Electro-Optical Attitude Sensors Basic concepts and application in

Transcription

Electro-Optical Attitude Sensors
Basic concepts and application
in the GNC of spacecrafts
Franco Boldrini – June 16, 2015
Space Attitude Sensor Activities – A long history
Since 1864
Selex ES activities in space date back
to the mid of the ’60 years when Officine
Galileo and FIAR have participated to
the first European programs promoted
by the European Agencies ELDO
(European Launcher Development
Organisation) and ESRO (European
Space Research Organisation).
At that time the Infra-Red and Electro-Optical technologies, mastered
by the Firenze plant for their Military Applications, where used to
develop the first attitude sensors to be used by European Satellites.
The first IRES generation was born, and was just the beginning …
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© 2014 Selex ES S.p.A. – All rights reserved
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What makes challenging the spacecrafts’ attitude sensing
Hostile Operative Environment:
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extreme survival/operative temperatures
need to withstand UV radiations and Cosmic Rays
micrometeorites
Launch stresses:
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vibrations and shocks
Need to guarantee:
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high reliability
long operative life
high performance
reduced mass, volume, power consumption
⇐ Tecnological Trend
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Attitude Sensors Today:
–  Selex ES is a Worldwide leader in the a2tude sensors field. –  Our a2tude sensors supply hundreds of missions with not a single in orbit failure. IRES – N2 (GEO)
IRES – C (LEO)
SSS Sun sensor
NAV-CAM
AA-STR
SPACESTAR
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With our Attitude Sensors we are on board the satellites of …
•  ASI (IT),
•  ESA,
•  ARSAT (Argentina),
•  Compagnia Generale Spazio (IT),
•  CONAE (Argentina)
•  JPL (USA),
•  JHU-APL (USA),
•  Lockheed Martin (USA),
•  Northrop Grumman (USA),
•  NASA GSFC (USA),
•  NASA Ames (USA),
•  OHB (Germany, Sweden)
•  Thales Alenia Space (FR&IT),
•  SSTL (UK),
•  CAST (China),
•  Astrium (UK, FR, GE),
•  Dutch Space (Holland),
•  NEC Toshiba (Japan),
•  SENER (Spain),
•  etc.
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At the beginning the satellites were spin stabilized
Since 1864
The easiest form of attitude stabilization is to give the rigid body an initial spin
around an axis of minimum or maximum moment of inertia. The body will then have a
stable rotation in inertial space
A spin-stabilized satellite is a satellite which has the motion of one axis held (relatively)
fixed by spinning the satellite around that axis, using the gyroscopic effect.
If satellite initially has a fixed orientation relative to inertial space, it will start to rotate,
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because it will always be subject to small torques.
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We started with the Earth Horizon Sensors
Since 1864
HORIZON CROSSING INDICATOR is
a single beam IR static sensor
designed and manufactured
under an ESA development
contract and flown on board the
ESRO IV satellite, launched in
November, 1972.
DUAL BEAM INFRA-RED sensor is
a dual beam IR static earth
sensor.
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Earth Horizon Sensors working principle
The Earth Sensor Unit consists of three
static IR beams and a processing
electronics which provides logic pulses,
corresponding to the Space/Earth and
Earth/Space transitions, when the
beam, rotating at constant speed about
the Z-axis (parallel to the satellite spin
axis) crosses the disk of the Earth.
The beams pointing axis lies in a
meridian plane
of1864
the satellite, at
Since
angles appropriate to satisfy various
earth coverage requirements.
The earth aspect angle β is function of
the angles ξ, the earth angular radius ρ
and the intercepted chord ϕc :
Warning: Requires significant Earth Radiance calibration efforts
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And then the brightest star … our Sun
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Since 1864
Wavelength band: silicium (0.4-1.1 µm)
Slit sensor FOV: ± 80 deg
Elev. angle measur. range: ± 60 deg
Detector: 8 silicon photodiodes
Spin rate: up to 120 rpm
Operating altitude: > 5000 km
Accuracy:
•  random error: < 0.01 deg
•  systematic error: < 0.035 deg
•  Insensitive to Earth albedo
X Beam Sun Sensor (XBSS) is a static sensor used on board spinning spacecraft to
produce azimuth reference pulses and to determine spin rate and sun elevation angle.
It consists of two slit sensors, each including four photodetectors placed behind a
narrow slit that defines an optical plane with four knife edges mounted in contact with
the photodetectors. The meridian slit is perpendicular to the mounting base; the skew
slit is 28 degrees tilted with respect to the meridian slit.
The optical planes of the slit define an X shaped field of view. The presence of the sun
in the optical plane generates an electrical output pulse when the spacecraft rotates
about its spin axis. The sun elevation angle is computed using the time separation of
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the pulses generated by the meridian and skew slit sensors. © 2014 Selex ES S.p.A. – All rights reserved
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Or a combination of the two …
EARTH ELEVATION SENSOR
•  Detector: immersed thermistor bolometer
•  Field of view: 1.5 x 1.5 deg
•  Wavelength band: 14 - 16.25 µm
•  Random error (3σ): < 0.15 deg
•  Systematic error: < 0.20 deg
SUN ELEVATION SENSOR
•  Detector: silicon photocells
•  Field of view: ± 80 deg
•  Wavelength
band:
Since
18640.4 - 1.1 µm
•  Random error (3σ): < 0.01 deg
•  Systematic error: < 0.035 deg
The Combined Earth and Sun Sensor (ESS) is constituted by a dual beam Infra-Red
sensor (Earth Elevation Sensor - ESS) and a two "V" slits sun sensor (Sun Elevation
Sensor - SES).
