Theory of Operation
The MARIE instrument is an energetic particle spectrometer
that can measure the elemental energy spectra of charged particles.
It is mounted on the science deck of the Mars Odyssey orbiter;
as the spacecraft orbits Mars, the angle between the axis
of the spectrometer's field of view and the mean interplanetary
field direction varies from 90° to 180°.
The power, mode control, and data download operations of
the MARIE instrument are controlled by the spacecraft. Commands
are sent from the ground to the spacecraft's central processing
unit (CPU) to cycle the power on MARIE and to change modes
between science mode and survival mode. When it is operating
in science mode, the MARIE acts as an autonomous data acquisition
device, and data is collected until the spacecraft issues
a mode change command. Once the instrument is in survival
mode, the spacecraft can issue commands to download data,
to change parameters, to power down, or to return to science
mode. During data download, the spacecraft controls the download
process and downlinks the data to the ground.
Image above: MARIE
Assembly, Closed (left) and Open (right).
Nine separate detectors inside MARIE are arranged in a stack
which acts like a telescope. This stack consists of two A
detectors, two position sensitive detectors
(PSD), four B detectors, and
one C (Cherenkov) detector. The A
and B detectors, which are made of silicon, are the principal
particle identifiers. Each detector records a signal proportional
to the energy deposited; the energy deposited is a function
of the particle anergy and the square of its charge (Z).  Particles
with sufficient energy pass through all the detectors; however,
some particles stop in the detector stack, and the charge
and energy of these particles can be inferred from the deposited
energy signals and the depth of penetration. If a particle
enters the telescope within the 60 degree sensitivity cone,
and it has enough energy to make it through both the A1 and
A2 detectors, it is considered a coincident event. In such
a case, all detector boards are polled by the CPU, and the
data for this event is recorded. Each detector records the
amount of energy deposited in itself. The PSD's also record
the position of the strike within the detector.
Image right: Cross-section of MARIE showing incident particle
and detectors.
The minimum proton energy required
to form an A1A2 coincidence corresponds to a proton with range
greater than the sum of the thickness of A1, PSD1, PSD2, and
a minuscule part of the A2 thickness. This adds up to 0.374
g/cm2 of Si and corresponds to a proton energy of 19.8 MeV.
Particles with energies above this threshold will be recorded
by the telescope.
The angular response functions are calculated for those particles
that give an A1A2 coincidence and also pass through PSD1 and
PSD2 detectors, since they are the only particles for which
the incidence angle can be measured. Note that not all particles
that give rise to A1A2 coincidence pass through PSD1 and PSD2
because the position sensitive detectors are slightly smaller
in size.
If a particle impacts only one of the A-detectors, the event
is discarded because the angle of impact and energy loss in
the other detector boards is not known. Also, any particle
entering the bottom of the telescope will not register an
event on the C-detector due to the directional properties
of the C-detector.
Component Parts
The chassis box of MARIE is made from machined aluminum with
an alodine coating. The exterior surfaces are painted white.
Input voltage to MARIE is 28 VDC and power requirements are
3 watts for survival mode and 7 watts for nominal operation.
There are no external controls. All control is from the orbiter
through an RS-422 interface.
Above: Exploded view of MARIE.
CPU & Power Boards
The Central Processing Unit (CPU) board has an Intel 80C188
microprocessor, detector interface circuitry and data communication
hardware for transferring data to the spacecraft from the
80 MB flash memory. The flash memory holds the program code
and any data which has not been transferred to the spacecraft.
The power from the spacecraft is nominally 28 volts. The Marie
instrument has Interpoint DC-DC converters to convert the
power to a usable level. Each detector has its own card, with
all of the electronics associated with the detector on it,
including a 12 bit analog-to-digital (ADC) converter, and
Field Programmable Gate Array (FPGA).
