2. TECHNICAL APPROACH
The IMAGE mission will address its magnetospheric imaging objectives in unique ways using existing imaging techniques: neutral atom imaging (NAI) over an energy range from 10 eV to 200 keV, far ultraviolet imaging (FUV) at 121-160 nm, extreme ultraviolet imaging (EUV) at 30.4 nm, and radio plasma imaging (RPI) over the density range from 0.1 to 105 cm3 throughout the magnetosphere. Science objectives developed in the Step-1 proposal result in the instrument requirements listed in Table 2.1. A minimum baseline mission, below which IMAGE would not be worth pursuing at the proposed cost, has been identified to include two of the three NAI instruments (MENA and HENA), the EUV instrument, and one of the two FUV instruments (either the spectrographic imager or the wideband imaging camera). With this minimum complement, IMAGE would accomplish the complete set of science objectives that have been identified for the core Magnetosphere Imager mission [Armstrong and Johnson, 1995]. Details of the proposed descope process are contained in Volume II.
Imager Measurement Critical Measurement Requirements
NAI Neutral atom FOV: 90 x 90 degrees (image ring current at
composition and apogee). Angular Resolution: 8 x 8 deg. Energy
energy-resolved Resolution (E/E): 0.8 (trade resolution for
images over three sensitivity). Composition: distinguish H, He and O
energy ranges: in magnetospheric and ionospheric sources,
10-300 eV interstellar neutrals and solar wind. Image Time:
(LENA) 1-30 5 minutes (resolve substorm development).
keV (MENA) Sensitivity: effective area 1 cm2 for each sensor.
10-200 keV (HENA)
EUV 30.4 nm imaging FOV: 90 x 90 degrees (image plasmasphere from
of plasmasphere apogee). Spatial Resolution: 0.1 Earth radius from
He+ column apogee. Image Time: several minutes to hours
densities. (resolve plasmaspheric processes).
FUV Far ultraviolet FOV: 16 degree (image full Earth from apogee).
imaging of the Spatial Resolution: 60 km. Spectral Resolution:
geocorona and the separate cold geocorona H from hot proton
aurora at precipitation ([[Delta]][[lambda]]~0.2 nm near
[[lambda]] = 121.6 nm); separate 130.4 nm and 135.6 nm electron
134-160 nm aurora emissions. Image Time: 2-5 minutes (resolve
(Wideband Imaging auroral activity).
Camera, WIC) and
[[lambda]] =
121.6, 130.4, and
135.6 nm
(Spectrographic
Imager, SI)
RPI Remote sensing of Density range: 0.1-105 cm-3 (determine electron
electron density from inner plasmasphere to magnetopause).
densities and Spatial resolution: 500 km (resolve density
magnetospheric structures at the magnetopause and plasmapause).
boundary Image Time: 1 minute (resolve changes in boundary
locations using locations).
radio sounding.
Table 2.1. IMAGE measurement requirements are achievable with
existing imaging techniques.
The IMAGE mission is designed to have very low risk because it uses (1) science instruments with a high level of technical readiness, with all subsystems having been either flight-proven or verified with extensive ground-based and laboratory testing; (2) a direct application of the FAST SMEX spacecraft; and (3) the existing Goddard Space Flight Center (GSFC) Science and Mission Operations Center (SMOC) with very simple on-orbit operations.
2.a. Mission Design
A full design study using an existing spacecraft architecture (FAST)
was performed by GSFC (Code 700). This analysis concluded that an adapted
FAST architecture with the instrument complement in Table 2.1. will meet
all IMAGE requirements within the MIDEX cost and spacecraft resource
limits. This mission design provides the highest possible science per unit cost
by making extensive use of existing flight and ground systems, software, and
procedures. Flight-proven hardware designs are used wherever possible,
while laboratory-proven new technologies enhance the science return.
Operational complexities such as stringent spacecraft pointing requirements
and narrow launch and operational windows have been eliminated.
2.a.1. Orbital Requirements. A full analysis of the required IMAGE
orbit has been performed by the GSFC Flight Dynamics Division. The orbital
requirements for the IMAGE mission are shown in Fig. 2.a.1. The initial
argument of perigee is 320 degrees, placing the initial apogee at 40 degrees
north latitude. The apogee will drift to be over the north pole after one year
and back to ~40 degrees north latitude at the end of the two-year mission. A
0 or 180 degrees right ascension of the ascending node (RAAN) places
the spacecraft spin axis at 23.5 degrees to the ecliptic plane. This
orientation minimizes the sun-angle variation, optimizing solar input to the
body-mounted solar panels.
The basic Med-Lite launch vehicle (Orbital Sciences Corporation Taurus
vehicle with a Star 37FM upper stage, two Castor IVB strap-on motors, and a
2.3-m fairing) will place a 373 kg payload into the required orbit. The IMAGE
spacecraft mass of 291.5 kg provides a substantial mass margin of 81.5 kg. The
Delta-Lite launch vehicle will also meet the IMAGE orbit requirements
because it provides an even greater mass-to-orbit capability.
An orbit error analysis performed by McDonnell Douglas for IMAGE
estimates 3 errors of 9 km on perigee and 1700 km on apogee. These small
dispersions allow IMAGE to be placed into the required orbit without the
need for an onboard propulsion system.
Fig. 2.a.1. The IMAGE orbit maximizes observation time over the
north pole during the 2-year mission and provides ample ground station
coverage. A downlink time of only 13.5 min./day is needed.
2.a.2. Launch and Operational Windows. The IMAGE spacecraft can
be launched any day of the year to attain the required orbit. The launch
window is constrained only by the time of day, with a minimum launch
window of one hour per day. For communications there is ample ground
station coverage for DSN northern hemisphere ground stations with contact
times of about 8 hours per 13.5 hour orbit using a DSN 34-m antenna (see Fig.
2.a.1.). However, a downlink time of only 13.5 min/day is needed. The
smaller 26-m DSN antenna also can be used owing to the considerable margin
in the downlink window. This smaller DSN antenna will necessarily require
a slower telemetry rate and a corresponding increase in the data downlink
interval to ~25 minutes.
2.a.3. Mission Lifetime. IMAGE is a two-year mission (see Fig. 2.a.2.).
Fig. 2.a.2. Natural evolution of perigee altitude for the IMAGE
mission results in the main mission being accomplished with perigees above
600 km, giving a safe margin of stability even for higher than expected solar
activity.
2.a.4. Communication Requirements. The IMAGE data will be
transmitted as packets in a store-and-dump mode except for real-time
transmissions during launch and early orbit (when instrument and spacecraft
checkout is performed). The telemetry rates and volumes will be sized to
allow automated ground operations. The telemetry protocol and formats will
follow the Consultative Committee for Space Data Systems (CCSDS)
standards for cost-effective use of existing equipment. The DSN is the baseline
tracking network. Schedulingof DSN will be relatively simple because of the
short downlink time relative to the length of time the spacecraft will be
visible to the DSN (see Fig. 2.a.1.).
Based on the IMAGE telemetry rate with 1.8:1 instrument data compression
(Section 2.c.3.), a downlink rate of 2.2 Mb/s, and a DSN 34-m antenna, the data
downlink time from the onboard solid-state memory is approximately 13.5
minutes per day (bit error rate 10-5). The 2.0 Gbit solid-state memory will
have the ability to store two full orbits (27 hours) worth of data. All rates and
times given include a 20% margin on total data.
2.a.5. Orbit Determination. Orbit determination will be performed
once per orbit in the GSFC SMOC by the GSFC Flight Dynamics Division.
Spacecraft position knowledge (above ~2 RE altitude) is required to within 50
km after the fact to determine the geographic location of individual pixels in
the auroral images.
2.a.6. Attitude Determination. Spacecraft attitude determination will
be performed on board by the IMAGE attitude control system (ACS) to within
the accuracy and stability listed in Table 2.a.1. Attitude determination will also
be performed on the ground using the FASTRAD system.
2.a.7. Spacecraft Pointing Requirements. Table 2.a.1. summarizes the
IMAGE spacecraft pointing requirements. Once the spin axis is fixed, the RPI
antennas are deployed, and the final spin rate is attained, no special
spacecraft maneuvers are required.
2.a.8. Mission Operations (MO) Scenario. Launch and Early
Orbit.