A family of these sensors was flown starting with SIRIO (1977) and OTS (ESA Orbital
Test Satellite, 1978) on board of all of the major European spinning and/or spinning
transfer orbit satellites.
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Later the satellites were 3 axis stabilized
Since 1864
The three-axis, body-stabilized design provides significant improvements. Separate
instruments may be operated independently, which enables the satellites to
continuously obtain data from multiple sources, instead of alternating between the
operating modes.
The 3 axis stabiles satellites are also capable of capturing higher resolution images.
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Earth Sensors for spacecrafts stabilized on 3 axis
Since 1864
Low Altitude Conical Earth Sensor LACES is a two-axis (pitch and roll) low
altitude (between less than 100 km and
more than 2000 km) Earth sensor.
The standard configuration has two
optical heads and one electronic unit with
two fully independent channels. Attitude
determination can also be obtained using
one optical head in case the S/C altitude
is known.
The principle of operation is based on optomechanical modulation, by conical scanning,
of the incident radiation from the Earth’s horizon in the 14-16.25 µm wavelength band.
The field of view is deflected by 45 deg by a continuously rotating prism and, when it
crosses the Earth edges, a bolometer generates an output signal from which S/E and
E/S signals are obtained by amplification and electronic processing.
LACES has been flown on ESA’s Eureca platform, ASI’s Tethered satellite and other
commercial programmes.
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LACES operation principle
Since 1864
The prism is continuously rotated by the scanning
system at a speed of 4 rev./sec. determining a
conical scan path.
When the FOV crosses the Earth edges, the
bolometer generates an output signal from which
the Space/Earth and Earth/Space (S/E and E/S)
signals are obtained after amplification and
electronic processing.
A phase comparison between the S/E and E/S
signals with the zero vertical reference allows the
Pitch angle measurement. The Roll angle is
obtained by the sensor measurement of the
chord, combined with a scale factor which is
function of the spacecraft altitude.
The outputs of the sensor are two chords Δ1 and
Δ2 which represent the angle between the
reference and the E/S, S/E crossings.
Also in this case … requires significant calibration efforts
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Earth Sensors for spacecrafts stabilized on 3 axis
Since 1864
IRES in 1970
today
IRES N2 is a GEO Earth sensor able to detect the Earth horizon and
to provide direct measurement of the Pitch and Roll attitude angles
(total error is smaller than 0.03 deg) of the Earth disk centre with
reference to the sensor Field Of View (FOV) centre.
IRES has been sold in more than 500 units and is flying since 1978.
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IRES N2 High Level DescripFon Pitch & Roll
measurement
based on Earth
border detection
•  The Earth thermal power (IR source) is modulated by the Scanning Mirror of the sensor and
recognised by four Thermal IR Detectors (Earth/Space transitions)
•  Using the angular scanning information coming from the Incremental Encoder, together with
the Earth/Space detected transitions, the Electronic Package is able to compute and to
provide:
•  Earth position in terms of Pitch and Roll angular coordinates
•  Earth Presence (EP)
•  Sensor Status information (that describes the sensor internal status)
•  With a dedicated Telecommand word, the sensor scanning mode can be configured (Wide/
Narrow), as well as the channel inhibition for each detection channel (four channels)
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IRES N2 Architecture •  Frictionless Scanning Mirror
suspended by flex pivots designed for indefinite life
applications. A limited angular torque motor provides the
motion of the scan mirror at a nominal frequency of
10Hz
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Four Thermal IR Detectors
scanning the Earth horizon along a scan path at 45°
latitude north and south that allowing no performance
degradation when the Sun enters the FOV. They are
allocated on the back side of an infrared optical system
fully packaged with the processing electronics.
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Incremental Encoder
to have the scanning angular data necessary to process
the Earth crossing information.
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Electronics Package
consisting of two separated PCBs: a Mainboard,
composed of three sections connected by rigid flex
cables and a separated DC\DC Converter PCB.
The DC\DC Converter PCB is allocated into a dedicated
aluminium box for EMI purposes.
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New Sta(c Earth Sensors Earth angular position estimation based on the signals coming from
the measurement pixel S and two reference pixels W (full Earth)
and B (cold Space) è Chord Information CI
S−B
ρ ⋅W + (1 − ρ )⋅ B − B
CI = X =
⋅L =
⋅L = ρ⋅L
W −B
W −B
0 ≤ ρ ≤1
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Coarse Sun Sensors for spacecrafts stabilized on 3 axis
Since 1864
The Coarse Analog Sun Sensor (CASS) is a fully
redunded two-axes sensor for measurement of the
angular position of the sun with respect to the sensor
reference frame. It is based on solar cells detectors;
2 (nominal) plus 2 (redundant) cells for each axis, 8
cells in total. The sensor is constituted of a central
truncated pyramid of square section with the lateral
sides inclined of an angle = 22deg, with respect to
the sensor line of site (x-axis). The cells are mounted
on the pyramid lateral sides; 1 (nominal) plus 1
(redundant) for each side..