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A
Detectors
Each of the A detector assemblies contains a 25.4 x
25.4 x 1 mm ion-implanted silicon solid state detector,
detector signal amplifiers, detector high voltage supply
and the interface circuitry between the detector and
the MARIE CPU. The MARIE CPU controls the interface
circuitry including high voltage control, collecting
digitized signal amplitude data and controlling signal
coincidence timing sources. The two A-detectors are
used to define a coincidence event. These detectors
are operated near 160 V.
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Image above
right: A Detector |
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B
Detectors
Each of the B-detector
assemblies contains a 5 mm thick lithium-drifted silicon
solid state detector, detector signal amplifiers, detector
high voltage supply and the interface circuitry between
the detector and the MARIE CPU. The MARIE CPU controls
the interface circuitry including high voltage control
and collecting digitized signal amplitude data. These
detectors are operated near 350 V.
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Image above:
B Detector |
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C Detector
The C detector consists of a Schot-glass Cherenkov detector
and a Hamamatsu photo multiplier tube (PMT). When a
charged particle with a velocity greater than the velocity
of light divided by the glass refractive index hits
the Cherenkov detector, the detector releases a photon
light burst proportional to the energy of the particle
which struck it. The photo multiplier tube receives
the light pulse and translates it into an electronic
pulse which is amplified by the tube and read by the
electronics on the C-detector board. The C-detector
assembly contains the PMT, detector signal amplifiers,
detector high voltage supply, and the interface circuitry
between the detector and the MARIE CPU. The MARIE CPU
controls the interface circuitry including high voltage
control and collecting digitized signal amplitude data.
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Image above:
C Detector |
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Position Sensitive
Detectors
Each of the position sensitive detector (PSD) assemblies
contains a 25.4 X 25.4 mm position sensitive detector
each with a 24 x 24 wire grid to define the incident
direction of the charged particle, detector signal amplifiers,
detector high voltage supply, and the interface circuitry
between the detector and the MARIE CPU. The MARIE CPU
controls the interface circuitry including high voltage
control and collecting digitized signal amplitude data.
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Image
above: Position Sensitive Detector |
Testing
Certifying the MARIE instrument for flight involves subjecting
the instrument to environments the instrument is expected
to see during the Mars mission. The instrument is also tested
for workmanship during the environmental testing. This process
is used to verify the instrument will work in the correct
environment and that there are no know latent defects. The
following is a list of tests used to certify the MARIE instrument.
EMI Testing (May 2000)
EMI (Electro Magnetic Interference) testing verifies the EMI
compatibility of the MARIE instrument with the other hardware
on the Odyssey spacecraft. This includes both conducted and
emitted interference.
 Vibration
Testing (April 2000)
Vibration testing is performed to verify workmanship and to
simulate the expected launch environment. The instrument is
attached to table(s) which vibrate the instrument side to
side as well as up and down.
Image right: Vibration testing.
Thermal-Vacuum Testing (April 2000)
During thermal-vacuum testing, the instrument is placed into
a special testing chamber. The unit is cycled between a high
and a low temperature at a nearly perfect vacuum. This is
done to simulate the thermal environment that is expected
during the mission. In this case, 1.5 cycles were performed
between -55°C and +50°C.
 Pyro-Shock
Testing (May 2000)
Pyrotechnic devices are used to separate rocket stages and
to detach the aerobrake shield on the orbiter. This test simulates
the shock to the instrument when the pyrotechnics fire.
Image right: Vibration testing.
Burn-In (March & April 2000)
The instrument is powered on and the environment controlled
to ambient and an elevated temperature. This test is performed
to test workmanship and expose parts that will prematurely
fail. This test was conducted in March 2000 for 96 hours at
an elevated temperature of 35°C. The ambient temperature test
was performed in April 2000 for 203 hours.
Thermal Cycle Testing (April 2000)
During this workmanship test the instrument is cycled between
a high and low temperature. The unit was tested for 1.5 cycles
between +7°C and -49°C.
Outgas Testing
This test ensures that the instrument will not release any
harmful materials when subjected to a vacuum.
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