After launch and vehicle separation, the ACS will align the spacecraft spin
axis perpendicular to the orbit plane and maintain a spin rate of 0.5 rpm.
Antenna deployment (identical to that of the FAST spacecraft) will occur in
the
Early Orbit phase (Figure 2.a.2.). During the ~25% of the orbit centered around
perigee, the antennas are slowly deployed while two spin torque rods
maintain the spin rate. A third torque rod negates spin rod precession and
maintains the spin-axis attitude, and a fluid damper reduces spacecraft
nutation. Analysis of the IMAGE spacecraft moment of inertia and perigee
magnetic field indicates that the antennas will be fully deployed in ~40 days.
This deployment period allows sufficient time for normal instrument high-
voltage turn-on and check-out.
Main Mission. The main mission starts after the antennas are
deployed (Figure 2.a.2.). Instruments will remain fully powered and will
operate autonomously throughout this phase. Most instruments have only
one science acquisition mode, and those that have more than one will still
operate autonomously with occasional preprogrammed mode changes
uplinked during data acquisition intervals. Only one contact per week will be
required for spacecraft commanding. Instrument health and safety checks will
be automated with data derived from the normal downlink.
2.b. Instrumentation
The IMAGE instrumentation is summarized in Table 2.1. above. The various
imagers communicate with the spacecraft through a Central Instrument Data
Processor (CIDP), making them appear as one instrument for command, data,
and electrical power purposes.
Table 2.b.1. NAI is accomplished with three sensors covering low,
medium, and high energies.
2.b.1. Neutral Atom Imagers (NAI). Science requirements driving the
NAI instrumentation for IMAGE are (1) to image the inner magnetosphere
including the ring current on a time scale of 300 s and (2) to resolve the major
species contributing to neutral atom fluxes. To meet these requirements a
suite of three NAI instruments will provide angle-, energy-, and
composition- resolved images at energies from 10 eV to 200 keV. Three
separate instruments are needed
because the physical processes used to convert neutral atoms to ions and then
detect the ions differ over the three energy regimes (0.01-0.3 keV, 1-30 keV,
10-200 keV). All three sensors are energy/mass/angle spectrographs, which
display the complete range of energy, mass, and polar angle simultaneously
and
do not require scanning high voltages. Sensor and electronics designs and
technologies for all 3 instruments draw heavily on Polar (TIDE, TIMAS, and
SEPS) and Cassini (MIMI and CAPS), as summarized in Section 2.b.5. Energy,
angle, and mass resolutions are kept at the minimum needed to deconvolve
anticipated particle populations whose composition (primarily H and O) and
charge-exchange cross sections are generally well-known. All three
instruments
use time-of-flight and energy/charge or total energy analysis to obtain
composition information. Angular information is obtained over 90 degree
fans with uniform 8 x 8 degrees resolution. Spacecraft spin is used to obtain
angular information in the orthogonal (azimuthal) direction. All three
instruments have collimators that consist of serrated, blackened surfaces to
reduce internal scattering. The collimators contain deflection potentials of 10
kV that deflect and absorb charged particles below 100 keV/e. Small broom
magnets remove electrons with energies <200 keV. All NAI sensors
contain Command and Data Interfaces, which control potentials on all high-
voltage supplies, accept logical pulses and data from the MCP and SSD units,
receive spin phase information from the CIDP, and accumulate events in
histogramming memory, which
is programmable to support calibration as well as flight operations.
High-voltage supplies and time-of-flight (TOF) electronics for all three
sensors have a common heritage based on Cassini/MIMI/CHEMS. Table 2.b.1.
summarizes NAI performance, while Section 2.b.5. summarizes NAI
resource requirements.
An important part of the NAI approach is the team's ability to calibrate the
sensors accurately. Facilities already exist at SwRI, GSFC, LPARL, and
University of Denver to achieve accurate absolute and inter-sensor
calibration.
2.b.1.a. Low Energy Neutral Atom [LENA] Imager.
We have examined in detail the merits of three possible techniques for
detecting neutrals in the energy range below several hundred eV. We find
that the only viable technique is surface conversion.
Fig. 2.b.1. LENA is a spherical electrostatic analyzer with TOF
mass analysis and a Cs conversion surface that converts neutral atoms to
negative ions.
Optics and Detectors. LENA [Ghielmetti et al., 1994] combines
flight-proven plasma analyzer techniques with well-established, high-
efficiency cesiated neutral-to-negative-ion surface conversion technology to
measure composition and energy spectra of neutral atoms (H, D,
3He, 4He, and O) at 10-300 eV. LENA consists of a
collimator, conversion unit, extraction lens and acceleration region,
dispersive electrostatic energy analyzer, and TOF mass analyzer with position-
sensitive particle detection (Fig. 2.b.1.). LENA's high geometric factor (0.2
cm2 sr) is achieved by simultaneously imaging in azimuthal
angle, energy, and mass/charge (i. e., the spectrograph approach).
Particles enter the instrument through a collimator with elevation acceptance
defined by the heights of the collimator and slit 1. A shutter located behind
slit 1 will be closed to protect the conversion surface during launch and
perigee portions of the orbit. Neutrals are converted to negative ions through
near specular reflection from a low-work-function cesiated tungsten
conversion surface [Wurz et al., 1995]. The surface is segmented to
cover a 90 degrees azimuthal acceptance, with the facets aligned on a conical
surface centered about the instrument axis of rotational symmetry (through
slit 1). Laboratory tests of a prototype LENA have confirmed operation of the
conversion surface, acceleration region, and electrostatic analyzer.
Laboratory testing of conversion surfaces at pressures significantly poorer
than will be achieved on orbit has shown ionization efficiencies of 10-15%
[Wurz et al., 1995]. Contamination of the surface will, in time, reduce
conversion efficiency. Because there is a direct relationship between work
function and conversion efficiency, in-situ monitoring of the work function
gives conversion efficiency at any time. This monitoring is accomplished
using
laser diodes and the photoelectric effect at three different wavelengths. Based
on laboratory data, we conservatively estimate that surface regeneration will
be required at 10-day intervals. During regeneration, the instrument shutter is
closed, high voltages turned off, and the surface heated to drive out
impurities. The Cs dispenser, containing a non-volatile Cs salt, is switched on
to deposit the single monolayer required. To avoid contamination within the
rest of the instrument, the Cs generator dispenses a collimated beam to the
surface. All nearby conductors are baffled.
Beyond the conversion surface, LENA is an ion mass spectrograph. Neutrals
converted to negative ions at the Cs surface are accelerated away from the
conversion surface and focused in the plane of slit 2. The negative ions are
further accelerated to 20 keV between the object slit 2 and the spherical
electrostatic analyzer. The electrostatic analyzer (operating at a fixed
voltage) is dispersive in energy and focusing in elevation angle, so that it
images the entire slit 2 onto the carbon foil of the TOF section.
The mass analyzer is similar to that of TIDE on Polar [Moore et
al., 1995]. Secondary electrons produced on the back side of the foil are
accelerated, deflected, and focused onto one MCP, producing the "start" pulse
for the TOF and providing energy and polar angle information. The signal
produced at the other MCP provides the "stop" pulse. The start and stop
signals are obtained from 50%-transmission grids placed behind the
respective MCPs. The position and TOF electronics are contained within a
high-voltage bubble. The encoded information is transmitted across the high-
voltage interface via fiber optics.
The three main sources of instrument background are UV photons,
photoelectrons from the conversion surface, and negative ions produced by
attachment of photoelectrons to residual gas molecules. UV photons scattered
at near specular angles from the conversion surface enter a light trap at the
back, while the electrostatic analyzer provides further multiple bounce
rejection. Photoelectrons from the conversion surface are separated from the
negative ions and diverted into a trap by a weak magnetic field between a
small magnet at the back of the conversion surface and another at the
dispenser. This magnetic field also traps electrons scattered through the
collimator. Ion background from residual gas is minimized by accelerating
photoelectrons away from the sensitive region. Any negative ions that are
thereby produced will be discriminated against by their lower energies
relative to those produced on the surface.
LENA has considerably lower mechanical tolerances and also lower detector
position resolution than required by the Polar TIDE and TIMAS instruments,
which are its heritage. A further simplification is that the LENA sensor and
detector high-voltage power supplies operate at a fixed voltage, eliminating
the need for high-voltage sweeping. The only portion of LENA without direct
flight heritage is the conversion surface, with which we have extensive and
ongoing laboratory experience [Wurz et al., 1995].