The sensor is designed to achieve 2π sterad field of view. A suitable baffling system
allows the avoidance of spurious light reflection on the detectors, from antennas and
satellite structures. The baffling system is constitued of three concentric rings fitted
around the detector pyramid and whose height must be adapted to meet the baffling
requirements of each sensor on the satellite.
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Fine Sun Sensors for spacecrafts stabilized on 3 axis
Since 1864
FINE DIGITAL SUN SENSOR (FDSS) is a
two-axis sensor for the accurate
measurement of the satellite position
with respect to the sun.
Two axes information is obtained by
means of two optical heads mounted
with 90 degrees between them.
The optical head consists of a light
entrance slit and of a linear array
detector disposed below the slit and
perpendicular to it: in such a way that
there is a light spot on the detector.
Suitable electronics processes the detector output in order to evaluate the center of
the light spot. The field of view (FOV) in the insensitive axis direction depends on the
slit length and then can be easily modified to meet various mission specifications.
The FOV in the sensitive axis depends on the optical head dimensions and can be
adapted to mission requirements.
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Modern 2 Axis Solar Sensors
based on APS detector
ADVANTAGES AND FEATURES
•  Broad FOV for accurate attitude measurement
•  Reduced weight and dimension (removable
alignment mirror),
•  Radiation hardening (up to 100 KRAD, MRAD
for the detector)
•  Cost Effective
ARCHITECTURE
•  Photoengraved pin hole with attenuation Filter
•  Optical Head for the detection of the Sun spot
image based on the APS
•  Proximity electronics for timing, interfaces
management, data and algorithm
processing, based on an ASIC
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Working Principle The Sun position is achieved by a centroid algorithm calculated on the spot
produced on the bi- dimensional APS matrix
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700
600
Digital Number
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x coordinate
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y coordinate
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On Axis Acquired Sun Spot 110
550
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x coordinate
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300
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Digital Number
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x coordinate
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532
100
530
y coordinate
528
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526
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522
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518
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110
x coordinate
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64° Incident Acquired Sun Spot © 2014 Selex ES S.p.A. – All rights reserved
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The future: Sun Sensor on Chip
CONTINOUS INNOVATION
•  Following the recent trend for miniaturization and the latest innovation steps in the
CMOS technology, the integration on the same chip of a detector pixel array and the
driving & processing logic has become mature for applications in the space field.
•  A prototype of a miniaturized Sun Sensor on Chip (SSoC) was developed, manufactured
and tested by SES with promising results. The SSoC activities demonstrated that a true
System on Chip can be made with a MEMS pin hole optics directly assembled on the
Silicon.
SSS Digital Sun Sensor
Mockup of Sun Sensor on Chip
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SSoC vs SSS: Key Benefits
•  1/3 the Size, 1/3 the Mass, Same Performance
•  Fully digital insensitive to albedo (by design)
•  A Coarse version (with no dedicated calibration) may be sold for a price comparable to
an analog sun sensors (truncated pyramid)
•  A full accuracy version is achieved applying a suitable on ground calibration
•  Environmental on-ground qualification will be completed Q4 2015
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Data Sheet - SSS versus SSoC
SSS SSoC Accuracy <0.03 deg (3σ) <0.016 deg (3σ) Power Bus 50V / pre-‐regulated +5V±10% pre-‐regulated +5V±10% Power Consump(on <700 mW (without DC/DC) <930 mW (with DC/DC) < 300 mW Radia(on Hardness 100 Krad with an extension to 300 Krad > 300 Krad FOV from ±64° to ±45° 64° half cone Mass < 330 gr 102 g Dimension 112 x 110 x 43 mm 58 x 60 x 23.6 mm Reliability 350 FIT 95 FIT I/O Interfaces RS422 RS422 SpaceWire Camera Mode Not available Only with SpaceWire Interface @ reduced rate 3,1Hz (8 bit) -‐ 1,5 Hz (10 bit) © 2014 Selex ES S.p.A. – All rights reserved
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Sun detection and tracking
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Autonomous sensor with two main operating modes
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Automatic “power on to Sun tracking” transitions
Searching…
…Sun is found
X
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X
XY Sun coordinates calculated as Sun spot barycentre
FPA
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FPA
Y
Sun Acquisition Mode (SAM), default at power on
Y
Sun Acquisition Mode (SAM)
•  Full frame reading (programmable, nominal 511x511)
•  Automatic transition when Sun is detected
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Sun Tracking Mode (STM)
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Sun is present
Sun is lost
Window reading (80x80 pixels, programmable)
Sun tracking with automatic window position update
Nominal cycle rate 10Hz
X
X
Automatic transition to SAM when Sun is lost
FPA
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Y
Optional 8 or 10-bit full image download, at reduced rate
FPA
Y
Sun Tracking Mode (STM)
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Sun Sensors CalibraFon: –  CalibraFon coefficients are provided to be externally applied to the sensor output for co-‐ordinate to angle conversion and mechanical/geometrical tolerance compensaFon –  Dedicated calibraFon can improve the accuracy in a reduced FOV –  Polynomial complexity within AOCS capability (6 and 21 coeff. in figure) Residual error 1sigma
for one axis
arcsec
120
100
80
60
40
20
0
6 coeff
21 coeff
98
67
51
30
17
25
74
51
34
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107
16
45
16
FOV deg
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55
25
36
65
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Reflections Rejection 1/2
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SSoC can operate in presence of objects in the field of view (like antennas,
appendages and booms) and reflections from spacecraft.
