The present invention relates to launching
5 satellites and putting them into orbit, and more
particularly it relates to deploying a constellation of
satellites.
STATE OF THE ART
10 Satellite constellations are used for numerous
applications requiring broad and continuous coverage.
Such satellite constellations comprise a set of
satellites describing distinct orbits around the Earth.
Nevertheless, deploying a plurality of satellites on
15 distinct orbits, which are commonly angularly offset
relative to one another, requires the use of a plurality
of launchers, which is very constraining.
By way of example, a constellation of observation
satellites in low orbit, with one observation per hour,
20 requires 12 satellites to be deployed on different
orbits. Unfortunately, deploying such a satellite
constellation presently requires 12 distinct launches,
which is very expensive, or requires a very large
quantity of on-board propellant in order to be able to
25 modify the or~bits of the satellites after they have been
deployed, which is likewise very problematic in terms of
on-board mass.
Solutions have been proposed for deploying a
plurality of satellites using a single launcher.
30 Nevertheless, those solutions rely essentially on the
ability of the launcher to reach the successive orbits in
order to deploy the various satellites on them, or
requires the satellites themselves to have the ability to
modify their own orbits once they have been deployed, and
' Translation of the title as established ex officio.
that remains problematic i.n terms of on-board mass, given
the quantity of propellant that is required.
The present invention thus seeks to propose a
solution to that problem.
5
S Y OF THE INVENTION
To this end, the present invention proposes a method
of dep10y;~ng a constellation of satellites, the method
comprising the following steps:
10 using a single launcher to deploy a plurality of
satellites at the same initial altitude on the same
initial orbit;
controlling said satellites in such a manner that,
depending on the initial orbit, each satellite reaches a
15 drift altitude selected from a drift set, with the orbits
of the various satellites shifting relative to one
another at the respective drift altitudes under the
effect of the gravitational potential of the Earth; and
controlling the satellites in such a manner as to
20 be moved sequentially in order to reach the same final
altitude, said sequential movement being performed in
such a manner that the satellites describe final orbits
that are angularly offset from one another, i.e. having
identical trajectories with the same angle of inclination
25 relative to Che equatorial plane, but presenting distinct
longitudes for their ascending nodes.
In a particular implementation, the final orbits of
the satellites are angularly offset relative to one
another about the Earth's axis of rotation.
30 Said final orbits then typically present a constant
angular offset between two successive final orbits.
By way of example, the drift set comprises a high
drift altitude and a low drift altitude, having
respectively an altitude that is higher and an altitude
35 that is lower than the initial altitude.
By way of example, the starting altitude, the high
drift altitude, the low drift altitude, and the final
altitude then lie in the range 150 kilometers (km) to
75,000 km.
The final altitude of said satellites then typically
lies between the initial altitude and the low drift
5 altitude.
By way of example, the final altitude lies in the
range 200 km to 800 km.
By why of example, the starting altitude is 800 km,
the high drift altitude is 1500 km, the low altitude is
10 270 km, and the final altitude is 420 km.
In another variant, the starting altitude lies in
the range 33,000 km to 38,000 km, and the final altitude
lies in the range 20,000 km to 25,000 km.
The kina1 orbits of said satellites typically have
15 angles of inclination that are different from their
initial orbits.
The present invention thus makes it possible to
deploy some or all of the satellites of a constellation
in a single launch, making advantageous use of the
20 gravitational potential of the Earth for the purpose of
modifying the orbits of some or all of the satellites as
deployed in this way.
S-Y OF THE FIGURES
2 5 Other characteristics, objects, and advantdges of
the invention appear from the following description,
which is purely illustrative and non-limiting, and which
should be read with reference to the accompanying
drawings, in which:
30 Figures 1 and 2 are diagrams showing the orbit
parameters of a satellite;
Figure 3 is a diagram showing the steps of the
method in an aspect of the invention;
Figure 4 is a diagram showing an example of the
35 method in an aspect of the invention for deploying a
satellite constellation;
Figure 5 shows an example of how the drift of the
various orbits of the satellites varies in the example
shown in Figure 4; and^
Figure 6 shows an example of a satellite
5 constellation deployed by means of a method in an aspect
of the invention.