Sensor Electronics. LENA contains a single static 20-kV high-voltage
supply from which all other ion deflection voltages are derived using taps on
the high-voltage multiplier. The MCPs use a separate supply. Four energy
levels and 11 angles are coded into a pixel array. Start/stop events are
measured in a time-to-amplitude converter, digitized, and then recorded and
tagged with pixel location. Events are binned in cadence with spacecraft spin
to produce 8 x 8 degrees angular pixels.
Operation. LENA has one operation mode and one surface
regeneration mode. In a typical orbit, the shutter will open after a perigee
pass, and low-energy neutral atom imaging will proceed throughout the orbit
until near perigee, when the shutter will close again. Approximately every 10
days the instrument will be cycled autonomously through the conversion
surface regeneration mode described above.
Calibration. Instrument calibration prior to launch will be performed
at GSFC and at the University of Denver neutral beam facilities using H and
O neutral beams.
2.b.1.b. Medium Energy Neutral Atom [MENA] Imager. The
MENA imager relies on thin carbon foils to convert neutral atoms into ions
[cf. McComas et al., 1994]. However, while McComas et al. describe an
energy spectrometer (which allows only one energy to be detected at a time),
MENA is an energy-angle-mass spectrograph that samples and resolves
simultaneously all energies, all polar angles within a 90 degree fan, and all
masses. This spectrographic feature is very important for magnetospheric
imaging because the weak fluxes of neutral atoms require an instrument with
a high duty cycle.
Optics and Detectors. The MENA analyzer (Fig. 2.b.2.) consists of
elevation and azimuthal collimators, stripper foil, TOF analyzer [Young et
al., 1992; Moore et al., 1995], and a cylindrically symmetric (with
axis of symmetry lying in the page) electrostatic mirror spectrograph with
well-known properties [Sar-El, 1967]. Collimator plates (described
previously) reduce charged particle background and internal scattering. The
other significant background that must be eliminated is ultraviolet light from
the Sun and geocorona. A combination of a stripper foil and -5 kV post
acceleration is used (1) to create positive ions that are deflected away from
the UV line-of-sight and (2) to reject secondary photoelectrons. When
transmission through the two foils, anti-reflection blackening of grids, and
MCP efficiency for UV are taken into account, the MENA UV rejection rate is
10-9.
Fig. 2.b.2.a. MENA is an
energy-angle-mass spectrograph that samples and resolves simultaneously all
energies, all masses, and all polar angles within 90 degrees. This elevation
view shows the ray-traced trajectories of ions created from neutrals with
incident energies of 1 and 30 keV.
Ray paths are shown in Fig. 2.b.2.a. for incident neutral atoms with incident
energies of 1 and 30 keV and in Fig. 2.b.2.b. for 30-keV atoms. The ions
produced by the stripper foil are accelerated by -5 kV in the adjacent
acceleration region. The cylindrical electric field forms an image in energy
and polar angle of the 8 x 8 degrees collimated beam. The carbon foils are
nominally 0.5 g/cm2 thick and of the same type and thickness
used in our other TOF instruments [Young et al., 1992; Moore et
al., 1995]. Energy-per-unit-charge and simultaneous velocity analysis
(through TOF measurements made with the start/stop MCP) yield mass per
unit charge with a resolution sufficient to separate H and O.
Fig. 2.b.2.b. Front view of
MENA analyzer with trajectories of 30 keV atoms at three different polar
entrance angles shows 90 degrees FOV.
Secondary electrons emerge from the back of the start TOF carbon foil placed
above the MCP and are electrostatically accelerated down onto the MCP,
which functions as both a start and stop detector (Fig. 2.b.2.a.). This
arrangement is feasible because the ion travel times (5 to 300 ns) are short
compared to the mean time between events (104 ns). The MCP records a start
pulse for TOF analysis and also gives position for energy and polar angle
identification. The stop pulse is produced by the same MCP to provide timing
information. Neutral atoms, which are not deflected by the electric field, pass
through a second cylindrical grid and are detected with the MCP shown near
the top of the figures. Negative ions, which also emerge from the stripper foil
in some fraction depending on atomic species and energy, will also be
detected by the upper MCP. Thus, even though the analyzer is designed
primarily for positive ions, particles that emerge as neutrals or negatives can
also be detected and analyzed to obtain nearly 100% of the total incident
neutral flux (the exact fraction is obtained from calibration). This feature gives
MENA a key advantage over other designs that analyze only positive or
negative ions [e.g., McComas et al., 1994]. A laboratory prototype of the
MENA analyzer has been built and successfully tested at SwRI.
Sensor Electronics. MENA contains a static +20 kV deflection supply
from which positive analyzer voltages are derived. A -5 kV potential derived
from the -10 kV collimator supply is used to accelerate positive ions leaving
the stripper foil at the entrance to the cylindrical analyzer. The neutral-atom
and positive-ion MCPs are powered from two separate supplies. MENA
detectors consist of two rectangular sections cut from standard Hammamatsu
"Z-stack" MCPs. Four energy levels and 11 polar angles are coded into a pixel
array. Start/stop events are measured in a time-to-amplitude converter,
digitized, and then recorded and tagged with pixel location. Events are binned
in cadence with spacecraft spin to produce images with 8 x 8 degrees angular
resolution.
Operation. All MENA high voltages are static except for infrequent
adjustments of MCP gain. After initial switch-on in orbit, MENA is simply
allowed to run in a single operational mode. A monthly calibration mode is
run to check MCP gain saturation, and pulsers are used to generate positional
patterns that test the detector signal chains and encoding logic.
Calibration. The SwRI Ion/Electron/Neutral-Atom calibration facility
will generate neutral atom beams of hydrogen and oxygen in the range 1 to 30
keV that will be used to test and calibrate MENA. The facility's
duoplasmatron source provides a collimated neutral flux equivalent to 10-14
Amp/cm2 (any larger current would saturate the MENA
detection system). The facility includes a 4 degree-of-freedom goniometer
capable of holding MENA-sized instruments.
2.b.1.c. High Energy Neutral Atom (HENA) Imager. At
higher energies (>10 keV for H and >40 keV for O), individual neutral
atoms can be directly detected with good energy resolution by low-noise solid
state detectors (SSDs). Each neutral atom's mass (M) and energy (E) are
identified unambiguously by the proven technique of measuring the speed by
time-of-flight (TOF) and residual energy by pulse-height (PH)a method first
published by Gloeckler and Hsieh [1979] and later employed in
various configurations in AMPTE/CCE and GGS/WIND, and to be flown on
Cassini/MIMI. Advanced imaging capability is achieved by using a 1-D strip-
SSD, each strip having a field-of-view (FOV) defined by the ion-rejecting
collimator to represent a polar ([[theta]]) sector of arrival for a given
azimuthal ([[phi]]) sector defined by the spacecraft spin. Thus, the detected
atoms are grouped into a matrix (M, E, [[theta]], [[phi]]), forming the basic data
base. HENA has two identical sensor heads, each having two identical
TOF/PH analyzers.
Optics and Detectors. HENA acquires angular images using a multi-
channel fan-shaped collimator that also serves to suppress charged particle
entry by biasing adjacent collimating plates at +/- 10 kV. Fig. 2.b.3. shows the
collimator and TOF/PH analyzer in half of a sensor head. A 1-D strip-SSD (20
strips per analyzer) over-samples the 10 angular channels of the collimator,
forming a 2D image in ([[theta]], [[phi]]) when binned by spin azimuth angle.
Collimator channels and strip-SSDs define 8 degrees [[theta]]-sectors over the
90deg. [[theta]]-FOV for each 8deg. [[phi]]-sector. A 5 ug cm-2
Si/polymead/C foil was chosen to reduce scattering and is possible because of
the immunity of the strip-SSDs to EUV. Secondary electrons generated on the
back side of the foil (Fig. 2.b.3.) are guided to the Start half of the MCP by
a carefully designed electric field. Symmetry of this field allows secondary
electrons generated on the surface of the SSD to be guided by the same electric
field to the Stop half of the MCP. The short TOF path (2.3 cm) allows
sufficient resolution (2 ns between 4 and 65 ns) for separating H from O in the
TOF/PH analysis and effectively reduces accidental coincidence rates. The
MCPs are rectangular "Z stacks" of three 100 mm 15 mm plates to provide
ample gain reserve. With its 1.6 cm2 sr geometrical factor and
overall sensitivity of 0.58 for TOF only and 0.41 for TOF/PH, simulations
indicate that HENA will gather ~3 events/pixel/minute when viewing the
quiet ring current and~300 events/pixel/minute during storm main phase.
