2. Preliminary requirements (1)
• Coverage: temperate latitude belt up to +- 50° (tbc). The coverage
of the polar caps is: optional
• Fully operational system: two or three satellite system, all
operating. Graceful degradation with just two spacecraft
operational
• Orbital planes: tbd, under evaluation;
• Pancromatic camera with better than 0.5 m grd resol. at nadir
• Images: square, around 6 x 6 km (goal: 12 x 12 km) ;
• continuous strips 6 km wide (goal: 12 km) x 70 km long
• Optical beam repointing: up to +- 35° cross-track
• Optional: up to +- 30° along track fast repointing ( for 3D
imaging)
• Operational duty: nominal average of 0.2% on a per-orbit basis,
with a maximum of 1 % on one orbit per day;
3. Preliminary Requirements (2)
• On board memory : compatible with 50% (tbc) of the data
acquired daily;
• Gathered data download data rate: < 150 Mbps, 1 or 2 channels
• On-board data download antennas: 1 or 2,directional, repointable,
providing simultaneous or independent operation;
• Ground receive data stations: multiple, provided with antennas
with diameter in the 4 to 5 m range;
• Full system: compatible with single VEGA launch and other
vehicles capable of multisatellite launches;
• Single spacecraft (demostration flight): launch vehicle and site is:
tbd
• Replacement or substitution satellite(s): under study
4. Impact of Requirements
• We face contrasting requirements:
a) To point the optical telescope beam at nadir while allowing
cross-track or along-track repointing, on demand;
b) To extract from the Sun the maximum energy possible
minimizing the costs;
c) To minimize the spacecraft and appendages cross-section area
to minimize drag;
d) To go down as fas as possible in altitude to increase ground
resolution while minimizing the telescope size and mass;
• Besides, the cost reduction pressure by Customers push towards
avoiding unnecessary sophistications
5. Alternate spacecraft configurations (1)
Nadir pointing
Sun vector
Pitch axis
This conventional approach deploys two solar wings which are kept Sun-pointed
via slip-rings throughout the orbit portion wherein the spacecraft is sunlit.The
telescope axis is , normally, nadir pointed.
This configuration presents a large cross-section to the air flow and is not the
‘optimum’ for low-altitude flying spacecraft. A positive factor is that the solar
panels can be kept sun-pointed thoughout the ‘active’ orbit portion, therefore the
solar cells are efficiently used. To achieve this the solar panels must be equipped
with a BAPTA and drive system, which are critical and costly items.
6. Alternate spacecraft configurations (2)
Sun vector
Pitch axis
Specular
flat surface
Optical beam direction
In this configuration a 45° flat mirror redirects the optical beam towards
nadir, Te cross section area-affecting the drag force- in minimized also
because the solar wings are edge-seen, with a positive impact on mass and
propulsion system
The solar arrays are folded onto the spacecraft body and in-orbit deployed.
The solar energy capture varies throughout the orbit thus deceasing the
mean energy collection.However the lower efficiency is counterbalanced by
a much simpler solar plant, needing only simple hinges to deploy the solar
wings.
7. Alternate spacecraft configurations (3)
Sun vector
Roll axis =
velocity vector
45° inclined flat
reflector
Nadir-looking
optical beam
This is the previous configuration rotated by 90° . The side solar panels are
provided with motorized hinges so that during the near polar portions of
the orbit one can get more power from the Sun rays than with the previous
configuration.
The advantages, to be better quantized depending on specific mission
requirements, are counterbalanced by a greater cross sectional area
impacting the drag, which results to be of the same order of magnitude ot
the first configuration.
