European Space Operations Centre
Rosetta.
Quick Mission re-Design of Europe’s comet chaser
ATA, Barcelona, July, 2004
J. Rodriguez-Canabal, ESA, OPS-GA
Page: 1
Contents



Rosetta, Comets, and Space
Missions
Rosetta Original Mission.
Spacecraft and Payload
Re-design of New Mission






Launch with Ariane 5
Gravity Assists. Optimization and
models.
Trajectory description. Navigation.
Fly-by of Lutetia and Steins.
Approaching 67P/ChuryumovGerasimenko
Landing of Philae
Page: 2
Rosetta ESA-Cornerstone


In November 1993, ESA’s approved Rosetta
as a cornerstone mission in ESA’s Horizon
2000 Science Programme.
Rosetta will be the first mission :






To orbit a comet nucleus.
To fly alongside a comet as it heads closer to the
Sun.
To observe from very close proximity how the
frozen comet nucleus is transformed by the heat of
the Sun.
To send a Lander for controlled touchdown on the
comet nucleus surface.
To obtain images from a comet’s surface and to
perform in-situ analysis
To fly near Jupiter’s orbit using solar cells asPage: 3
In situ measurements
Page: 4
Why the name Rosetta?



The Rosetta stone
(1799) was the key
to deciphering the
old hieroglyphics
writing of ancient
Egypt.
Decree to honour
Ptolemy V (210-180
BC)
Obelisk from Island
of Philae (1815)
Page: 5
Why to go to a comet?


Comets have always attracted the attention of
mankind. The apparitions are recorded in
documents going back millennia.
Comets appear suddenly and have been
interpreted as good signs or as bad omens
announcing great disgraces.
Battle of Hastings (1066 AD)
Page: 6
Why to go to a comet? (2)

Are comet dangerous for us?.
What happens if a comet hit the Earth?.
Dinosaurs extinction event Chicxulub impact
crater in Yucatan (discovered 1991).
We cannot do too much about it !
Meteor
Page: 7
Why to go to a comet? (3)

A comet is a celestial body originating very far
away from the Sun



Oort cloud, far beyond Pluto (50000 AU)
Kuiper Belt,
beyondofNeptune ( 30-100 AU)
nucleus
composed
ice, dust, of a size
between a few
hundred m up to a
few km. Carbon
compounds.
Near the Sun it
develops a coma (
100000 km), and
tails (dust, ion)
several Mkm
Page: 8
Why to go to a comet? (4)



Scientist wants to study comets because
these are what is left of the “primitive cloud”.
They are time capsules preserving the
physical and chemical conditions that existed
when the planets were formed 4.5 billions of
years ago.
Comets could have provided water and
organic material to the Earth.
Comets can help to understand
conditions of formation of the
solar system
Page: 9
Space Missions to Comets

To Halley



Giotto, 1986, 600
km,
68 km/s and comet
Grigg-Skjellerup,
1992, 200 km.
(ESA)
VEGA-1 & VEGA-2,
9000 km, 78
km/s1986. (RUS)
Sakigake & Suisei, 7
Giott
Mkm, 150000
o
km,1986.VEG
(JAP)
A
Page: 10
Space Missions to Comets (2)

Halley nucleus was full of surprises (size,
albedo 0.03, jet activity)
Giott
o
Page: 11
Space Missions to Comets (3)



ISEE-C/ICE to comet Giacobini-Zinner, 1985,
NASA, 8000 km
Deep Space, 2001, comet Borrelli
Star Dust comet Wild-2, 2004, 240 km, 2.6
AU
Page: 12
Rosetta
Ready for Launch Jan 2003





Launch Jan. 2003 with Ariane 5 G+ using
EPS delay ignition.
Use of 3 Gravity Assists (Mars-Earth-Earth).
Fly-by of 2 asteroids: Siwa and Otawara.
Large distance from Sun, 5.3 AU, and from
Earth for long periods.
Arrival at Wirtanen on Dec. 2011. Orbiting
around the comet nucleus for 1.5 years (up to
perihelion)
Fully optimised for the mission to Wirtanen
fixed:


Max. min. distances to Sun. (0.9 AU – 5.3 AU)
Page: 13
Propellant (660 kg of MMH. 1030 kg of NTO)
Spacecraft






Wet launch mass
3064 kg
Solar power (300 W-8
kW)
24 x 10 N bipropellant
thrusters
2 Navigation cameras,
2 Star trackers,
4 Sun sensors,
9 Laser gyroscopes,
9 accelerometers
HGA of 2.2 m, MGA,
LGA, S-X band
Data storage 20Gbits.
Page: 14
Scientific Payload

Remote Sensing





OSIRIS (Optical, Spectroscopic and Infrared Remote
Imaging System)
Wide and Narrow angle camera.
ALICE (UV spectrometer) Analyses gases in the coma and
tail. Production rates of water and CO and CO2. Comet
surface.
VIRTIS (Visible and IR Thermal Imaging Spectrometer).
Maps solids and temperature of comet surface.
MIRO (microwave Instrument). Abundance of major gases,
surface outgassing rate, nucleus subsurface temperature.
Composition Analysis



ROSINA (RO Spectrometer for Ion and Neutral Analysis)
Composition of atmosphere and ionosphere, velocities of
charged particles, and reaction between them.
COSIMA (Cometary Secondary Ion Mass Analyser). Dust
grains characteristics
MIDAS (Micro-Imaging Dust Analysis System) Dust Page: 15
environment; grain morphology
Scientific Payload (2)

Nucleus large structure


Dust flux, mass distribution


GIADA (Grain Impact Analyser and Dust Accumulator).
Number, mass, momentum and velocity distribution of dust
grains.
Plasma environment


CONSERT (Comet Nucleus Sounding Experiment by
Radiowave Transmission). Nucleus tomography
RPC (Rosetta Plasma Consortium). 5 sensors measure the
physical properties of the nucleus, structure of the inner
coma, cometary activity, interaction with solar wind.
Radio science

RSI (Radio Science Investigation). S-X band, measure
mass, density of nucleus. Solar corona during conjunction
events.
Page: 16
Spacecraft
OSIRIS
VIRTIS
COSIMA
MIDAS
ALICE
CONSERT
MIRO
ROSINA
GIADA
RPC
Page: 17
Scientific Payload (3)

Rosetta Lander








CONSERT
ROMAP (RO Lander Magnetometer and Plasma Monitor).
Local magnetic field and comet/solar wind interaction.
MUPUS (Multi-Purpose Sensors for Surface and Subsurface
Science). Sensors to measure density, thermal and
mechanical properties of surface.
SESAME (Surface Electrical, Seismic and Acoustic
Monitoring Experiment). Electric, seismic and acoustic
monitoring. Dust impact monitoring.
APXS (Alpha, Proton, X-ray Spectrometer). Elemental
composition of surface.
ÇIVA/ROLIS (visible & IR imaging). 6 cameras and
spectrometer. Composition, texture, albedo of samples from
the surface.
COSAC (Cometary Sampling and Composition). Gas
analyser for complex organic molecules
Modulus Ptolemy. Gas chromatography; isotopic ratios of
Page: 18
light elements.
Rosetta Recovery

Failure of Ariane Flight 157 on 11.12.2002 led
to intense work to study alternative scenarios
in case of cancellation of Rosetta launch on
Flight 158.





Fixed constraints on spacecraft: mass, propellant,
power, thermal, mechanical, Telemetry
Use of periodically up-dated database of extended
alternative mission.
Very good collaboration of ESA, Industry, and
Scientists
January,7, 2003, launch of original Rosetta
cancelled
Recommendation of first ESA internal review
27.01.2003:

Page: 19
No Venus swing-by; Maintain mission schedule;
Rosetta Recovery (2)

25-26 Feb. 2003 ESA’s Science Programme
Committee



Intense activity on:





67P/Churyumov-Gerasimenko; launcher Ariane 5;
launch Feb. 2004 with launch backup in 2005
using Proton.
46P/Wirtanen; launcher Proton; launch Jan. 2004.
Observation of 67P/Ch-G using HST, and ESO
Lander constraints. Rebound on 46P/Wirtanen,
crash on 67P/Ch-G
Spacecraft constraints. Unloading of MMH, but not
of NTO. Danger of tanks corrosion
Launcher performances: payload, fairing
dimensions
13-14 May, SPC decided 67P/Ch-G withPage: 20
Rosetta Recovery (3)