Fig. 2.b.3. Half of a HENA sensor head (top) and schematic of the
TOF/PHA detector (bottom).
Sensor Electronics. The HENA TOF electronics are essentially
identical to those used on Cassini/MIMI, which in turn are based on earlier
experience with AMPTE/CCE and GGS/WIND. The SSD electronics utilize
the flight-proven AMPTEK A250 hybrid charge-sensitive amplifiers (44 per
sensor head) followed by shaping and logic based on the microelectronics of
Polar/SEPS [Blake et al., 1995]. Pre-amplification is done within the
sensor heads, while TOF and PH analysis are performed on boards within the
rectangular box sandwiched between the heads.
Operation. All HENA potentials are static except for infrequent
adjustments of MCP gain. After activation and checkout, HENA runs in a
single operational mode. Some choices of data product priorities can be made
as software options, providing, for example, a periodic (e.g., monthly)
calibration mode to check the TOF/PH identification of atomic species, the
pulse-amplifier chains, and the binning logic.
Calibration. HENA will be calibrated using ion accelerators at GSFC
and Taylor University and a solar-constant UV source at MSFC. The ion
accelerators provide a small collimated neutral atom flux in the presence of a
much larger ion flux, which is a good simulation of conditions in which
HENA will operate.
2.b.2. Photon Imagers. There are three photon imagers on IMAGE.
The EUV instrument images resonantly scattered solar emissions from
plasmaspheric He+ at 30.4 nm. Two FUV instruments, SI
(Spectrographic Imager) and WIC (Wideband Imaging Camera) image the
Earth's auroral emissions in several FUV wavelengths. General
characteristics of the imagers are listed in Table 2.b.2.
2.b.2.a. He+ 30.4 nm Imager (EUV). Effective
imaging of plasmaspheric He+ requires global "snapshots" in
which the high apogee of the IMAGE mission and the wide FOV of the EUV
imager provide, in a single exposure, a map of the entire plasmasphere from
the outside with a sensitivity of 0.2 count/s-pixel-Rayleigh (R), a spatial
resolution of 0.1 RE, and a time resolution of several minutes. The 30.4-nm
feature is easy to measure because it is the brightest ion emission from the
plasmasphere, it is spectrally isolated, and the background is negligible.
Measurements are easy to interpret because the plasmaspheric
He+ emission is optically thin, so its brightness is directly
proportional to the He+ column abundance. Since the 1970's,
EUV photometers in LEO have produced "inside-out"images of the
plasmasphere with a brightness of ~10 R.
Fig. 2.b.4. Schematic of one of the three identical EUV
(He+ 30.4 nm) imager sensor heads. Internal baffling is not
shown. The design is that of EUVI, a He+ imager that will fly on
the Shuttle in July 1997.
EUV consists of three identical sensor heads (Fig. 2.b.4.) serviced by a
common electronics module. It employs elements of new technology,
multilayer mirrors. Because it is simple and lacks moving parts, the EUV is
rugged and reliable. Each sensor head has a field of view of 30 x 30 degrees.
The three sensors are tilted relative to one another to cover a fan-shaped
instantaneous FOV of 90 x 30 degrees. As the satellite spins, the fan sweeps a
90 x 360 degrees swath across the sky.
Optics and Detector. Each EUV sensor achieves high throughput and
a wide field of view by using a large entrance aperture and a single spherical
mirror. A multilayer reflective coating on the mirror selects a narrow 5-nm
passband around the 30.4 nm line. To circumvent the red leak in the
multilayer mirror, the filter blocks H Lyman [[alpha]]
contamination from the geocorona. The detector consists of two curved,
tandem MCPs with an alkali halide front
surface photocathode. The detector's spherical input surface minimizes the
effects of spherical aberration. Readout from the detector is from a 128 128
wedge and strip anode. The sensitivity (accounting for the duty cycle inherent
in a spinning spacecraft) is 0.2 count/(sec pixel) per Rayleigh, where the
pixel size is taken to be 0.1 RE. By summing pixels to make a spatial
resolution element (called a resel) of 0.5 RE, the count rate is
5 counts/(sec resel) per R.
Electronics. Each sensor has its own preamplifiers and high- voltage
power supply. Signals from the preamps are processed by position- finding
circuitry in the central electronics module, which also includes the control
microprocessor (CMP), a RAM buffer, and A/D converters for housekeeping
information. The CMP accepts commands from the CIDP to select operating
modes, spatial resolution, etc. The satellite provides a spin-phase sync signal
to the EUV.
Operation. A key element in the operation of the EUV is the RAM
buffer, which stores a 2-D array of brightness measurements. Each element in
the array corresponds to a particular spatial resolution element in the sky.
EUV keeps track of the satellite's spin phase, so the CMP can relate the
position of each detected photon to an absolute position in the sky. It then
increments the value of the corresponding array element. After a number of
spins corresponding to the desired time resolution, the CIDP commands a
read-out of the map buffer, compresses the image, and formats it for the
telemetry stream. EUV's data rate is 1.5 kbps for about 90% of an orbit and
increases to 4.5 kbps near perigee.
When the satellite is near apogee, the plasmasphere will fill about 30% of the
sky. The range of the map buffer and the angular size of a resel will be set
appropriately by autonomous command. When the satellite is near perigee,
plasmaspheric emission will fill much of the sky. Then the range of the map
will be increased, and the angular resolution will be reduced.
For some orientations of the spin axis, the Sun may enter the FOV of one of
the
sensor heads at some spin phases. No useful data can be recorded from the
affected sensor head with the Sun in the field, and the EUV control
microprocessor will automatically reduce the high voltage to the MCPs to
avoid
excessive counting rates. The filter will prevent damage to the detector by
focused visible sunlight.
Heritage. The sensor heads are similar to the soft x-ray telescopes
developed for the Earth-orbiting satellite Alexis. The multilayer coating
techniques are now quite mature. The detector is similar to others used
routinely, including those developed by the University of Arizona IMAGE
team
members. We are presently fabricating the EUVI, a He+ imager of the design
proposed here, that will fly on the Shuttle in July 1997. Test results will be
available prior to the IMAGE instrument fabrication phase.
Calibration. The EUV will be calibrated at the University of Arizona
Lunar and Planetary Laboratory (LPL) EUV Calibration Facility. We will
calibrate the sensor heads separately, by mounting each head in a gimbal that
permits rotation about two orthogonal axes and translation along two
orthogonal
axes. With the entrance aperture illuminated by a collimated beam of 30.4-nm
radiation of known intensity, we will record the response of the EUV as the
head is rotated to probe the full field of view.
2.b.2.b. Far Ultraviolet Imagers (FUV). Science requirements
driving FUV imager designs are (1) to image the entire auroral oval from a
spinning spacecraft at 7 RE apogee altitude, (2) to separate spectrally the hot
proton precipitation from the statistical noise of the intense, cold geocorona,
and (3) to separate spectrally the electron and proton auroras. FUV consists of
two imagers that combine high spectral discrimination, high spatial
resolution, and the greatest possible sensitivity to meet these requirements.
In the FUV range up to ~160 nm, there are several bright auroral emission
features that compete with the dayglow emissions. For the electron aurora,
the brightest is 130.4 nm OI, which is multiply scattered in the atmosphere
and thus cannot be used for auroral morphology studies. The next brightest is
the 135.6 nm OI emission [Strickland and Anderson, 1983]. Separation
of the 130.4 and 135.6 nm lines necessitates the use of a spectrometer because
even reflective narrow-band filter technology cannot satisfy the ~3 nm
wavelength resolution requirement. Above 135.6 nm, weak LBH lines can be
detected using narrow-band filter technology. Separate imaging of the intense,
cold geocorona (Lyman [[alpha]] emissions at 121.6 nm) and the less intense,
Doppler-shifted Lyman [[alpha]] auroral emissions requires significantly
higher spectral resolution (0.2 nm).