8. Alternate spacecraft configurations (4)
• Assumptions: spacecraft body lenght x width: 1.5 x 0.6 m
• Solar panels: 1.5 x 0.6 m each
Computed mean solar panel effective area (over the sunlit orbit
portion) :
Conf. #1 = 2.7 m^2
Conf. #2 = 1.9 m^2
Conf #3 = 2.15 m^2
Computed equiv. Area for drag during the sunlit (*) orbit portion:
Conf. #1 = 2.7 m^2
Conf.#2 = 0.42 m^2
Conf. #3= 1.55 m^2
(*) in eclipse all config. are modified to minimize air drag
9. Preferred solution
Ae for drag Solar array
effectiveness
Risk and
maturity
Config. #1 High Good Reference
Config. #2 Very Low Medium Equivalent to
reference
Config. #3 Medium to
High
Slightly
better than
medium
A bit worse
than config. #2
Configuration # 2 can be advantageously pursued from a drag (and
propulsion) viewpoint , Solar Array type ( simple mechanisms) and
related cost savings. The flat specular mirror will have to be designed
and built with utmost care, however.
10. Basic formulas for optical instrument
• IFOV = Rs/H ( Rs= ground resolution; H= satellite altitude)
• F = ps/IFOV ( F= focal lenght; ps= pixel size)
• Do = 1.22 *F* /ps ( Do= pupil diameter; = wavelenght)
• Telescope size is mainly affected by Do and F.
The equations can be rewritten as:
• Do = 1.22* * H/Rs : a low H reduces the pupil size
• F= ps* H/Rs : the smaller the pixel size the better
• Most existing and planned satellite fly high to avoid counteracting the
drag using the heavy chemical propulsion. But doing so increases F and
Do with attendant significant cost and mass increases.
• Lowering the orbit altitude restores manageable values for telescope and
spacecraft mass, power and cost, but this cannot be cheaply achieved
with chemical propulsion (high mass impact) so one has to resort to
electric propulsion
11. Impact of spacecraft altitude on camera parameters
12
10
8
6
4
2
0
altitude, km
focal lenght, m
pupil diameter for 0.5 m grd resolution
1.2
1
0.8
0.6
0.4
0.2
0
350 400 450 500 550 600 650 700
orbit altitude , km
diameter, m
The Focal lenght is valid for a
specific value of the pixel size,
here 7 microns
12. Athmospheric drag compensation
• Alternatives:
• Chemical propulsion, Isp of 210 sec
• Electric propulsion, Isp of 1000 sec. (conservative, Hall-thrusters)
• The drag was computed for 7 years starting from 2012 assuming
mean values for F10.7 and Kp coefficients .
The ballistic coefficient was
computed for different
spacraft configurations
The V required to
counteract the drag was then
computed vs orbit altitude
and S/C configuration.
13. Athmospheric drag compensation
• From published forecasts of the expected Sun activity for the next
Sun cycle, which seems to be considerably more quiet than the
previous one, a mean F10.7 value around 120 can be computed.
This was increaded to 140 for margin
• With these assumptions the 5 years velocity increment required to
counteract the drag effect for the three configurations was
computed
• For the two propulsion systems the propellant mass result as
follows:
chemical electrical
• Conf # 1 163.6 kg 29.8 kg
• Conf. # 2 18.9 “ 3.9 “
• Conf. # 3 74.8 “ 14.7 “
These data show that Config. #2 could be a valid candidate in all
cases
14. Propulsion system considerations
• The data shown above imply that Configuration #2 could be
supported by a less costly chemical propulsion system, while
Configuration #1 would likely have to be equipped with electric
propulsion not to increase too much the launch mass and
spacecraft size that could make difficult the compatibility with
less expensive launchers, VEGA included;
• Nevertheless other requirements, including allowances for
‘missions of opportunity’ requiring orbit changes, or demanding
orbit injection manoeuvres, and a more precise assessment of the
spacecraft dimensions and mass – depending from the optical
instrument configuration and features- might lead to different
conclusions;
• The use of a mixed chemical and electric propulsion cannot be
excluded at this stage and will be reconsideed in the following;
15. Spacecraft architecture (1)
The satellite #2
architecture would
include a cradle
supporting the telescope, a
propulsion module (in
yellow) carrying either the
chemical or the electrical
propulsion items, a frame
supporting a 45° inclined
flat mirror, and the solar
array in three panels, two
of which stowed during
launch and deployed in
orbit by means of
motorized hinges
16. Spacecraft architecture (2)
This is a pictorial view of
the spacecraft with the solar
array deployed. Heat
rejection can occurr via the
side panels. The back of the
45° mirror is available for
carrying an ISL package for
data relay to a tbd satellite
system. Two additional
panels could be installed
between the mirror and the
cradle to support high
datarate X_band link with
Earth, and also at S_band
for TT&C and GPS
reception.
