measured data by magnetometers and sun sensors
during sun pointing operation mode, which is very close to contingency
operation mode. The results are verified by GPS data.
M.Ovchinnikov, V.Pen'kov, I.Kiryushkin, R.Nemuchinsky,
A.Ilyin, N.Tretyakova
Keldysh Institute of Applied Mathematics of RAS,
Russia
Abstract
Attitude control systems which provide
small satellites with a required angular motion developing the control torque
through interaction with the geomagnetic field are considered. They can be
passive or active in regard with the purpose of a satellite. Relatively low
accuracy of orientation achieved by such systems can be ether accepted due to
the purpose of the satellite or can be compensated through preprocessing the
payload data. Next, the magnetic attitude control systems developed for the
latest years followed by results of fly testing are presented.
20 June 2003, Friday
Session 15
Celestial Mechanics
08:20-10:00
Alexei Golikov
KIAM, KIA Systems
Not
available
V.I. Denisov,
Institute of Applied and Theoretic Investigations of the Moscow State
University, Russia
A.V. Bagrov
Institute of Astronomy of the Russian Academy of Sciences, Russia
Now two similar astrometrical projects
are prepared in Russia and in USA. Both are based on pupil interferometry that
allows to provide space observations with accuracy up to 10-5-10-6 arcseconds.
This accuracy is so high that some gravitational effects, that not long time ago
were subjects for experimental investigations for testing of basic nature lows,
can put limits to the real accuracy of astrometrical catalogues and inertial
co-ordinate system.
This is why we need to understand well
all the effects that can distort positions of celestial objects and to consider
them at processing of observational data.
Sergey Kudryavtsev
Sternberg Astronomical Institute of Moscow State University, Russia
Accurate expansion of the perturbation
function due to attraction of the "third-bodies" (Sun, Moon and
planets) to a Poisson series is done. The series is built through an improved
harmonic analysis of the ephemerides DE/LE-406 over 3000BC - 3000AD. Unlike the
results of classical Fourier analysis, both the amplitudes and frequencies of
the obtained series are high-degree polynomials of time. The new expansion is
applied to analytical prediction of both high- and low-altitude Earth's
satellites motion. The work is supported in part by a grant number 02-02-16887
from the Russian Foundation for Basic Research.
Victoria I. Prokhorenko
Space Research Institute Russian Academy of Sciences, Russia
We turn our attention to the problem of
the long-live AES orbit selection. The geometrical method is developed for the
parametric analysis of the satellite orbits long-term evolution under action of
the external bodies gravitational perturbations.
As a conceptual basis we use the
analytical solution of the double-averaged circular Hill's problem obtained by
Lidov (1961). We use, as an example, the high-apogee orbits of the AES ROGNOZ -
series launched in 1972 to 1995. To apply the above-mentioned analytical
solution to the problem of the four bodies (Earth, AES, Moon, and Sun) we use
the assumption of the co-planarity of the Moon orbit and ecliptic planes. We
compare the analytical estimation with the results of the numerical integration
of the complete system of the differential equations taking into account the
lunisolar gravitation perturbations.
Valentine Prokudina, Nadezhda Kolodiazhnaya, Darya Petukhova, Darya
Mikhaylova
Moscow City Palace for Children (Youth) Creativity, Russia
It is known that during solar activity it
is observed the disturbance of the Earth magnetosphere, leading to the break of
normal work of satellite and to the changes of its orbits.
We estimated the change of density at
high atmosphere during the period of largest solar and geomagnetic activity at
July 2000, using the model of Jacchia with parameters F 10.7 and Ap -indicies.
In addition to geomagnetic indicies we
analysed the Belt Index which increased on several orders during the period of
high level of magnetospheric activity because of the precipitation of energetic
particles to atmosphere.
We investigated the periods, when several
satellites worked with considerable breaks (for example May, 1998). Besides we
analyzed the solar and geophysical data during the sharp descent of the KA MIR.
Session 16
LOW THRUST
TRAJECTORIES
10:20-12:00
Jesus Cegarra,
GMV at ESA/ESOC, Germany
Gottlob Gienger,
European Space Operations Centre, European Space Agency, Germany
The Cluster mission is a set of four spin
stabilized spacecrafts forming a tetrahedron. Since their launch in July and
August 2000 three constellation changes have been prepared and commanded.
A precise modelling, covering models for
the thermodynamical behaviour of the helium and propellant, combined with a
calibration method based on covariance analysis, has given very good manoeuvre
performance. A new method is presented to calibrate a set of two axial Delta-V
manoeuvres and one attitude slew with up to four different thrusters
simultaneously. Implementation, results and benefits for the Cluster
constellation manoeuvres are described in this paper.
R.Z.Akhmetshin , T.M.Eneev
KIAM, Russia
The spacecraft which is now worked out
for the "Fobos-Ground" project, cannot deliver the payload to
asteroid of the Main Belt, necessary for relic ground sample return. Mission
becomes possible if
the assist of Mars will be used (about
twenty variants of missions were found for asteroids of all the main types
C,E,I,M,O,S,U), or
the power of solar arrays will be
increased (twice increasing allow to realize straight flights to many
asteroids).
Supported by RFBR, grant 01-01-00015.
Vyacheslav Petukhov
MAI, Russia
It is considered spacecraft insertion
into target orbits using light-class launch vehicles and electric propulsion.
Typical flight profile includes insertion into a parking orbit using launch
vehicle, transfer into an elliptical intermediate orbit using high-thrust upper
stage, and transfer into target orbit using electric propulsion. The
optimization of low-thrust transfer and intermediate orbit parameters is
carried out. Available launch vehicles, upper stages, and electric propulsion
are considered. There are presented dependencies of delivered mass with respect
to transfer duration, spacecraft electrical power, and electric propulsion
specific impulse. Presented results could be used for feasibility study of
low-cost space missions.
Viacheslav V. Ivashkin, Alexander V. Chernov
KIAM, Russia
Optimal trajectories for a space flight
to a near-Earth asteroid (NEA) are presented. The flight aim is an
impact-kinetic effect of the spacecraft (SC) upon the NEA, a correction of the
NEA orbit and its deflection from the Earth. This deflection is maximized under
the optimization. The SC has a combination of a chemical jet propulsion system
and an electric one. Numerical results are given for the asteroid Toutatis
orbit.