Fig. 2.b.4.a. The IMAGE Spectrographic Imager (SI) is actually two
imagers in one: high spectral resolution of Lyman emissions is achieved
with slit 1; lower spectral resolution of LBH bands is achieved with slit 2.
Spectrographic Imager (SI). The relatively high wavelength
resolution requirement is satisfied by the SI. The 0.2-nm wavelength
resolution drives the size of the instrument and consequently the number of
mirrors in the optics system. Also the narrowness of the slits in the
spectrometer limit the dwell time during which a pixel is in the field of view.
Optics and Detector. The SI (Fig. 2.b.4.) is actually two spectrometers.
It has two closely spaced parallel slits located in a cut-out in the center of
the grating. The two slits are positioned parallel to the satellite spin axis
and scan the Earth once per rotation. The narrow slit (slit 1) provides a
wavelength resolution of 0.2 nm. The narrow width of this slit, combined
with the 0.1 nm spectral resolution of the system, allows the imager to
distinguish between, and separately measure, the central core of the Lyman
[[alpha]] emissions (geocorona) and the Doppler-shifted components (proton
aurora). The wider slit (slit 2) produces a wavelength resolution of 3 nm. This
slit is used to measure the OI 135.6 nm emission with adequate spectral
resolution to distinguish it from the brighter OI 130.4 nm emission and has a
blocking transmission filter to suppress the bright Lyman emission. There
are two regions of interest on the detector. The first region contains the
Lyman [[alpha]] (121.6 nm) emissions from the geocorona and proton aurora
through slit 1. The second region contains the 135.6 nm emissions from the
electron aurora through slit 2. This region actually consists of a variety of
electron aurora emissions: about 80% OI 135.6 nm, about 20% 135.4 nm LBH,
and also a negligible amount (~2%) of additional LBH at 140 nm that enters
through the narrow slit 1.
The detector uses a KBr photocathode on a MgF2 window image tube with
MCP intensification. The intensified image is detected by a crossed delay-line
type detector with two 32 x 128 pixel active areas. The narrowness of slit 1
assures that the detector is not saturated even with full nadir Lyman [[alpha]]
dayglow/geocoronal emission in the field of view.
Electronics. The detector preamps and high-voltage power supplies
are housed in the sensor box near the detector (Fig. 2.b.4.a.). Like the other
photon sensors, the SI detects the position of every photoelectron by
position-finding circuitry, and the images are produced in RAM memory by
integrating the detected pulses according to a look-up table that is dependent
on the rotation phase of the spacecraft.
Operation. The SI operation is very similar to that of EUV, except that
the RAM memory stores a 2-D array of brightness measurements for two
separate wavelengths (121.6 and 135.6 nm) instead of one. The SI has only one
mode of operation. For some orientations of the spin axis, the Sun may enter
the field of view of SI at some spin phases. As with EUV, the control
microprocessor will automatically reduce the high voltage to the MCPs to
avoid excessive counting rates. The filters will prevent damage to the detector
by focused visible sunlight.
Heritage. The SI design is derived from the University of Arizona
Glow
Spectrometer and the Dual Range FUV spectrograph in the Solar Plasma
Diagnostic
Experiment. A full ray tracing design has been performed for the system. The
detector design was developed jointly by UC Berkeley and Lockheed Palo Alto
Research Lab for the SOHO spacecraft.
Calibration. The calibration of all the UV instruments is relatively
conventional. All imagers will be calibrated for absolute sensitivity as a
function of wavelength in band and out of band. It is important to determine
the response of the instrument to natural sunlight as reflected from objects
such as white clouds. Also, every pixel will be measured in the calibration to
determine any geometric distortion as required for the flight software.
Facilities exist at all photon imager institutions to perform these
calibrations. For SI, the University of Liège will perform mechanical
alignment, Lockheed will perform detector and initial calibration, and GSFC
will perform final preflight calibration.
Fig. 2.b.4.b. The SI, shown here in a cutaway view, employs novel,
low-risk, fixed-optics technology to achieve dual-wavelength imaging.
Wideband imaging camera (WIC). The relatively high
sensitivity requirement for auroral imaging is satisfied by the WIC. This
imaging camera uses the basic design flown on the Viking and Freja satellites
(Figure 2.b.5.) to measure the auroral LBH emissions in a relatively
broad band from 134 nm to 160 nm [Anger et al., 1987]. The large field
of view permits a long dwell (or integration) period and increases the
apparent sensitivity.
Optics and Detector. The WIC optics design is identical to that of the
Freja camera. Incident photons pass through a filter that blocks the Lyman
[[alpha]] emissions and protects the detector from direct, focused sunlight
(Fig. 2.b.5.). The primary and secondary mirrors have a coating that is highly
reflective (>60%) in the FUV but has minimum (<3%) reflectance out of
band. The detector uses a CsI photocathode on a BaF2 window. MCPs are used
to
intensify the image, which is produced on a phosphor and fiber-optically
coupled to a diode array.
Electronics and Operation. Electronics and operation of the WIC are
essentially identical to those of SI. Readout occurs once per 0.1deg. of
rotation for a frame rate of 30 frames/s at 0.5 rpm. The camera data are
digitized and co-added in memory; the addresses are selected according to the
rotational phase of the spacecraft. This technique minimizes the distortion
correction required by the imager.
Heritage. The WIC is a direct copy of the Viking and Freja wideband
cameras with the CCD detector replaced by a diode array detector to minimize
radiation susceptibility. Such diode array detectors have been flown by
Lockheed in the DoD ERIS program.
Fig. 2.b.5. A copy of the WICs flown successfully on Viking and
Freja, the IMAGE WIC will provide high-sensitivity images of auroral LBH
emissions. A photo of the
wideband cameras on Freja.
Calibration. The WIC will be calibrated by the University of
Calgary
using methods similar to those described above for SI.
2.b.3. Radio Plasma Imager. The Radio Plasma
Imager (RPI) is a transmitter/receiver system that responds to the science
requirement for the continuous remote sensing of plasma densities,
structures and dynamics in the magnetosphere and plasmasphere. The
instrument measures the time delay, angle-of-arrival, and Doppler shift of
magnetospheric echoes over the frequency band from 3 kHz to 3 MHz. This
frequency range makes possible remote sensing of plasma densities from 0.1
to 105 cm-3. Programmable operational modes
selected within the limits listed in Tables 2.b.3. and 2.b.4. will focus on specific
magnetospheric and plasmaspheric features.
Table 2.b.4. RPI operational parameters.
RPI Configuration. RPI will have two crossed 400-m tip-to-tip thin
wire dipole antennas in the spin plane, and a 10-m tip-to-tip tubular dipole
antenna along the spin axis (Fig. 2.b.6.). All three antennas will be used for
reception to determine the angles of arrival of the echoes [Calvert et
al., 1995].
RPI Electronics. The RPI electronics will be housed together with the
CIDP in an electrically shielded box. RPI is a space-qualified adaptation of
the state-of-the-art Digisonde Portable Sounder (DPS) [Reinisch et al.,
1992, 1995; Reinisch, 1995]. Eighteen DPS units are operating
worldwide, and 15 more are currently being deployed. The transmitters use
solid-state MOSFET push-pull final amplifiers with ferrite core transformers.
The two spin-plane antennas are driven +/-90 degrees out-of-phase, resulting
in right/left-hand circularly or elliptically polarized signals. The signal
waveform (pulse width and repetition rate, intra- and inter-pulse phase
coding and wave polarization) is software controlled on a pulse-by-pulse
basis, providing the needed mode selection flexibility. The digitally
synthesized super-heterodyne receivers are derived from the DPS design.
Fig. 2.b.6. RPI is a straightforward spaceborne adaptation of the
Digisonde Portable Sounder system.
Signal Processing, Signal-to-Noise Ratio, and Spatial and Time
Resolution. The large distances, low power, and short antennas (relative
to the wavelength) require onboard signal processing. Pulse compression and
coherent spectral integration techniques developed for the DPS will be used to
achieve the required signal-to-noise (S/N) ratios. The calculated worst case
S/N before (right axis) and after (left axis) digital processing for echoes
from the magnetopause, plasmapause, and plasmasphere are shown in Fig.
2.b.7. for a typical digital processing gain of 21 dB. For a large part of the
frequency range, S/N is larger than 100, assuring 1 degree angular
resolution. The range resolution of 500 km is defined by the 3.3-ms width of
the transmitted pulses. The number of sounding frequencies selected for a
given measurement, together with the coherent integration time, determines
the time resolution, since measurements are taken continuously.
