• Author: Hurtado Jairo Alberto PhD
  • Description:

    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)
    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].
    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.
    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
    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).

  • Year: 2012
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