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The logic can autonomously detect and identify up to 4 clusters of over threshold
pixels. These 4 clusters can be filtered out according to specific dimensions and
energy criteria or maintained and tracked, with relevant information reported in
telemetry, while the real Sun is correctly identified and tracked.
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“Large objects” are, by default,
discarded
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In presence of multiple Sun-like
objects:
•  The one closest to the last
known Sun position is
automatically tracked
•  A flag is set (solution not to
be trusted)
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Reflections Rejection 2/2
• 
In addition the user can define and dynamically update through telecommand two
“blind” rectangles in the detector array. All the pixels contained in the area inside
these rectangles are automatically excluded by the SSoC logic, both in terms of
background calculation and over threshold pixels determination.
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The blind rectangles may be defined based on the telemetry data (size and position
of the identified clusters) or on the raw image that can be dumped through the
SpaceWire interface.
Reflection
Blanking
rectangle
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HR-‐STR: we started from the highest accuracy • 
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in the 80ies SES started to develop a high accuracy star tracker, suitable for being used in the ESA telescopes. HR-‐STR is a high accuracy (beKer than 1 arcsec), narrow field star tracker based on an opFcal system with cathadioptric objecFve built with radiaFon hardened glasses. In December 1995 the ESA telescopes ISO and SOHO were launched, bringing to space our High ResoluFon Star Trackers (HR-‐STR) for the first Fme. Many other SES star trackers’ launches have followed since then, cumulaFng to the date more than 220 years of successful in flight operaFons. It is worth to men(on that SOHO HR-‐STRs are s(ll working nominally aQer 18 years of flight at L1, allowing this World-‐class science mission to be extended un(l the end of 2016. © 2014 Selex ES S.p.A. – All rights reserved
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CASSINI SRU: our first “autonomous” Star Tracker • 
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The first large FOV Star Tracker of Selex ES was developed for the CASSINI satellite, launched in October 1997 and sFll behaving nominally to the date. The Stellar Reference Unit (SRU) performance has been evaluated by JPL (Selex ES customer) through ground qualificaFon tests whose results were confirmed by the in-‐flight experience From this first acFvity with the parFcipaFon of JPL (in charge of the relevant Electronic Processing Unit), SES developed a new product series which is currently on board Rosepa (where we also have the NavigaFon Camera), Mars Express, Venus Express. Cassini SRU
Rosetta, MEX & VEX STR
Rosetta Navigation Camera
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Star Sensors Development Roadmap
Miniaturized
Sensors “on chip”
Image detector, signal
processing, power supply,
communication interface on the
same chip
APS & MEMS APS CCD AA-STR APS Autonomous STR
Attitude measurement (autonomous)
CCD A-STR CCD Autonomous STR
Mass 2.6 kg
Attitude measurement (autonomous)
Flying since 2009 (Proba 2, Alphabus,
Bepi Colombo, Astro-G, SpaceBus 4000,
ExoMars ...)
Mass 3.5 kg
Mass 7.2 kg + baffle
Flying since 1997 (Cosmo, Stereo,
Messenger, Radarsat, MRO, Pluto,
Herschel/Planck, Phoenix, LRO,
LCROSS, SDO, GAIA, Sentine-1l,
Grail, GPM JWST…)
Flying since 1995
(ISO, SAX, SOHO, Integral ...)
CCD based Optical Head (OH)
Conventional HR STR
Star position measrurement
APS based Optical Heads (OH)
Juno
1990 + Dedicated
PFC Board
Option
SPACESTAR (on board IRN)
2001 2009 2012 © 2014 Selex ES S.p.A. – All rights reserved
toward 2020... 32
Case Study:
Autonomous Star Tracker
An autonomous star sensor is able to
provide robust and accurate three axis
attitude determination without any a
priori information of attitude and
angular rate
The attitude determination is nominally
performed by comparing Star Images
measured on a detector to known star
positions and characteristics stored in
a Star Catalogue within the sensor
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Star Tracker Main Elements
baffle
optics
detector
Proximity &
Processing
Electronics
Power Supply
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A-STR ELECTRONIC ARCHITECTURE
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A-STR SOFTWARE
Stars Field
Star position &
Brightness
Stars
Identification
On Board
Stars Catalogue
Attitude
Determination
On Board SW
Calculated Attitude
§ Attitude Acquisition
§ Attitude Updating&Tracking
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Autonomous Star Trackers Main Opera(ve Modes Lost In Space Mode
Tracking Mode
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A-STR: an example of products’ flexibility
The modular design of the A-STR product allowed SES to employ a common design for
a broad range of missions.