Fig. 2.b.7. Digital signal integration will allow remote sensing of
most of the magnetosphere [after Calvert et al., 1995].
2.b.4. Central Instrument Data Processor (CIDP). The CIDP
consists of a RAD-6000SC central processor, 250 MB solid-state memory with
telemetry interface, instrument communications board, RPI signal processing
electronics, and a DC/DC converter subsystem. All circuit boards are packaged
in 6U VME form factor. The CIDP is based on an existing flight computer
developed by SwRI for the CSAT commercial communications satellite (Fig.
2.b.8.) and under development for the International Space Station Alpha
(ISSA) Furnace Facility.
Employing the Loral Federal Systems RAD-6000SC 32-bit, radiation hardened,
reduced instruction set (RISC) CPU, the CIDP is capable of over 25 MIPS
performance at maximum clock frequency. Power consumption can be
reduced by lowering the CPU's clock frequencies by changing control registers
in software. With a total dose resistance of >106 rads (Si), a single event
upset LET of >80, and totally latchup-free operation, the RAD-6000SC is the
optimum control processor for IMAGE. The RAD-6000SC was developed for
the Mars Environmental Survey Mission and is fully flight qualified. With
the exception of the instrument communications board, all CIDP subsystems
are fully developed. The CIDP will be produced by the SwRI spacecraft
computer development group, an organization with over 30 successful
missions to its credit.
The CIDP provides the following support services to the IMAGE science
instruments: (1) instrument data acquisition, compression, and storage; (2)
onboard data processing; (3) uplink command processing, routing, and
scheduling; (4) formatting of stored data for telemetry downlink; (5) health
and status monitoring; and (6) communications interface to spacecraft
Mission Unique Electronics (MUE). In operation the CIDP will acquire
instrument data over RS-485 serial communications lines; compress, bin, or
sort the data as required for each instrument; and store the data in
compressed format in the solid-state memory. Orbit state-vector information
and time will be acquired from the spacecraft MUE and included with the
instrument data packets. Ground
commands will be passed to the CIDP from the spacecraft transponder. After
error checking, the command loads will be passed to the science instruments'
embedded microcontrollers. Commands can also be stored for later execution.
When enabled to do so by the spacecraft MUE, the CIDP will transfer the
contents of its solid state memory for downlink at 2.2 Mb/s.
Fig. 2.b.8. CIDP design will be based on the C&DH system for
the CSAT commercial communications satellite, shown here, and the ISSA
Furnace Facility processor now under development.
Software resident in the CIDP will consist of the VxWorks operating
system (produced by Wind River Systems, Inc.), I/O device drivers, and the
applications software used in instrument operations. This same operating
system is currently in use in two other RAD-6000-based flight computers
being built by SwRI. Applications software for both the microcontrollers and
the RAD-6000 will be written in ANSI Standard C. IMAGE software will be
developed in a controlled, structured, and well-documented manner with
emphasis on configuration control using the ISO 9001-3 software assurance
system.
2.b.5. Instrument Requirements Summary. A summary of the
resource requirements of the complete set of IMAGE instruments is shown in
Table 2.b.5.
2.c. IMAGE Spacecraft
The IMAGE mission will use a NASA-provided spacecraft based on the Small
Explorer (SMEX) Fast Auroral Snapshot Explorer (FAST). FAST is an existing
spacecraft of known cost. A full spacecraft design study completed at GSFC
shows that the FAST system, with a larger solar array and battery, can be
adapted to meet all IMAGE requirements within the MIDEX cost and resource
limits.
FAST was chosen as the baseline for IMAGE because both missions require a
spin-stabilized spacecraft with deployable wire booms and a body-mounted
solar array, and both must withstand a high radiation environment. The
IMAGE spacecraft diameter is 60% larger than FAST but will copy the simple
FAST architecture (Fig. 2.c.1.), all components being either exact copies of, or
based on, existing hardware. IMAGE will benefit from FAST and SMEX
systems already developed.
Specifically, IMAGE will make use of:
This design approach provides ample mass and power margins, while
minimizing risk and cost (Table 2.c.1.).
Fig. 2.c.1. IMAGE uses the FAST spacecraft architecture. (FAST
subsystems are shaded.)
2.c.1. Instrument Accommodation. All instruments mount around
the perimeter of a single deck (Fig. 2.c.2.). The NAI, EUV, and FUV
instruments have fan-shaped fields of view parallel to the spin axis and
centered on the spacecraft spin plane. The four RPI wire antenna deployers
are mounted 90 degrees apart, oriented so the deployed antennas are
perpendicular to the spin axis. The two RPI rigid axial antenna deployers are
mounted at the center of the deck, one each on the top and bottom, and
deploy along the spin axis.
Fig. 2.c.2. IMAGE instrument complement is easily
accommodated by an enlarged FAST-based spacecraft.
Openings in the spacecraft bellyband allow instrument viewing and antenna
deployment. Instrument fields of view are unobscured by the RPI antennas,
and fine alignment is not required. Sunshades, baffling, and protective covers
are integral to each sensor.
Table 2.b.5. IMAGE science payload resource requirements
summary.
2.c.2. Structure. The IMAGE structure (Fig. 2.c.3.) uses the FAST
concept. The aluminum structure mates to the vehicle with the spin axis in
the thrust direction. The design will be dynamically balanced and provides a
spin to lateral moment of inertia ratio of 1.37 prior to boom deployment,
which is greater than the 1.05 needed for spin stability.
2.c.3. Command and Data Handling (C&DH) and Communications
System.
The C&DH and electrical system (Fig. 2.c.1.) is based on the FAST
architecture and uses the FAST MUE. The MUE houses the spacecraft
computer, power system, and ACS electronics. Use of the MUE allows
extensive re-use of FAST software. Only two relatively minor changes to the
MUE are required for this mission: (1) a modification to drive three torquers
versus the two on FAST and (2) a modification to use two horizon-crossing
indicators versus one on FAST.
The C&DH system provides ample data storage and downlink capacities
(Table 2.c.2.). IMAGE will execute both real time and stored commands.
Execution of stored commands will be within 100 msec of onboard Mission
Elapsed Time (MET).The command load will be less than 30 kBytes with one
command load per week sufficient for normal operations. The IMAGE
instruments will use the timing capabilities provided by the MUE. The MUE
provides MET and a 1-Hz clock. The required MET resolution of 10 ms and
absolute accuracy of 5 s is within the MUE capability of 15-s resolution and 3 s
accuracy.
The IMAGE communications system uses the SMEX 5-W S-band
transponder. A medium-gain belt array antenna on a FAST type boom
transmits and receives. Analysis shows a downlink margin of 8.8 dB with a
bit error rate of 10-5 using a downlink rate of 2.2 Mbit/sec.
Fig. 2.c.3.IMAGE spacecraft exploded view shows its
simple FAST-based design.
2.c.4. Power System. The power system consists of the SMEX
21 A-Hr Ni-Cad battery, a 6.6 m2 body-mounted
solar array, the MUE power system electronics, and the FAST shunt control
electronics. The solar array is GaAs with 1.5-mm coverglass to minimize
charged particle damage. Resistors on the back of the solar array allow
shunting of excess power. Unlike those on FAST, the upper and lower solar
arrays are identical with simple geometry, flat sides, and minimal cutouts to
reduce costs.
Table 2.c.2. IMAGE requires only once daily DSN contacts.
Power system analysis shows the output from the solar arrays varies as a
function of the spin axis to the sun-line angle (Fig. 2.c.4.). This angle never
falls below 23.5 because of the 90 degrees inclination orbit with a RAAN of 0
or 180 degrees. The battery provides ample capacity for initial attitude
acquisition and allows the spacecraft and instruments to operate nominally
through eclipse periods of up to 89 minutes. GSFC orbit analysis shows the
longest eclipse period to be 78 minutes.
Fig. 2.c.4. IMAGE Power System provides ample end-of-life power
during all spacecraft attitudes and eclipses.