In the figures, elements that are common are
identifie3 by identical numerical references.
10 DETAILED DESCRIPTION
Figures 1 and 2 are diagrams showing the parameters
of the orbit of a satellite.
In these figures, the Earth A is represented by a
sphere, b) the equator E, and by the axis P passing
15 through the poles of the Earth.
These figures also define:
an equatorial plane PE containing the equator E;
the vernal point g defined as being the
intersection between the ecliptic and the celestial
20 equator; a vernal point direction is also defined as
connecting the vernal point and the center of the Earth.
The orbit O of a satellite around the Earth is shown
diagrammatically in Figures 1 and 2.
In these figures, there can be seen:
2 5 the angle of inclination - i of the orbit plane PO,
i.e. the plane that contains the orbit, relative to the
equatorial plane PE. An angle of inclination of zero
means that the orbit plane PO coincides with the
equatorial plane PE;
30 the travel direction of the satellite, identified
in arbitrary manner by an arrow;
the ascending node NA corresponding to the point
of intersection between the orbit and the equatorial
plane PE when the satellite goes from the southern
35 hemisphere to the northern hemisphere;
the descending node ND corresponding to the point
of intersection between the orbit and the equatorial
plane Pi? when the satell-ite goes from the southern
hemisphere to the northern hemisphere; and
the apogee B and the perigee T, corresponding for
an elliptical orbit respectively with the point having
5 the highest altitude and the point having the lowest
altitude. The angle of inclination of the perigee
relative to the line of the nodes is measured by the
argument of the perigee. In the embodiment shown, the
orbit 0 is circular, and the apogee T and the perigee B
10 are substantially at the same altitude, with w = 90".
The longitude 0 of the ascending node is also
ich longitude is the angle between the
f the vernal point -g and the line connecting
ing node ND to the ascending node NA. The
15 longitude !2 of the ascending node is measured going from
the direction of the vernal point -g towards the ascending
node NA, corresponding to the direction of rotation of
the Earth, likewise indicated by an arrow in Figure 1.
It is thus possible to define an orbit plane by
20 means of the angle of inclination - i and the longitude 0
of the ascending node.
The satellite then describes a circular or
elliptical orbit in the orbit plane PO as defined in this
way.
2 5 For a circular orbit, the altitude is then
substantially constant.
For an elliptical orbit, altitude varies between a
maximum value when the satellite is at the apogee of its
trajectory, and a minimum value when the satellite is at
30 the perigee of its trajectory. For such an elliptical
orbit, the term "altitude" is used to designate an
altitude at a given point of the orbit, e.g. the altitude
at the ascending node, at the descending node, at the
perigee, or at the apogee.
3 5 Figure 2 shows two circular orbits 01 and 02 having
the same angle of inclination - i and having distinct
ascending node longitudes, thereby leading to an offset
dC2 between the orbits 01 and 02.
Figure 3 is a diagram showing the steps of the
method in an aspect of the invention.
Figures 4 and 5 show an implementation of the method
of an aspect of the invention. Figure 4 shows the
variation of altitude for the satellites of a satellite
constellation deployed by means of a method in an aspect
of the invention. Figure 5 shows the variation of the
longitude 0 of the ascending node of each of these
satellites as a function of tlme.
In a flrst deployment step 10, a plurallty of
satellites of a constellation are deployed by means of a
single launcher.
This plurallty of satellites may correspond to all
of the satellites of a constellation or to only some of
them. In the example described below and shown in
Figures 4 and 5, consideration is given to deploying a
set of six satellites in a single launch.
The plurality of satellites are deployed at a common
initial altitude in a common initial orbit.
Figures 4 and 5 show this deployment step at t=O;
all six satellites deployed during the single launch have
the same altitude and the same orbit.
Once the deployment step 10 has been carricd out, a
step 20 is performed of controlling the altitudes of the
satellites as deployed in this way so as to bring each
satellite to a drift altitude, while remaining in the
initial orbit plane.
The drift altitudes are selected from a drift set
comprising a plurality of altitude values, and including
by way of example the initial altitude.
