Lecture Module 4 (Lectures 1-3) This module gives an

1
Lecture Module 4 (Lectures 1-3)
This module gives an overview of various attitude determination sensors
and hardware.
Attitude determination sensors:
An attitude sensor measures the orientation of what is known as a reference vector with
respect to a frame of reference fixed to the satellite. One or more reference vectors are
usually required to determine attitude within acceptable degree of accuracy. Reference
vectors are defined by the lines joining the center-of-mass of the satellite (also origin of
satellite body fixed frame of reference) and some external relatively inertial bodies. Among
the external bodies which can be considered as inertial (for practical purposes and for
acceptable engineering estimates) are Earth’s magnetic field, Sun, a known star, or the center
of Earth. Two types of sensors are widely popular: (i) Sun sensors, these are based on
detecting electromagnetic radiation from Sun and, and (ii) Earth horizon sensors, based on
detecting infrared radiation. Usually, a satellite is equipped with a number of such sensors of
various types.
Attitude sensors are named based on the reference inertial bodies they are supposed to
track. For example, sun sensors are named so because the light source from sun is reference
signal for measuring the orientation of a satellite in this case. Similarly, horizon attitude
sensors are based on detecting horizons of the reference object. Star tracker attitude sensors
are used for maximum accuracy, wherein, attitude is measured with respect to a distance star.
Depending upon the convenience of measurements, these sensors are placed either on the
rotating part of the satellite or on the stationary part. We learn some details about various
types of attitude sensors and their hardware in the following sections.
Sun sensors:
Sun being sufficiently far away, brightest, and nearly stationary for most operational
satellites, it can be treated as an inertial reference and a point source. Its angular radius is
0.267deg at 1AU which is nearly orbit independent. Another unique advantage with Sun is
that it also is a free source of energy to the satellite and solar panels are positioned on satellite
to absorb maximum radiation from Sun for satellite power requirements. The line joining the
center-of-mass and Sun thus can be considered as an inertial reference vector, which is used
for attitude determination. Three basic types of sun sensors are:
Prof. Nandan K Sinha
Aerospace Engineering
IIT Madras
2
a. Analog sun sensor: has an output signal which is a continuous function of the Sun
angle and is usually monotonic;
b. Sun presence sensor: provides a constant output signal whenever the Sun is in the
Field of View (FOV); and
c. Digital sensor: provides an encoded discrete output which is a function of the Sun
angle.
Analog Sun sensors
Also called ‘Cosine detectors’ function on the principle of energy deposited in a photocell
which outputs an electric current varying as cosine function of the Sun angle as shown in Fig.
4.1.
SUN
P