2.c.5. Attitude Control System (ACS). The IMAGE ACS is identical to
that on FAST, except that three torque rods are used instead of two magnetic
coils on FAST, and two Horizon Crossing Indicators (HCIs) are used instead of
one on FAST. The MUE houses the ACS electronics. The ACS sensors consist
of the two HCIs, a Fan Sun Sensor (FSS), and three-axis magnetometer. The
HCIs provide spin angle and rate by detecting the Earth's horizon every
spacecraft revolution and provide spin-axis attitude from the difference
between the two HCI readings. The FSS provides the Sun's attitude, from
which coarse spin-axis attitude is determined. The magnetometer measures
the Earth's magnetic field as required to perform magnetic torquing. A fluid
damper is included to reduce spacecraft nutation.
ACS simulations performed by GSFC for FAST with 2 HCIs have
demonstrated spin-angle and spin-axis knowledge accuracies that meet
IMAGE requirements (Table 2.c.3.). The Ithaco HCIs have been used at
altitudes of up to 20 RE, greater than the 7 RE maximum altitude of
IMAGE.
2.c.6. Thermal Control System. The IMAGE thermal design is similar
to FAST in that all components mount to the aluminum deck, which also
serves as a heat sink. The deck conducts heat to the bellyband radiator (Fig.
2.c.3.). The spacecraft interior is covered with multi-layer
thermal blankets, and the solar arrays are thermally isolated from the
spacecraft. Analysis shows that with an orbit average power of 94.2 W and a
silver Teflon radiator of about 0.5 m2, the spacecraft is kept
within its operating limits with minimal heater power (Table 2.c.4.). The
spacecraft battery is thermally isolated with a separate radiator to maintain
its tighter temperature limits. If required, thermostatically controlled
heaters will maintain operating and survival temperatures.
2.c.7. Contamination Control. Several IMAGE instruments are
susceptible to molecular contamination, particularly from hydrocarbons and
water vapor. By using clean rooms with instrument purging and standard
materials selection practices, instrument integrity will be protected.
2.c.8. Radiation Protection. IMAGE will copy the FAST approach to
radiation protection of using 30-krad rated parts and providing shielding to
reduce total dose to the 30-krad level. A total radiation dose analysis was
performed for the IMAGE orbit by SwRI. By providing minimum shielding at
the box level equivalent to 3.8 mm of aluminum, the predicted radiation dose
on the spacecraft and instrument parts is limited to 30 krad. Additional
shielding is required on the MUE and transponder, which has been accounted
for in the spacecraft mass budget.
2.c.9. Electromagnetic Interference and Electromagnetic Compatibility
(EMI/EMC). The IMAGE spacecraft does not contain any highly EMI-
susceptible instruments such as low-energy plasma instruments. Thus,
standard spacecraft engineering practices are expected to provide adequate
protection even in the presence of the RPI instrument, which transmits peak
radio pulses of 10 W at frequencies from 100 kHz to 3 MHz, falling linearly
with frequency to 0.1 W at 3 kHz. To assure that this output does not interfere
with spacecraft or instrument systems, all instruments and spacecraft systems
will be tested in radiation fields well in excess of the RPI levels throughout
the entire frequency range of RPI operation.
2.c.10. Government Facilities. The IMAGE spacecraft development
team plans to use NASA GSFC integration and environmental test and UV
calibration facilities. No new or mission unique facilities are required.
2.d. Manufacturing, Integration, and Test.
The IMAGE manufacturing strategy involves assigning responsibility for
support systems used in multiple instruments (e.g., HVPS, detectors, etc.) to
one team member. The team members assigned such responsibility have
demonstrated their abilities to build high-reliability systems on multiple
missions. Examples include the production of high-voltage power supplies by
the University of Maryland and the production of MCP detector systems by
ISAS. Table 2.d.1. shows the institutions responsible for development, test,
and calibration of each instrument.
2.d.1. Schedule. The schedule for the IMAGE science planning and
instrument development, test and integration is shown in Fig. 2.d.1. Details
are contained in Volume II.
Table 2.d.1. Seven U.S. institutions will have primary
responsibilities for development of the IMAGE science payload. Significant
industrial and foreign contributions to the hardware are also listed.
2.d.2. Manufacturing Processes. Manufacturing processes for
IMAGE instruments are well-established, mature processes performed by
experienced operators. The list of common processes includes the following:
(1) machining; (2) anodizing; (3) alodining; (4) soldering (surface mount and
through hole); (5) conformal coating; (6) crimping; (7) electrostatic discharge
machining; and (8) aluminum and gold blackening. All processes are
performed with written procedures and are monitored by team-member QA
inspectors. Post-processing inspection by QA inspectors is a normal part of the
process close-out.
2.d.3. Software Production. As described in Volume II, our software
development process is based on MIL-STD-498 and the guidelines of the
Software Engineering Institute's (SEI) Capability Maturity Model (CMM).
Requirements for software will be documented in a software requirements
document. Design for software will be documented in a software design
specification. We will produce our software in ANSI Standard C developed
under the Greenhills Corp. software development environment hosted on a
SunSparc or IBM R-6000 work station. A "chip -on- processor" (COP)
emulator port on the RAD-6000SC allows comprehensive test and debug
support. The C++ applications code will run under VxWorks Ver. 5.1. Device
drivers for most I/O devices are available as COTS item from Wind River
Systems or IBM. Custom device drivers can be written if necessary. All
software procedures will be tested using written test procedures. Software will
be maintained under configuration control using SEI-CMM Level 2,
beginning at the software CDR.
Fig. 2.d.1. IMAGE instrument definition, development,
test and integration schedule.
2.d.4. The Transition from Design to Production. Instrument designs
are transitioned to production only after a thorough design review process
has been completed. For each instrument, two IMAGE team design reviews
are scheduled. In addition, the IMAGE Independent Review Team (see Vol.
II) will be asked for a recommendation on the suitability of a design for
production. The P.I. will make the final decision on suitability for production
based on the recommendation from the responsible instrument team leader
and his management team.
2.e. Mission Operations and Ground Data Systems
The IMAGE data system has been designed with an end-to-end approach,
eliminating the problems associated with separating spacecraft design, ground
system, and mission operations. The design eliminates duplication of effort
and addresses the needs not only of the IMAGE team but of the scientific
community and the public as well. The IMAGE mission operations and data
flow systems are characterized by seamless integration with the Internet and
World Wide Web (WWW). Nearly all transactions among science
investigators, participating scientists, engineers, and the public will be
handled via Web-browser interfaces. All relevant mission documentation
will be published on-line, as well as by conventional means when needed, to
ensure the widest and most timely availability possible.
A functional diagram of the ground data system is shown in Fig. 2.e.1. The
telemetry stream will be routed into the IMAGE Science and Missions
Operation Center (SMOC) at GSFC. The SMOC is based on similar facilities for
the SMEX program and will consist of a number of workstations and support
systems. The SMOC handles the health and safety functions of the spacecraft.
IMAGE team members will check on the health and safety of their
instruments through their SMOC Internet interface. The SMOC will also take
the input data stream and decommutate and format the data into arrays for
each instrument. Each array corresponds to a raw image browse product,
which is generated by the SMOC from software provided by the instrument
team. The level-0 data and browse products are immediately sent to NSSDC
and to the instrument teams. The IMAGE team will provide to the NSSDC
various higher-level data products (fully calibrated images in geophysical
coordinates), along with all associated software tools. The scientific
community and general public will access IMAGE data directly
from NSSDC through the WWW. The browse products will be the main
source of data for the general public and educational outreach programs as
described in Vol. II. Staffing has been identified to maintain these public areas
and provide technical support.
Fig. 2.e.1. IMAGE uses existing Explorer ground data systems.
Instrumenters will develop command sequences at their own institutions
and send them to SwRI for integration into command loads. The SMOC will
receive the loads and will verify them before uploading to the spacecraft over
a secure network.
Spin Axis Direction
* less than 1 degree perpendicular to the orbit plane
Spin Axis Stability
* 0.005 degrees/s
Attitude Knowledge (at apogee)
* 0.1 degree spin angle
* 0.1 degree spin axis
Spin Rate (final)
* 0.5 +/-0.1 rpm
Table 2.a.1. The IMAGE spacecraft pointing requirements
are within the FAST spacecraft capability (see Section 2.c.5.).