In the example shown, the drift set has three
altitudes:
the initial altitude;
a high drift altitude, higher than the initial
altitude; and
a low drift altitude, lower than the initial
altitude.
In this example, the initial altitude is 800 km, the
high drift altitude is 1500 km, and the low drift
5 altitude is 270 km.
Other implementations are possible, e.g. having a
larger or smaller number of altitudes in the drift set,
and optio$ally including the initial altitude and/or the
final altitude in the drift set.
10 In general manner, the starting altitude, the high
drift altitude, the low drift altitude, and the final
altitude all lie in the range 150 km to 75,000 krn.
By way of example, the starting altitude may lie in
the range 33,000 km to 38,000 km, and the final altitude
15 may lie in the range 20,000 km to 25,000 km.
When some of the satellites in the set have a drift
altitude that is not very high, it may be necessary to
apply a thrust force in order to compensate for
atmospheric drag and thus maintain the satellite at the
20 drift altitude.
Figure 4 shows an example of the spread of durations
for the step 20 of controlling the altitudes of the
satellites.
As shown in this figure, two satellites are
25 controlled to leave the initial altitude in order to
reach their respective drift altitudes as soon as the
deployment step 10 has been performed.
Subsequently, three other satellites are taken in
succession to their drift orbits; two satellites are thus
30 controlled in succession to reach their respective drift
orbits at about 50 days after the launch step 10, and a
fifth satellite is controlled to reach its drift orbit at
about 100 days after the deployment step 10. The sixth
satellite remains on the initial orbit.
3 5 As soon as their altitudes differ, the various
satellites shift progressively relative to one another
during a shifting step 30.
Specifically, the Earth is not a perfect sphere; in
particular it is flattened at its poles, thereby leading
to significant disturbance in its main gravitational
potential.
5 This disturbance leads to the orbits of the
satellites orbiting at different altitudes becoming
modified progressively, with the force that is exerted on
a body as a result of the Earth's gravitational field
depending on the body's distance from the Earth.
10 Thus, if the initial orbit is considered as a
reference orbit, then the satellites that are taken to
drift altitudes distinct from the initial altitude have
orbits that become modified progressively relative to the
initial orbit. This modification of orbit leads to a
15 modification to the longitude C2 of the ascending node,
which increases for satellites having a drift orbit of
altitude lower than the initial altitude, and that
decreases for satellites having a drift orbit that is
higher than the initial altitude. The trajectories of
20 the various orbits of the satellites nevertheless remain
identical; only the longitude C2 of the ascending node
changes.
It should be observed that the initial orbit in this
example is selected as the reference orbit, but that this
25 selection is arbitrary and serves only for describing the
shifting of the orbits of the satellites relative to one
another.
The spread in the times at which the various
satellites are sent to their respective drift altitudes
30 serves to obtain different shift values, even though
several satellites are sent to drift altitudes that are
identical.
Figure 5 shows how the longitude C2 of the ascending
node varies over time for the six satellites under
35 consideration.
In Figure 5, there can be seen several changes of
slope in the curves showing variation over time in the
longitude Q of the ascending node, with these changes in
slope corresponding to variations in the altitudes of the
satellites under consideration.
The satellites are subsequently taken to a final
orbit in a final control step 40, in which the various
satellites are controlled so as to be taken to a final
altitude from their respective drift altitudes.
The altitude typically lies between the
initial altitude and the low drift altitude. The final
altitude may also belong to the drift set; all or some of
the satellites then do not need to change their altitude
during this final control step 40. The final altitude
lies typically in the range 200 km to 800 km, which
corresponds to the altitude commonly used for observation
satellites.
In the example shown, the satellites are taken
successively to a final altitude of 420 km.
In a variant, the satellites may be taken to the
final altitude simultaneously, or they may be taken in
groups.
This final control step is configured so that once
the satellites have been taken to the final altitude,
their respective orbits are shifted relative to one
another, e.g. in such a manner that the various orbits
present the same identical angle of inclination -i.
relative to the equatorial plane, while the longitude Q
of the ascending node of each orbit is different. By of
example, the angle of inclination may be 96". The angle
of inclination of the final orbits may be identical to or
different from the angle of inclination of the initial
orbits.