ds
Solar cell
Figure 4.1: Electromagnetic radiation from Sun incident on solar cell.
In Fig. 4.1, “Poynting vector” P gives the magnitude and direction of the energy flow of
electromagnetic radiation from Sun onto the photocell patched on the surface of satellite. The
energy flux E, through an elemental surface (of the Solar/photovoltaic cell) ds with unit
normal nˆ can thus be found using the expression
E  P.nˆ ds
Prof. Nandan K Sinha
(4.1)
Aerospace Engineering
IIT Madras
3
0
+ 90
Figure 4.2: Output current as a function of angle  .
The output current from the energy deposited in the photocell is given by the expression
I ( )  I (0) cos 
which is plotted in Fig. 4.2. I (0) is the value of the current when “Poynting vector” P is
directly normal to the photocell surface, i.e.   0 .  is the angle which may be used to
determine the orientation of the satellite with respect to Sun. There are some losses due to
transmission (Fresnel reflection), the effective cell area, and angle dependent reflection at the
air-cell interface, which have not been accounted for in the above expressions.
Reference axis
FOV Detector 2
Detector 1
FOV
Figure 4.3: Group cosine detectors.
A variant of analog sensor is group cosine detectors or “eyes”. Each solar cell in this case
has a different field of view (FOV) as shown in the Fig. 4.3. The output current in this case
Prof. Nandan K Sinha
Aerospace Engineering
IIT Madras
4
cover a wider range of angle due to different field of view of the two detectors as shown in
the Fig. 4.3. The output current is plotted below in Fig. 4.4. Note that now the coverage angle
of each photocell is shifted by an angle  and together two photocell provide coverage angle
beyond  90deg. The blue line is the summed output of the two sensors.
0
+ 90
Figure 4.4: Output current as a function of angle  .
Another variant of analog sun sensor uses a bar or mask to shadow a portion of one of more
photocells. One-axis and two-axis mask sun detectors are common.
Sun presence detectors: For these sensors, Sun light in the Field of View of the sensor
generates a step function response. These are used to protect instrumentation, to activate
hardware, and to position the satellite or experiments.
Digital sun sensors: consists of two components, a command component and a measurement
component. The command component acts much like a sun presence detector while the
measurement component generates an output which is a digital representation of the angle  ,
whenever Sun is in the field-of-view (FOV).
Prof. Nandan K Sinha
Aerospace Engineering
IIT Madras
5
Spin axis
Sun
FOV
FOV
Figure 4.5: Attitude sensor mounted on the body of a satellite.
Horizon sensors/Infrared Earth sensors:
These are typically used for estimating a satellite’s orientation with respect to Earth. Since
the field of view of Earth for a near-Earth satellite is quite large as compared to that of some
of the stars and Sun treated as point source, correct attitude determination of a satellite using
Earth is a challenging task. This is the reason why most Earth based attitude sensors are
designed to locate Earth’s horizon. Locating horizon of planetary bodies with atmosphere
such as that for Earth poses another challenge to horizon sensor designers for the simple
reason that there is a variation (usually a gradual decrease with altitude) in the radiation
intensity (coming from the true or hard horizon of the solid surface) through atmosphere.
Most horizon sensors are infrared detectors (for Earth based in the 14- to 16- m CO2 band)
because they are unaffected by night (less susceptible to sunlight). Horizon sensors are
normally characterized by the way they are mounted on the satellite or the scanning
mechanism. Some are mounted on the satellite at an angle to the spin axis as shown in Fig.
4.6. Wheel mounted horizon sensors are attached to the momentum wheel of the satellite.
Sensor components
A typical horizon sensor consists of four basic components:
Prof. Nandan K Sinha
Aerospace Engineering
IIT Madras
6
a. A scanning mechanism: consists of a mechanism to scan the horizon of the Earth or
another reference planet, for example, a satellite in Lunar orbit will use horizon of
Lunar surface.
b. An optical system: consists of a filter to limit the spectral band and a lens to focus the
target image on the radiance detector.
c. A radiance detector: used to detect the presence of a horizon. It is classified by its
region of spectral sensitivity. Different types of radiance detectors are: photodiode
(working principle is light falling onto it increases the number of electrons and holes
in the junction area thereby increasing the leakage current), a bolometer (a resistance
thermometer), a thermopile, and pyroelectric detectors.
d. Signal processing electronics.
Scanning mechanism: Horizon sensors have to be mounted on the spinning part on a
satellite for scanning purposes. For a spinning satellites, sensors can be mounted on the body
of the satellite at an angle relative to the spin axis, with a circular or square field of view of
diameter approximately 2 deg for good accuracy of measurements. Wheel mounted horizon
sensors are similar to the body mounted sensors except that they are attached to the
momentum wheel of satellite and spinning of momentum wheel provides the scanning
motion.
Optical system: The optical system of a horizon sensor consists of filter to limit the observed
spectral band and a lens to focus the target image on the radiance detector.
Radiance detector: Different types of radiance detectors used on satellites are photodiode,
bolometer, thermopile and pyroelectric detectors.
-
A photodiode consists of a P-N junction operated under reverse bias. Light falling on
the photodiode increases the number of electrons and holes in the junction region,
thereby, increasing the leakage current.
-
A bolometer is a very sensitive thermometer or thermistor used to detect infrared
radiation.
-
A thermopile detector consists of a string of thermocouple junctions connected in
series.
-
A pyroelectric detector consists of a thin crystal slab, such as triglycine sulphate
sandwiched between two electrodes. Impinging radiation on crystal slab increases its
temperature and thus instantaneously polarizing the crystal material yielding into a
Prof. Nandan K Sinha
Aerospace Engineering
IIT Madras
7
measurable potential difference across the electrodes. Pyroelectric detectors are fast
and have a high signal-to-noise ratio and no low-frequency noise.
In-triggering
Spin axis
Sensor field of view cone
Out-triggering
Figure 4.7: Detection of Earth by a Horizon scanner.
Output from a scanning horizon sensor is a measure of time between the sensing of a
reference direction and the electronic pulse generated when the radiance detector output
reaches or falls below a selected threshold value.
Prof. Nandan K Sinha
Aerospace Engineering
IIT Madras
Detection output
8
TREF
TAOS
TMID
TLOS
Electronic threshold
tI
tO
tw
tM
Time
Figure 4.8: Output from a body mounted horizon sensor with a Sun reference pulse.
The reference direction for a body fitted horizon sensor is often a Sun pulse picked up from a
separate sensor. When the output signal is increasing across the threshold, this phase is noted
as acquisition of signal beginning at time TREF. When the output signal is decreasing across
the threshold, the phase is called at loss of signal, beginning at time T LOS. The AOS
(acquisition of signal) and LOS (loss of signal) pulses are also referred to as in-crossing/intriggering and out-crossing/out-triggering, respectively. The output signal detected above the
electronic system threshold determines the time between the TAOS and TLOS and subsequently
define the width of Earth calibrated in time. Thus, tw = TLOS – TAOS, which is sometimes
referred as Earth width. Various electronic systems provide reference to AOS time tI = TAOS –
TREF, the reference to LOS time, to = TAOS – TREF, tw and the reference to midscan time, tm =
[(TLOS + TAOS)/2 – TREF]. Percentage of the total time that the signal is above threshold is the
duty cycle.
It should be obvious that depending on how a satellite body is oriented with respect to
Earth, the horizon size (the arc segment on Earth surface) detected and subsequently the time
taken to scan the horizon would be different. Thus tw calibrated against the attitude angle
directly determines the attitude angles.
Prof. Nandan K Sinha
Aerospace Engineering
IIT Madras
9
For momentum wheel fixed horizon sensors detection output is as shown in the Fig.
Detection output
4.9.
TAOS
TREF
TMID
TLOS
Electronic threshold
Time
Figure 4.9: Output from a momentum wheel mounted horizon sensor with a magnetic pickoff reference pulse.
In the above Fig. 4.9, TREF is counted from the time at which amplitude of output signal is
maximum.
Prof. Nandan K Sinha
Aerospace Engineering
IIT Madras

Download Report