Missions considered for recovery
Launc RV
Vinf De DV
Perih.
km/ de m/s
min/M
dist.
Aster
Baselin
Siwa-
2003/
2012/
2013/
3.4
0
149
0.93/5. Siwa
Chur-
2004/
2014/
2015/
3.5
-3
159
179
0.88/5.
Wirtane
2004/
2012/
2013/
5.0 10.
112
175
Lutet
Rhod
Siwa
0.9/5.3 Isis
Julia
Page: 21
Ariane 5 EPS
Delayed Ignition



The engine of the
upper stage, EPS, of
Ariane 5 is ignited
after cut-off of the
central core engine,
but it can be restarted or its ignition
delayed.
A delayed ignition
increases the time
from launch to
injection, but
substantially
increases the
performance
Flight software for
Page: 22
The Big Jump
Page: 23
AR 5 Delayed EPS ignition


Only 2 Launcher Flight Programs needed for
a launch period of 21 days (26.02 –
17.03.2004) with 2 launch attempts per day.
Original mission had 14 FP.
Earth escape targets: V = 3.545 km/s,  = 2°
Page: 24
AR 5 Delayed EPS ignition
Page: 25
Gravity Assists



Gravity Assists have been used since 1973
Mariner 10 mission, that flew by Venus in its
way to Mercury.
Later Pioneer 11 to Saturn, Voyager 1 & 2
(Jupiter, Saturn, Uranus, Neptune), Galileo to
Jupiter, Ulysses out of the ecliptic, Vega, ICE
to comet Giacobini-Zinner, Giotto, etc.
Gravity Assist or swing-by is a significant
trajectory perturbation due to a close
approach to a celestial body.
Foundations laid down since early 20th
century. Applications to missions described by
1965.
Gravity assist is based on the deflection of
the arrival relative velocity, Va, to the Page: 26
departure relative velocity V d,
Gravity Assists (2)
VPlanet

Vrd
Vd
VPlanet
VPlanet
Vra
Va
Swing-by
Va
VPlanet
Vd
V EGA
Page: 27
Gravity Assists (3)


The change of velocity is Vd = Va + (Vd - Va
).
The deflection angle is given by:
sin/2 = 1 / (1+r V2 /)
The change of velocity is:
2)
v
=
2
V
sin/2
=
2
V
/(+
r
V




Planet
/2
v (km/s)
Venus
61.6
4.77
Mars
39.9
3.40
Jupiter
85.8
11.25
r = planet radius, Va = Hohmann transfer
Page: 28
Gravity Assists (4)


The VEGA (V-Earth Gravity Assist) is the
use of a swingby of the Earth after a V
manoeuvre. (Hollenbeck 1975).
Launch from Earth into a 2 or 3 years
heliocentric trajectory
(V < 5 km/s), followed by a manoeuvre near
aphelion (few hundred meters) to target either
before or after perihelion produces a relative
V ~ 10 km/s.
Page: 29
Finding the good way there


Comet of interest: perihelion  1 AU, Aphelion
 5-6 AU
Departure from Earth or last Earth swing-by
with relative velocity of 9-11 km/s. Gravity
Assists is needed




Delta-V + Earth GA high propellant consumption
(3 years round trip, with launch at V ~3.4 km/s,
900 m/s needed to reach the 9 km/s)
Mars GA + Earth GA: - launch at 3.5 km/s, one
revolution before Mars, or at3.9 km/s, one
revolution between Mars and Earth return.
Venus GA: thermal problems with the spacecraft
are confirmed.
The strategy Launch – Earth within one Page:
year
30
COMET RENDEZ-VOUS STRATEGIES
01/2003: Mars GA (A window)
Page: 31
Finding solutions

Sequential approach:




Feasible missions
Optimization using simple models
Full numerical optimization with all mission constraints
Given a sequence of swing-by, and the number of
revolutions between swing-by, a discrete search
provides the swing-by times.



