Development of a low cost Sun sensor and star tracker for Aramis (Modular nanosatellite platform)

as, astronomy, atmospheric studies, communications, navigation,
reconnaissance, remote sensing, search and rescue, space exploration, and weather. All
of them provide a great deal of information in the sciences of earth and life.
Usually, a normal sized satellite weighs more than 500kg, it is a few meters long
and can cost more than 1million Euro to build and launch it.[1].
In later years, small satellites (MicroSat, NanoSat, CubeSat and PicoSat) have
been present as the beginning of the space age.
Nowadays, due to advances in microelectronics (especially in microprocessors
and the lower cost on the launching, small LEO (Low Earth Orbit) satellites are an
attractive alternative feasible instead of traditional big GEO (Geostationary Earth Orbit)
satellites.
Small satellites are cheaper to construct and launch, they can do several specific
tasks even better than large satellites, aside from the small size of the satellite.
Furthermore, all onboard components are energy thrifty, and their entire surface is used
to collect energy from the sun through solar panels.
All these advantages provided for small satellites compared with big satellites
can be summarized with the slogan “faster, smaller and cheaper”.
The difference between MicroSat, NanoSat and Pico Sat is basically the size of
each one. CubeSat is a a type of miniaturized satellite for space research with small
volume and mass (less than 1 liter and 1,33 kg). NanoSat is a satellite slightly bigger
than a CubeSat, with a mass between 1 and 10 kg [2].
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One of the most important parts in a satellite is the Attitude Determination and
Control Subsystem (ADCS). It consists of a set of attitude sensors, actuators and a
microcontroller with control algorithms.
To determine attitude, several sensors can be used: Sun sensor, Horizon scanner,
magnetometer, Star tracker, and Gyroscope. Some of these can be expensive, others
must work together because acting as standalone devices they cannot obtain
adequate/reliable results, and others are too complex or too big for small satellites.
Sun sensor is one of the most common parts in an ADCS. It determines the
satellite’s orientation relatively to the sun by measuring the amount of light that hits it.
It also can be used for instrument pointing, solar array pointing, and safe mode or sun
acquisition systems. There are several types of Sun Sensors, analog or digital, one or
two-dimensional position, all of them, with different specifications and cost. In this
project, a CMOS sensor based is used as Sun sensor.
1.1. Design Goals and Proposed Solutions
The main goal of this thesis is to estimate the attitude of a small satellite using
only image sensors in a low cost satellite.
The main objective is to process the images coming from the image sensor,
which is designed with a low cost Sun sensor that can be used also as a star tracker to
get the attitude of a small satellite (modular architecture, type ARAMIS) without use
any other attitude sensor.
There are two possible ways to solve the problem of the attitude in the satellite
using a sun sensor; the first one is based in the knowledge of the orbit parameters, along
with the measurements from the magnetometer to estimate the rotation angles. This is
the classical solution.
The second approach is presented in this thesis, using only image processing
without knowing the orbit parameters or using any other sensor.
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Another goal of this thesis is to design and evaluate a real sensor structure for
nano satellites using CMOS image sensor, as a attitude sensor and to evaluate the
possibility to estimate the position of a small satellite using only image sensors.
Such a direct approach can allow for implementations of single sensors to
estimate the attitude, and to reduce cost or space in the satellites.
1.2. Main Contributions
The first period of the PhD program was invested in the study of the literature
and state-of-the-art on the field of Sun sensors for nanoSat, complemented with the
studies of the physical implantation CMOS sensors and design of the Sun sensor for the
satellite. This was an essential part, in order to ensure a proper knowledge base before
starting working towards the aforementioned goals.
The following step was to dedicate the time to design, implementation of the
circuit and board of the Sun sensor, as a part of the modules in Aramis.
Once the schematics circuit was designed, it was changed and redesigned
because the sensor provider (Kodak) in its bankruptcy process cancelled the production
of new sensors; consequently, another sensor from a different provider had to be
selected.
The work with the circuit and the sensor consisted in capture images from the
Sun to calibrate the sensor and to adjust the parameters when neutral filters were used.
The last part of this thesis was done at Space System Laboratory at MIT
(Boston, MA. USA), where the mathematical analysis and captured image simulator
were developed to estimate the attitude of the satellite based in the measurements with
the real Sun sensor (image sensor).