17. Spacecraft architecture (3)
• The spacecraft electronics is rather conventional with few exceptions.
• The short term pointing stability implies the elimination or reduction of
microvibrations, and suitable technologies will have to be used;
• The image center should be within a 10% of the image size which is felt
to be sufficient for surveillance / defense tasks in populated areas. This
would imply an angular pointing indetermination of 1.5 mrad, implying
the use of star sensors;
• The X_band transmission system would be based on 8PSK modulation,
effective coding and the use of on-board compression systems. The
approach of having two transmission chains centered on different center
frequencies and radiating through two separately pointed directive
antennas, provides operational flexibility, redundancy and graceful
degradation;
• The use of multiple TT&C S_band and of GPS antennas for signal
reception, is also envisaged to widen the accessibility coverage;
18. Spacecraft High Data rate transmitter
The high data rate transmission system is an outstanding spacecraft
subsystem. A candidate block diagram is shown below, with two
independent transmission chains, each carrying a wideband moulated
carrier centered on different frequencies and linked to an independently
repointable directive antenna The latter can transmit the same, or
different, data towards the same, or different, destination stations within
the istantaneously available but time-variable access area.
19. Spacecraft budgets
• Spacecraft mass and DC power budgets cannot be properly
elaborated without a precise knowledge of the Camera
geometrical envelope and of the key electrical interfaces. The
Camera lenght has, indeed, a deep impact on the structure , solar
array and other subsystems.
• However, and for reference purposes, we have estimated the mass
budget for the Camera version that we propose for an advanced
implementation based on a 0.5 m diameter pupil, a focal lenght
around 5 m , an APS-CMOS detector matrix with 7 micron pixels.
We assumed a spacecraft flying around 400 km and provided
with chemical propulsion system (worst case). The mass depends
on the redundancy implementation but we are well below the
target of 400 kg at launch.
• Once the Camera will have been chosen, we plan to reconsider
both the configuration choice ( either the #2 or the more
conventionall #1) and the overall mass budget.
20. System coverage (1)
The coverage, intended as extension of the access area, has been assessed with
one, two and three satellites. The access area limits was taken, respectively, 30°
and 40° half-cone angle. With one satellite the access are is extremely limited
and does not allow to decrease the revisit interval below 3 or 4 days. The
situation improves with 2 satellites and can be considered satisfactory with three
spacecraft .
21. System coverage (2)
These graphs show the access area improvement feasible using three satellite
spaced by 120° in true anomaly,same orbital plane. The satellite must rotate in
roll by, respectivly, +- 30° ( with a resolution loss of 22% in the cross –track
direction) or +- 40° ( with a resolution loss , cross-track, of 30% ) . This system
configuration allows a revisit interval of 1 day for almost all points of the
Earth, except small areas close to the poles.
22. VEGA compatibility
•The system coverage can be
achieved incrementally, launching
satellites one-at-the-time.
•However if achieving a very short
revisit interval is a priority issue,
then multiple simultaneous launches
are mandatory.
•In case of a three-satellite system
they could be accommodated inside
the VEGA shroud and launched
together: but the satellites cross
sections should be kept (see the side
picture) within a 650 x 950 mm
envelope.
The VEGA inner shroud diameter is
2380 mm and can accommodate one
satellite or even two spacecraft with
cross-sections around 1.1 x 1 m.