Techniques to accelerate the search: keep tables of Lambert
solutions, prune trajectories, order results.
Pay attention to:
- Number of revolutions in between swing-by, and cases;
- singular cases: multiple swing-by of same body at 180°
or 360°
Using a constrained non-linear parameter optimisation
method, optimise sequence of events, launch conditions,
and introduce Deep Space Manoeuvres to force to zero any
manoeuvres at swing-by.
Page: 32
Finding solutions




Parameter optimisation min F(x), xEn, with
qi(x) = 0, gi(x) > 0.
To ensure convergence, it is important to
make a good selection of the variables, the
constraints, and the cost function.
The cost function is typically the useful mass,
or the sum of the modulus of the V with
weighting factors.
The variables can be position and velocity
vectors at some points in the trajectory, dates,
impact vectors, angles, orbital elements, etc.
The constraints describe the initial/final
conditions, trajectory matching at selected
points, minimum swing-by height, technical
constraints to control behaviour of the Page: 33
Optimization



Problem is defined as:
min F(x), xEn, gk(x)=0, k=1,…q, gk(x)>0,
k=q+1,…,m
Sequential Quadratic Programming is a
generalized Newton’s method that, starting in
a given point, finds a better point by
minimizing a quadratic model of the problem.
Packages: OPTIMA, MATLAB, NPSOL,
NLPQL, SQP
OPTIMA: penalty function P(x,r)=F(x)+ g(x)T
g(x)/r.
Quadratic sub-problem:
g
min ½ pT B p + fTp, with Ap=-½ r  Page: 34
Selection of model

Rosetta – 67P. Patched conics. No asteroids.
Free Launch date and comet rendezvous
date.
E2
DSM3
TL
TE3
TE2
TDSM3
Variable
s 18
L DSM1 E1 DSM2 M
E3 DSM4
Comet
TDSM1
TE1
TDSM4
Constrai
nts 7
Ea1
Ea1
TDSM2
Tc
TM
Ma Ed2
Ma Ed2
RDSM4 < 4.4 AU (Solar Power)
RpE1 , RpE2 , RpE3 > RminE , RpM > RminM;
=0.
TC > TC min, ( RC < RC min )
Ed3
Ed3
Vswing-by
Page: 35
Selection of model (2)










Arc M-E2, Lambert  VdM , VaM , VE2a , VE2d
Arc DSM2-M, back propagation  RDSM2
Arc E1-DSM2, Lambert  VDSM2 , VE1d , VE1a
Arc DSM1-E1, back propagation  RDSM1
Arc L-DSM1, Lambert  VL , VDSM1
Arc E2-DSM3, forwards propagation  RDSM3
Arc DSM3-E3, Lambert  VDSM3 , VE3a , VE3d
Arc E3-DSM4, forwards propagation  RDSM4
Arc DSM4-C, Lambert  VDSM4 , VRDV
VL  ML   Vi /ISp  MRDV
Page: 36
AR 5 Delayed EPS ignition
Estimated performances


Page: 37
Selection of model (3)


Similarly can be solved the Launch Window
problem where the fixed parameters are TL ,
TC , VL , L .
18 variables, 5 constraints.
Page: 38
The acrobatics




Launch-Earth-MarsEarth-Earth-Comet
L-E1, 370 d, 170 m/s
E1-M, 730 d
M-E2, 260 d
E2-E3, 727 d
E3-DSM,
540 m/s
DSM-67P, 1110 d
Near comet, 445 d
7160 M km !!
Earth 940 Mkm/year
Page: 39
Trajectory
Trajectory Earth-Earth
 Manoeuvre Optimisation
1.5
DSM1.2