  The Herschel and Planck programs of the European Space Agency (ESA) are an
example of the flexibility of the A-STR design
•  The A-STR is used both in the extremely accurate pointing Herschel Telescope (3
axis stabilized satellite) and in the slowly spinning (1 rpm) Planck spacecraft.
•  This has been achieved by implementing in the standard A-STR product an
“interlaced” tracking mode and a TDI (Time Delay Integration) mode respectively.
•  The TDI mode of the A-STR is also used in the New Horizons Pluto Kuiper Belt
mission of JHU-APL, with S/C spinning at 10 rpm
  For the JUNO Mission of JPL SES developed a derivative of the Pluto Kuiper Belt
configuration, allowing to cope with both harsh radiation requirements and spinning
satellite configuration.
  Several
custom baffles allowed the unit to be used in Lunar Missions (LRO,
LCROSS) and GEO missions (GPM)
A-STR: High Accuracy Configuration
Herschel implemented the “interlacing” mode, which allows to virtually increase the
number of stars tracked at each cycle.
In flight pointing performances showed that the A-STR and the ACMS are working well,
in some regards exceeding expectations:
The same sensor is currently in production for JWST.
Parameter
Req.
Goal
Perf.
APE Point
3.70
1.50(a)
1.80÷1.01(a)
APE Scan
3.70+0.05ω
1.50+0.03ω(a)
≈1.80
RPE Point (60 sec)
0.30
---
0.19÷0.29
RPE Scan (60 sec)
1.20
0.80(b)
0.29
ω is the scan rate in arcsec/sec
a) using the option to track up to 18 stars [STR interlacing], and after full calibration and thermal stabilization
b) STR interlacing.
ALL VALUES ARE ARCSEC 1 SIGMA
A-STR Optimitazion and tailoring for JWST
Mission needs
•  Guarantee adequate radiation hardening for
mission at L2
•  Achieve high accuracy pointing stability against
thermo elastic variations
Tailored production A-STR with
Titanum baffle in place
Thermal
isolation
Aluminum stage A-STR design and modeling
•  Use of increased exposure time for accurate star position
measurement (2x standard integration time)
Titanium stage
•  Increased box wall thickness for radiation hardening
•  Detailed mechanical and optical modeling of pointing
Increased wall
stability
thickness
•  Dedicated design of 2 stage baffle to achieve high pointing
stability (0.38" 1σ) over 24 hours, with variable Sun input
Heritage box
angles:
•  Baseplate almost isotherm (S/C controlled)
•  Lower stage in Titanium with no solar input
•  Upper stage in Aluminum with solar input
•  Thermal isolation between Al / Ti and Ti / Box
•  Specific baffle flange design to improve thermo elastic
response (flange cuts)
•  Accuracy predictions by model verified and validated by
A-STR: Same core, different mission requirements
GPM Mission (NASA)
GEO baffle and increased box wall
thickness for radiation hardening:
LRO & LCROSS Missions (NASA)
Custom baffle design:
Pluto & Plank Missions (NASA & ESA)
Capability to work in spinning (up to 10
RPM) and 3 axis stabilized mode
JUNO a custom design for a challenging mission
(Harsh radiation environments and spinning satellite)
•  The JUNO Stellar Reference Unit (SRU) operates nominally on a spinning
spacecraft launched toward Jupiter on August 5th, 2011.
•  The SRU is suitable for operations during the interplanetary cruise and the
Jupiter orbits, providing information (at 5Hz) for attitude determination at
spin rates from 0.5 rpm to 2.5 rpm, with nominal spin rate of 2 rpm.
•  While the radiation environment during the interplanetary cruise is not a
concern for the SRU operation, the high fluence of particles (i.e. electrons
and protons) in the Jupiter orbits produces permanent damage and
transient noise in the optics, in the CCD detector and in the electronics
components, which might heavily affect the performance and functionality
of the star sensor. In this instance, the unit is capable of withstanding this
level of degradation without affecting the capabilities required for the
specific mission.
•  CCD tested in flight conditions (timings and temperatures) with
protons and electrons beams at the expected Jupiter radiation levels
•  GEANT4 simulation tool used for evaluation of radiation environment
and for Optical Head structure design
•  SRU mathematical model deeply revised in order to simulate all the
radiation effects
JUNO Stellar Reference Unit
• 
• 
• 
• 
• 
Optical Head separated from the
Electronics Unit
CCD operated in Time Delay Integration
(TDI) mode
Total mass 12 Kg (focal plane shielded
by tungsten shell)
Power consumption 10W
Dedicated Operative modes to cope with
radiation (permanent and transient
effects)
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Software countermeasures
Spatial filtering and data from multiple frames are used to
discriminate stars from transient events.
This method, already developed and in-flight tested on the Selex ES
APS based autonomous star tracker for operations in a rich proton
environment, as the one expected during peak solar flares, has been
suitably modified to take into account that the stars position is not
fixed but widely changes frame by frame for the JUNO SRU
operation in a spinning spacecraft.