The final orbits of the various satellites thus
follow trajectories that are identical but that are
shifted relative to one another by an angle of rotation
about a common axis, specifical1.y the axis of rotation of
the Earth.
In the embodiment shown, the control steps 20, the
drift step 30, and the final control step 40 are
configured in such a manner that the variation in the
longitude CZ of the ascending node between orbits of two
5 adjacent satellites is constant once the satellites have
been taken to the final altitude, and in this example it
is equal to 15".
As can be seen in Figure 4, the control steps 20,
the shifting step 30, and the final control step 40
10 overlap; for example, some satellites are performing
their final control step 40 while others are still in the
shifting step 30.
Figure 6 shows an example of a constellation of
satellites deployed using a method in an aspect of the
15 invention.
This figure shows the orbits 01 to 06 of six
satellites as described above with reference to Figures 4
and 5. As stated above, the control step 20, the drift
step 30, and the final control step 40 are configured so
20 that the variation in the longitude fi of the ascending
node between the orbits of two adjacent satellites is
such that once the satellites have been taken to the
final altitude, the resulting dQ is constant between any
two successive orbits.
25 The method as described above thus enables a
plurality of satellites in a satellite constellation to
be deployed in a single launch, and enables their
respective orbits to be shifted mutually by making use of
the gravitational potential of the Earth. The method may
30 be used to deploy satellite constellations with orbits of
any type: circular, elliptical, low orbit, or high orbit.
This method thus makes it possible to avoid
consuming a large amount of propellant in order to modify
the orbits of the satellites, and to reduce significantly
35 the number of launches that are needed for such
deployment.
As a function of the desired final orbits, a
constellation of satellites can thus be put into place in
less than one year.
By way of example, the deployment of a constellation
5 of 12 satellites by means of two launches, each carrying
six satellites, and by performing the proposed method can
be achieved within less than one year.

CLAIMS
1. A method of deploying a constellation of satellites,
the method comprising the following steps:
using a single launcher to deploy (1.0) a plurality
5 of satellites at the same initial altitude on the same
initial orbit;
controlling (20) said satellites in such a manner
that, depi5nding on the initial orbit, each satellite
reaches a drift altitude selected from a drift set, with
10 the orbits of the various satellites shifting (30)
relative to one another at the respective drift altitudes
under the effect of the gravitational potential of the
Earth; and
controlling (40) the satellites in such a manner
15 as to be moved sequentially in order to reach the same
final altitude, said sequential movement being performed
in such a manner that the satellites describe final
orbits having identical trajectories with the same angle
of inclination relative to the equatorial plane, but
20 presenting distinct longitudes for their ascending nodes.
2. A method according to claim 1, wherein the final
orbits of the satellites are angularly offset relative to
one another about the Earth's axis of rotation.
25
3. A method according to claim 2, wherein said final
orbits present a constant angular offset (dQ) between two
successive final orbits.
30 4. A method according to any one of claims 1 to 3,
wherein said drift set comprises a high drift altitude
and a low drift altitude, having respectively an altitudem that is higher and an altitude that is lower than the initial altitude.
5. A method according to claim 4, wherein the starting altitude, the high drift altitude, the low drift altitude, and the final altitude lie in the range 150 km
to 75,000 km.
b. A rnetnoa according to claim 4 or claim 5, wherein the
final altitude of said satellites lies between the
initial altitude and the low drift altitude.
•7. A method .according to any one of claims 1 to 6,
wherein said final altitude lies in the range 200 km to
800 km.
8. A method ^according to claim 6, wherein the starting
altitude is '8 00 km, the high drift altitude is 1500 km,
me low aiLimde is 270 Ion, and the final altitude is
420 km.
9. A method according to claim 5, wherein the starting
altitude lies in the range 33,000 km to 38,000 km, and
the final altitude lies in the range 20,000 km to
25,000 km.
10. A method according to any one of claims 1 to 9,
wherein the final orbits of said satellites have angles
of inclination that are different from their initial
orbits.