1.0

EARTH
1
Y (AU)
0.5

DSM1.1 – Perihelion (6/2004)
DSM2.1 – Aphelion (12/2004)
Variation with launch day
LAUNCH
0.0
-0.5
DSM1.1
-1.0
-1.5
-1.5
-1.0
-0.5
0.0
0.5
1.0
1.5
X (AU)
Page: 40
Events
Event
V (m/s)
Deep Space Manoeuvre 3
Date
2004/02/26
2004/05/25
2005/03/02
2006/10/21
2007/02/27
2007/11/15
2009/03/16
Earth gravity assist
2009/11/11
Deep Space manoeuvre 4. @ (4.4 AU)
2011/05/10
533
Rendezvous with 67P. @ (4.0 AU)
Start nucleus operations at 3.25 AU
Perihelion pass
2014/05/23
2014/10/03
774
Launch
Deep Space Manoeuvre 1.1
Earth gravity assist
Deep Space Manoeuvre 2
Mars gravity assist
Earth gravity assist
Vrel. (km/s)
3.545
Peric. H (km)
173
3.9
4300
8.77
9.33
200
14000
9.98
550
65
130
2015/08/11
Page: 41
Distances to Earth & Sun
Page: 42
Rosetta got an extra

The propellant left for Near Comet
operations, after rendezvous with 67P, varies
by 20 kg, (33 % of allocation at comet).
After a delay of 5 days, Rosetta was launched
on March, 2.
Page: 43
Planet swing-by


Conditions at the
first Earth swing-by
depends on the day
of launch
Conditions at Mars
swing-by or at
subsequent Earth
swing-by are very
fixed
Earth
-1
Mars
Page: 44
Planet swing-by (2)
Earth -2
Earth
-3
Page: 45
Navigation


Orbit Determination and Trajectory Correction
Manoeuvres
Measurements:





Distance measurement (radio tracking range) (2-5
m error)
Relative velocity spacecraft – Ground Station
(Doppler) (1 mm/s error)
Delta-DOR (Differential one-way ranging) ( ~ 20
cm error)
Onboard Optical Measurements (Camera, star
trackers)
Delta-DOR measurements use spacecraft
signal simultaneously received by 2 ground
stations. It is a type of Very Long Baseline
Page: 46
Interferometry measurement and determines,
Navigation (2)



By using the signal from a nearby quasar,
both GS cancel the common error sources
(atmosphere, propagation media, clocks)
Delta-DOR measurements are very useful in
critical phases of a mission: planet approach,
prior to a swing-by, orbit insertion, landing,
etc.
Other sources of errors are:



Station position ( < 1 m )
Signal propagation (troposphere, ionosphere,
spacecraft transponder)
Modelling of forces (planets, solar radiation
pressure, out-gassing, open thrusters, ..)
Page: 47
Navigation (3)



Effect of biases. Measurement equations:
z=Ax+By+
where: z : measurements residuals
(observed – computed).
x : variables to be estimated.
y : variables known to be biases and
not estimated.
Estimated:
xe = (AT W A) -1 (AT W) z
W-1 = E (  T)
Computed error covariance:
P = E (xe – x, (xe – x)T = (AT W A+C1)
–1

Consider covariance:
Page:, 48
Pc = P + S Py ST , S = -P AT W B
Py
T
Correcting the Launcher

Launcher injection errors corrected by
manoeuvres that re-optimises the full
trajectory.
Large correction manoeuvre may be needed.
Difficult first acquisition from Kourou
Page: 49
Interplanetary Navigation

COVARIANCE ANALYSIS – Knowledge and
Dispersion Matrixes
 Deterministic
Manoeuvres
- Implementation
Errors
- No re-optimisation
- Degradation of
Knowledge

Dispersion and Knowledge mapped at pericentre of
1st Earth swing-by
Mid-Course
Corrections
- Improve dispersion
errors at target
- Implementation
Page: 50
Interplanetary Navigation

Mars swing-by is critical. Minimum altitude
selected at 250 km.
Very good experience with Mars Express
Page: 51
Interplanetary Navigation (2)

Last Earth swing-by should be as low as
possible, baseline 530 km, but not critical
Page: 52
Interplanetary Navigation