Expected SEU rate at CCD:
up to 2 E6 particles/cm^2/sec(!!)
are able to reach the CCD and
corrupt the images
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SEU corrupted image
Example of a simulated image at the expected particle flux rate:
500
500
Zoom of the star
among SEUs
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400
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300 380
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375
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365
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345
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0
385
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Here,
there
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385
is a star !
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0
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Hardware countermeasures
•  The Juno SRU design derives from the A-STR consolidated design, with
changes to provide sufficient shielding to the CCD focal plane and reduce
the particles fluence expected at EOL and during the perijove peaks
•  The CCD will be operated at low temperature to recover the transfer
efficiency affected by the cumulated displacement damage
Tungsten shell shielding in Brown
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AA-STR: All the advantages of adopting an APS detector
The use of the Active Pixel Sensors (APS) as image detector into the Star
tracker design allowed the development a fully digital sensor, with the
following benefits wrt CCD trackers:
•  Number of electronic components strongly reduced, benefiting overall
reliability figure, size, mass and costs.
•  Complete elimination of analog electronics reduces sensitivity to EMC/
EMI disturbances and simplify intermediate tests (no trimmings of
analog levels is required during PCB level testing)
•  APS technology does not suffer from Charge Transfer Efficiency
degradation as occurring in CCDs, thus APS technology is more suitable
especially in missions where proton and heavy ions are dominating
radiation sources.
•  APS have an intrinsic high anti-blooming
capability, allowing star tracker to operate without
performance degradation with bright objects in the
FOV (Moon, Jupiter, etc.)
•  Direct pixel readout lead to robustness to SEU. In
CCD the SEU are accumulated also during the
phases of exposure, shift and readout, while in APS
the direct readout feature limits the SEU
accumulation only to the exposure phase.
Software Improvements: Multi-Head configuration
In the frame of MTG SES has developed new algorithms that improve the overall
accuracy of the star trackers “system” (composed of 3 AA-STRs), by also removing the
effects of thermo-elastic deformations. This is achieved by running in the PFC a simple
SW (provided by SES) that takes the quaternion outputs from the 3 AA-STR and
combines them properly.
The overall accuracy obtained by the multi-head SW after the quaternions data fusion is:
•  LFE: 1 arcsec.
•  HFE: 2.5 arcsec
Multi-Head data fusion algorithm
48
SW Improvements: Background Noise Filtering
•  Dark Signal-Non Uniformity produces a systematic error in the star spot photometric
center determination, for a fixed position of the star in the FOV. This error is
dependent on star magnitude, signal distribution over the pixels used to compute the
barycenter, and on the level of non uniformity.
•  To improve the star position measurement accuracy, a dedicated background
estimator process has been implemented. The concept is based on the fact that
during tracking the star tracker knows with good accuracy the current position of the
star, the rate and the direction of the star movement. Therefore an estimation of the
non-uniformities in the area where the star will move in the next tracking cycle is
possible.
•  In the assumption that the star rate is sufficiently low (@ 0.1 deg/sec) it is possible to
sample pixels several times before the star signal is superimposed. Therefore at
every frame, the pixels of the tracking window not interested by the star signal are
sampled and a filtering is done reducing noise effects and mitigation of eventual SEU
corrupting the pixel signal during one of the frame used for background estimation.
49
SW Improvements: Background Noise Filtering
Area
inhibited for
background
estimation
2
4
Star
movement on
the detector
6
Warm
pixel #3
8
10
• 
It must be noted that this algorithm works only
when moderate angular rates are considered
(a limit around 0.1 deg/sec can be assumed).
• 
At higher rates, the efficiency of the algorithm
is null. The SW autonomously disables this
function when the angular rate higher than a
given threshold is measured.
• 
Figure shows the resulting error in the star
position measurement in the case of “standard”
algorithm (in red) and in case of the algorithm
with background compensation (in blue).
• 
The error significantly reduces as star passes
across warm pixel #2 and warm pixel #4.
Improvements are also achieved in other
areas, due to the continuous background
removal. The rms error over the 50 pixel scan
was 0.2 pixel in case of standard algorithm,
while it reduces to 0.1 pixels in case of
background removal algorithm.
Warm
pixel #2
12
14
16
18
20
2
4
6
8
10
12
14
16
18
20
1
0.8
0.6
0.4
0.2
error (pixels)
star
0
-0.2
-0.4
-0.6
-0.8
-1
50
55
60
Star passing on
pixel #2
65
warm
70
75
star
notposition
compensated
compensated
80
85
90
95
Star passing on
pixel #4
100
warm
50
Improvements from the Proba 2 Flight Demo
• 
AA-STR is flying as an experiment and is not in
the PROBA-2 AOCS control loop.
• 
Although some limitation on available resources
shall be accepted (telemetry data, mass
memory, power, etc), this allows to execute all
STR tests without impacting the spacecraft’s
operations.
• 
Several improvements have been brought to the
AA-STR products thanks to Proba 2 Flight
Experience (or validated on board the Proba 2
Flight Demo).