LENA MENA HENA
Energy range 0.01-0.3 1-30 10-200 [keV]
Energy 0.8 0.8 0.7 [[[Delta]]E/E]
resolution
Instantaneous 8x90 8x90 8x90 [deg]
FOV
Total FOV 90x360 90x360 90x360 [deg]
Pixel 8x8 8x8 8x8 [deg]
resolution
Total 0.2 0.33 1.6 [cm2sr]
G-factor
Pixel 7x10-4 2x10-3 1.610-3 [cts/atom cm2sr eV/eV]
sensitivity
Image time 300 300 300 [s]
Pixel array 2Mx4E11A 2Mx4E11A 2M4E11A
dimensions*
UV rejection 10-7 10-9 10-9
Electron rej. 10-10 10-9 10-5
Ion rejection 10-8 10-8 10-5
*3-dimensional:
M = mass, E = energy, A = polar angle
Inst Measurement Focal Aper- lambda la Instan Tot. Pixel Sensi Time
Spatial
. Type Length ture bda]] mbda t. FOV FOV -tivi Res. Res.
(mm) (mm2) (nm) ]] FOV (deg) (deg) ty (min @7RE
Res. (deg) (c/R/ ) (km/pix
(nm) pixel )
)
EUV Plasmaspher 75 25 30.4 5.0 3090 90 0.5 2.4 2 640x640
ic He+ ( 3) 360 0.5
30.4 nm
FUV/ Auroral 89 3.9 121.6 0.2, 166.5 16 0.13 0.04 2 100x100
SI LBH ,135. 3 360 0.13
(narrow 6
band),
Ly-[[alpha]
]
FUV/ Auroral 22.4 60.8 134-1 26 3022 30 0.07 0.1 2 60x60
WIC LBH (broad 60 360 0.07
band)
Table 2.b.2. IMAGE photon imagers provide images of
geocoronal, auroral, and plasmaspheric emissions with a 2-min. time
resolution (one image per spin).
Measurement Nominal Res. Limits
Echo Range 500 km 0.1 to 20 RE
Angle-of-Arrival 1 deg. function of S/N
Ne Density 10% 0.1%
Doppler 0.125 Hz 0.004 to 0.3 Hz
Time 8 sec. 4 sec to 34 min
Table 2.b.3. RPI measurement capabilities.
Parameter Nominal Limits
RF Power 10 W 10 W+
Pulse Width 53 msec 3.3 to 210 msec
Receiver 300 Hz fixed
Bandwidth
Pulse Rate 2 pps 1 to 5 pps
Frequency 10-1000kHz 3 kHz to 3 MHz
Range
Frequency 5% > 1Hz
Steps
Coherent 8 sec. 4 sec to 34 min.
Integ. Time
+Antenna voltage is maintained 3 kV; power drops below 10 W at
frequencies below 100 kHz.
Instrument Mass Power Data Dimensions* FOV Heritage
(kg) (W) Rate (L W H) (cm) (deg) Instr./Spacecraf
(kbps) t
NAI
LENA 8.0 5.3 1.0 33 29 31 8 90 TIDE,TIMAS/Polar
MENA 7.0 7.0 1.0 38 20 17 8 90 TIDE/Polar,
CAPS/Cassini
HENA 8.0 12.0 2.5 35 38 34 8 90 MIMI/Cassini,
SEPS/Polar
FUV
SI 8.7 6.0 6.0 62 36 16 16 1.2 SEPAC/AEPI
WIC 1.9 3.0 4.0 26 15 13 22.4 30 Viking, Freja
Electronics 2.6 3.0 n/a 15 20 16 n/a SOHO, TRACE
EUV
Sensor (3) 6.8 6.0 4.5 15 (dia) 17 30 90 Alexis, Viking
(ttl) (ttl) ea.
Electronics 3.5 (total) n/a 15 15 10 n/a Alexis
RPI
Antennas X 3.6 0 0 0 0 0 0 36 18 18 n/a ISEE, WIND,
axis (2) Y ea. 3.6 36 18 18 WISP Rocket,
axis (2) Z ea. 8.2 4.8 22.9 Stacked Oscars
axis (2) 0.55 ea. on Scout (SOOS)
Electronics, 11.0 11.8 7.0 36 27 14 n/a DPS, GP-B,
incl. CIDP Space Station
Furnace
Totals 73.0 54.1 26.0
*Excludes collimator dimensions for LENA and MENA.
Component Mass Orbit Item Source or Heritage/Additional
(kg) Avg. Information
Power (W)
Instruments 73.0 54.1 See Table 2.b.5.
Spacecraft Subsystems
C&DH/Electrical
Mission Unique Electronics 18.5 FAST MUE, with additional 3 kg
(MUE) 18.0 shielding
Wire Harness & Umbilicals
15.0
Communications
S-band transponder 5.0 ACE, TRACE, & WIRE (add 1 kg.
5.1 shielding)
Antenna & cabling Medium Gain Belt Array
2.5
Power
Solar Arrays 56.0 6.6m2 GaAs on aluminum honeycomb
substrate
Shunt control electronics 1.4 5.0 FAST Shunt Control Electronics
Battery (21A-hr) 23.1 MIDEX/SMEX 21 A-hr super Ni-Cad
Attitude Control
ACS Torque rods 15.4 5.3 Ithaco Torqrods #D4311 (2) &
#D42553
(1)
Magnetometer (total) SMEX 3-axis Magnetometer
1.0
ACS sensors Ithaco HCI(2) & Adcole FSS(1) as on
1.0 FAST
Nutation damper FAST-type fluid damper
5.0
Mechanical
Primary structure 33.9 FAST-type aluminum structure
Bellyband Structure & Aluminum with silver Teflon
coating
Radiator 9.0
Battery Mount & Radiator Sized for 21 A-hr battery, incl. 1
3.0 face of bellyband
Antenna Boom FAST-type composite antenna boom
2.0
Fasteners, Misc. bracketry Based on FAST and SMEX S/C
10.0 percentages
Balance weights Based on preliminary C.G. offset and
10.0 POI's
Thermal Control
Thermal blankets Aluminized Kapton multi-layer
4.0 insulation
Heaters and thermostats 10.0 Minco film heaters and Elmwood
2.0 thermostats
Transponder heatsink per MIDEX AO Appendix B
1.1
Spacecraft Bus Total 218.5 43.8
IMAGE Total 291.5 97.9
Capability 373.0 129.8 Worst case, end of life
Mass & power margins Percent 81.5 31.9 25% ACS torquers are not
used after the
22% first 40 days (spinup), resulting in
5.3 W additional margin.
Table 2.c.1. IMAGE mass and power budgets meet MIDEX 20%
margin requirements.
Combined instrument data rate 26kb/sec
Housekeeping data rate 2 kb/sec
Total data rate 28 kb/sec
Daily storage required (20% 1.78 Gbits
margin, 1.8:1 instrument
data compression)
CIDP bulk memory 2 Gbits
Daily Downlink time 13.45 min
@ 2.2 Mb/s
Available contact time ~14.5 hr.
per day
Attitude Knowledge FAST IMAGE
Error at Apogee Performance(2 HCI's) Requirement
Spin Angle 0.042 degree 0.1 degree
Spin Axis 0.060 degree 0.1 degree
Table
2.c.3. FAST ACS system meets IMAGE pointing knowledge
requirements.
Component Operational Survival
Limits Limits
S/C -10 to 40 -30 to 50 [deg. C]
Electronics
Instruments -10 to 30 -30 to 50 [deg. C]
Battery 0 to 20 -10 to 30 [deg. C]
Table 2.c.4. Wide temperature limits allow for passive thermal
control.
SwRI
* Integration of science payload
* Development of CIDP
* Development of RPI electronics
* Development of MENA
* Dev. of flight software
GSFC
* Development of LENA mechanical systems
* Development of LENA LVPS
* Calibration and test of LENA
* Calibration of HENA and FUV SI
MSFC
* Fabrication and test of FUV WIC
* Development of HENA controller
LPARL
* Development of FUV SI
* Development of LENA conversion surface
* Design of LENA particle optics
U. Md
* Dev. of HVPS and TOPS electronics for NAI (LENA, MENA, HENA)
U. Ariz.
*Development of EUV
* Development of HENA senor
U. Mass.
* Design of RPI
* Test of RPI
AEC-Able Engineering will provide RPI antennas and deployers.
U. of Bern will provide design of LENA conversion surface.
U. of Calgary will provide design and calibration of FUV WIC.
ISAS will provide MCPs for NAI (LENA, MENA, HENA).
U. of Liege will provide mechanical design of FUV SI.