Propellant Assessment



Ariane 5 Launching
Accuracy
- Position – 39 Km,
mostly AlongTrack
- Velocity – 36 m/s,
mostly radial
LIC - Launcher Injection
Correction
- 4 days after injection
Mid-Course Corrections
- About 17 targetting
conditions at
pericentre of planets
swing-by
ROSETTA Churyumov-Gerasimenko 2004
INTERPLANETARY NAVIGATION BUDGET ESTIMATION
TC #
Mean
(m/s)
1-s
(m/s)
Min
(m/s)
99%
(m/s)
Max
(m/s)
LIC
1
45.2
30.5
0.7
138.0
218.6
Phase Launch-Earth 1
4
3.2
1.3
0.08
8.9
13.9
Phase Earth1-Mars
4
1.4
0.6
0.04
3.7
5.6
Phase Mars-Earth2
2
3.9
2.0
0.07
10.6
16.6
Phase Earth2-Earth3
5
1.6
0.8
0.04
4.5
7.0
Phase Earth3-Comet
2
1.7
0.8
0.06
4.5
6.9
Total without LIC
17
11.8
2.7
0.3
32.1
50.1
Page: 53
Asteroids

The asteroids to be explored were decided
after launch.
The excellent performance of Ariane 5, error
in V< 1.8 m/s, and the optimal launch day
allow to include 2 asteroids fly-by along the
mission: Diameter
Period
Type
21
130x104x7
8.17
M, Mo,
2867
17.5 – 5.5
?
C?
Page: 54
67P/Churyumov-Gerasimenko





Comet discovered in 1969 by K. Churyumov
and S. Gerasimenko.
Up to 1840 the perihelion was 4 AU. A Jupiter
encounter reduced it to 3 AU. In 1959 another
Jupiter encounter reduced it to its current
1.28 AU.
Orbital period is about 6.6 years.
Well observed in 1976, 1982, 1989, and
2002.
Estimated diameter of nucleus 5 x 3 km.
Relatively active object. Dust production 60 –
220 kg/s.
Ratio gas / dust ~ 2.
Page: 55
67P/Churyumov-Gerasimenko

Comet models: 2 km radius, 12 hr rotational
period
Page: 56
Approaching 67P






Start of comet rendezvous when distance to
Sun = 4 AU
Operations based on using only the NAVCAM
for comet detection.
Earliest start at (3 Mkm)
As an improvement OSIRIS could be used for
comet detection.
Early comet detection can be used to
advance the Orbit Insertion Point (OIP). Start
of comet Global Mapping Phase.
Power will not drive the earliest start of RV
operations.
Power limit is at 4.4 AU.
Page: 57
Earliest start of phase is driven by available
Approaching 67P (2)

Near comet operation phases up to Lander
delivery do not depend on the comet
characteristics.
Page: 58
Approaching 67P (3)


The approach from 600000 km to 40 km, and
the reduction of the relative velocity from 780
m/s to 0.3 m/s will be performed in about 3
months.
During this period Rosetta will:




get comet images to determine: nucleus size,
shape, rotation, relative position+velocity,
identification of landmarks;
avoid cometary debris, and eclipses;
Keep Earth communications;
Keep safety;
Page: 59
Mapping 67P

Mapping and selection of landing sites:
- Orbit safety
- avoid debris, jets
- ensure no eclipse, no occultation
- cover at least 80% of illuminated
surface, good illumination conditions,
- volume of data to be transmitted
to Earth
- fly over 5 selected areas at
required illumination conditions,
and resolution.
Page: 60
Mapping 67P (2)
Page: 61
Mapping 67P (3)
Page: 62
Philae Landing
Page: 63
Philae Landing (2)

Montecarlo simulations by MPIAe (M.
Hilchenbach, Cologne 2003)
Montecarlo
calculation for
target comets:
Variation of radii
and densities*
no landing
possible
landing
possible
*still assuming landing on inactive
comet, about 3 AU away from the sun.
Page: 64
Philae Landing (3)

Lander Delivery, 12 d:
- arrive at delivery point in a safe orbit, with
the proper attitude
and velocity.
- Constraints on: Ground visibility, Eclipse,
Solar Aspect
- Mechanical Separation System constraints:
ejection V.
- Active Descent System constraints: V
vertical
- SSP landing constraints: V impact, angles,
landing errors
Page: 65
Philae Landing (4)

Delivery Trajectory and Landing errors (3-s). Vimpact < 120 cm/s
Page: 66
Philae Landing (5)
Page: 67
Conclusions

Thanks to a very intensive collaboration
between all people and institutions involved in
Rosetta a new mission to 67P/ChuryumovGerasimenko has been defined in a very
short time.
THANKS
Page: 68
FOR YOUR ATTENTION