Proba-2 Launch on Nov. 2nd, 2009
51
In flight autonomous focal length refinement
• 
After the on ground calibration a residual error remains in the focal length knowledge,
mainly due to set up induced errors.
• 
As worst case, a 10 µm maximum error is predicted, considering the set up characteristics.
• 
In flight results have demonstrated that errors in the order of 6 µm are present.
Star pairs error (uncalibrated on left and calibrated on right) vs. interdistances using reference positions of the stars
52
In flight autonomous focal length refinement
• 
In the figure the focal length residual error achievable with the AA-STR self-calibrating
function is reported, starting from an initial error due to on ground calibration of 5 µm.
• 
After about 500 seconds the residual error reaches the steady state value.
• 
The focal length autonomous calibration function can be enabled/disabled via
telecommand. In case it is not activated the value of focal length stored in the application
SW is used.
53
Tested behavior during solar flare (onboard Proba2)
• 
During a March 2012 solar storm, Proba 2 went into Safe Mode on Wednesday 7th at
11:14:55 due to GPS corrupted data (their GPS equipment is known to be radiation
sensitive). The satellite was put back in operation at 16:50:27.
• 
The Safe Mode switched off the Selex ES APS star tracker (AA-STR) flying on board
Proba 2 as a “technology demonstration” payload.
• 
Proba 2 AA-STR Star tracker was Switched ON at 12:46 (*)
• 
AA-STR Star tracker was commanded in Stand_by mode at 12:47 (*)
• 
AAD (lost in space mode) was commanded at 12:48 (*)
• 
Pattern recognition was successfully performed within 1 second
• 
Star tracker autonomously entered in tracking mode. Accuracy during the period is
essentially unchanged with respect to other test sessions outside solar flare. 15 stars
have been continuously tracked.
• 
At the end of the experiment the AA-STR was again commanded to Stand by Mode at
13:03 (*)
(*) hours are referred to UTC of the 8th March 2012.
54
Solar Flare Proton Flux during SES STR Test
• 
The test was performed in correspondence of the “peak” of the solar activity (see
proton flux):
12:48
(source: http://www.swpc.noaa.gov/ftpmenu/plots/proton.html)
55
SPACESTAR overview
SPACESTAR means:
Satellite Platform Avionics Computer Embedding Star Tracker Algorithms & Resources
The SPACESTAR system consists of:
•  Up to three Optical Heads, each containing a baffle, optical system, focal
plane (APS based) and digital electronic module (FPGA based).
•  A Software running on the
Platform Computer (PFC) able to
elaborate the images acquired by
the OHs and output the attitude
data.
An optional configuration includes a
PFC board with µP, memory and
TEC driving circuit.
© Copyright Selex ES. All rights reserved
56
SPACESTAR – A Novel Star Trackers Architecture
  In 2010 SES was awarded a contract to provide the Star Trackers for a large
commercial constellation program
•  This program has developed a highly optimized Star Tracker
–  Centralized control processor
–  High volume test infrastructure
–  Optimized redundancy
–  Cost optimizations
  The new architecture (SPACESTAR) has the potential to provide significant
value to new space programs.
•  Optimized hardware and elimination of unnecessary redundant hardware
•  Cost efficient
•  Reduction in Size, Weight, and Power
  New
architecture has the potential to provide significant benefits, but
modifications to most legacy flight computers are necessary to interface to
the new architecture
SPACESTAR description
The SPACESTAR uses a centralized processor to control multiple ST
Optical Heads
•  Allows for a single processor to control 3 or more OHs, eliminating the
control electronics traditionally included in each OH.
•  Optics and focal plane already developed for the AA-STR are the core of the
SPACESTAR OH
•  Optical system is assembled in a titanium structure (barrel), to match
thermal expansion coefficient
•  Optical system manufactured with radiation resistance glasses, hard mounted
in the titanium structure. No use of adhesives or optical cements
•  20° full cone FOV on the HAS-2 APS detector,
•  Internal layouts of the optical barrel and the rings designed considering stray
light performance
•  Optical barrel also supports focal plane assembly, realizing the electrooptical module
58
Optical Head description
•  The APS is cooled down by a Peltier which
hot side is fixed to a copper heat sink and a
two arms thermal strap. The thermal strap is
fixed to the structure again.
•  The thermal strap design is aimed to
maximize the surface for heat exchange and
to maintain flexibility in order not to stress the
focal plan
•  The detection module (optics plus focal
plane) is included in an OH structure that
supports also the OH electronics, and
through a dedicated thermally isolating
spacer, the baffle
•  Two Spacewire connectors, one power
connector and a test connector, connected
by means of flex connections to the PCB are
then directly assembled on the external case.
•  No alignment mirror on the box – updating
with optical mirror can be performed.
59
MHSTR functionalities general overview
  Images are acquired and compressed by the OH and sent via Space-Wire
to the SBC for Multi Head STR (MHSTR) SW elaboration
  Autonomous
recognition of first attitude is possible through the
“Acquisition mode”
  Fine
attitude calculation is performed in “tracking mode”, after the
initialisation from acquisition
  Two
on board catalogues are included in the MHSTR SW for pattern
recognition and prediction of stars in FOV:
•  Triads catalogue: about 8000 entries
•  Stars catalogue: about 3000 entries
  Architecture
and algorithms based on existing SES Single Head Star
Trackers’ experiences
  Robustness
SoftWare
increased by the presence of a Multi-Head management
SPACESTAR SW
•  Multiple Head management, with modular functionality (self adapting
in-flight to availability of OH data):
•  Single head operation
•  Two heads data fusion
•  Three heads data fusion
•  In-flight relative OHs alignment calibration for compensating thermoelastic deformation and ground to orbit shifts
•  Autonomous management of OHs for blinding and optimal
performance achievement
•  Customized SW ICD in order to minimize changes to existing avionics
software
61
Performances / Examples
Data fusion (worst case with 2 OHs):
•  Bias Error
•  Pitch/Yaw: 8 arcsec (3 σ)
•  Roll: 11 arcsec (3 σ)
•  Low frequency
Error (3σ) Pitch/Yaw
Error (3σ) Roll
±5°C temperature range
2.5 arcsec
4.5 arcsec
-25°C to +50°C temperature range
10.2 arcsec
4.9 arcsec
•  High Frequency (without filtering)
Spacecraft Rate
X
(arcsec, 3σ)
Y
(arcsec, 3σ)
Z
(arcsec, 3σ)
0.1 deg/sec
7.5
8.4
7.2
0.5 deg/sec
10.8
11.6
9.9
1 deg/sec
17.3
18.6
15.9
4 deg/sec
45
45
45
62
Multiple Heads Management general overview
 Multiple Heads Management SW based on:
•  “data fusion” SW algorithms and SW developed for the
Meteosat Third Generation AA-STR product, aimed to improve
quaternion accuracy and compensation of thermo-elastic
misalignments among optical heads.
•  Up to 15 stars per OH (up to 3 OHs in the loop).
 Additional functionality to cope with “high angular rates” achieved
implementing two modes:
•  Normal mode (for moderate angular rate < 1.5 deg/sec)
•  High angular rate mode (maintain tracking with rates up to 4
deg/sec.)
High Angular Rate Mode
Used mainly for Earth Observation missions and agile satellites.
Optical Heads data fusion performed at star vector level:
 Increased robustness to rate using stars as detected by each OH
 OHs Misalignment estimation frozen to last valid values.
 Star vectors from OH_j are rotated on OH_1. Q-method on
augmented set of star vectors is applied.
CURRENT STATUS
•  Qualification achieved in 2013
•  Ramp Up Production phase completed in June 2014
•  More than 100 FMs delivered to date
•  Production phase running:
•  Current Delivery rate 9 FMs per Month
•  242 FMs to be delivered by January 2017
65
Selex ES: the Rosetta Navigation Camera
Technical Objectives
«Common» Focal
Plane
•  Determination of satellite absolute
positioning in space (Attitude Sensor
task)
•  Measurement of satellite position wrt
a “relatively near” object (Navigation
Camera task)
Mechanism for
filters’ change
Baffle
The instrument
STR OH
The Selex ES solution for Rosetta is a
sensor with a common electronic unit
and, optical head with two different
FOVs and two bespoke software. It
performs two functions:
•  Star Sensors
•  and Navigation Camera thanks to a
special mechanism which acts on
the instrument optics.
66
During the trip – First Earth Fly-by (2005)
The Earth and the Moon seen by NAVCAM
67
NAVCAM Images of the comet …
August 19th 2014 – NAVCAM ~79 km
August 21st 2014 – NAVCAM ~ 69 km
68
w rt
d
e
e
et sp 00 km/h
m
o
C
0
: 55.
n
u
S
ESA: orbiting the comet
The ʺ″orbitingʺ″ phase
setta
o
R
et to w cm/s
m
o
C
: fe
d
e
e
sp
69
Scientific results from NAVCAM
NAVCAM images
October 26th 2014
Distance:
Resolution:
Credits:ESA/SELEX ES
7.7 km
65 cm/pixel
70
Orbiting around the
comet
Comet 67P on 15 October 2014
Distance: 9.9 km from centre of
comet
Mosaic image – NavCam
Total area: 1300 m x 1300 m
(0.63 m/pixel)
ESA/Rosetta/NAVCAM
71
JUICE Navigation Camera – Operating Scenarios
In-cruise navigation:
optical measurements based upon Line of Sight (LOS)
acquisition of far objects, which can be planets, satellites or
asteroids and comets, used as beacons. Filtering techniques
are used to estimate the S/C velocity from the observations of
positions at different times, allowing the processing of nonsimultaneous (subsequent) measurements.
Far approach:
observations of the target itself provide the best positioning
accuracy in the direction perpendicular to the LOS vector.
Close approach:
target size in the camera FOV increases so that its shape can
be resolved, and its Centre of Brightness computed. When the
distance to the target is such that it is possible to identify the
body limb, limb measurements can be added to the
observables.
72
Selex ES: JUICE Navigation Camera
LOS acquisition concept (left) and Limb identification and measurement (right)
73
Thanks!
[email protected]
74

Electro-Optical Attitude Sensors Basic concepts and application in

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