Date: Wed Feb 23 1994 12:08:00 Subj: Clementine The DSPSE mission involves approximately t
Date: Wed Feb 23 1994 12:08:00
From: Royston Paynter
Subj: Clementine
The DSPSE mission involves approximately two months in lunar orbit and then
departure on a trajectory that will cause the spacecraft to pass by the
asteroid 1620 Geographos at close range. The time in lunar orbit provides an
opportunity to collect lunar images and data for scientific investigation. The
close range flyby of 1620 Geographos provides an opportunity to collect images
and data from the asteroid for scientific investigation. The mission is
composed of the following operational phases:
o Low Earth Orbit (LEO)
o Lunar Transfer Trajectory: Translunar Transfer Injection (TTI) to
Lunar Insertion
o Lunar Orbit
o Geographos Transfer Trajectory:
- Lunar Departure to Lunar Swingby
- Lunar Swingby to Geographos Pre-Flyby
o Geographos Pre-Flyby and Flyby
o Post-Flyby Operations
o Post-Mission Support
Lunar Orbit Operations:
The lunar mapping phase of the DSPSE mission is planned to last 70 days and
will include 57 days of mapping of the lunar surface with the imaging sensors.
The first five days in lunar orbit will be used to establish the lunar mapping
orbit and for pre-mapping mission readiness activities including calibration of
the sensors, autonomous position estimation tests, and imaging special areas of
interest on the lunar surface. The next 27 days will be devoted to lunar
mapping. After the first lunar sidereal day of mapping, approximately one
earth day will be used to change the latitude over which periselene occurs and
the next 27 days will be devoted to the last lunar sidereal day of mapping.
During the lunar mapping phase, mapping will be temporarily interrupted for
orbit adjustments and may be interrupted to divert the sensors from lunar
mapping to observe higher priority targets of opportunity.
The final 8 days following mapping operations will be used for
preparations for injection into the Geographos transfer trajectory, autonomous
position estimation tests, and possible observations of targets of opportunity.
The targets of opportunity may involve a change of orbit if adequate fuel
reserves remain. Another possibility, without changing the orbit, is to divert
the sensors from nadir pointing to observe targets of opportunity.
The Lunar Orbit
On 21 February 1994, the lunar injection maneuver will place the
spacecraft into a preliminary lunar orbit passing over the lunar south pole.
The orbit injection will be accomplished with two burns using an ~ 80%/20%
split in total Fv with time in between for orbit determination. After the first
burn the spacecraft will be in an orbit with an
approximate period of 8 hours. The inclination of the orbit will be 90!11!
with reference to the lunar equatorial plane. DSN range and range-rate and
Pomonkey range-rate and spacecraft accelerometer data will be used by GSFC to
calculate the final orbit shaping maneuver parameters.
After approximately 24 hours (2 orbits) the second burn will place the
spacecraft in the initial mapping orbit with a periselene altitude of 425125 km
occurring at -27! to -30! lunar latitude (TBD) on the sunlit side of the moon
with a 5 hour orbital period to allow for 131.5 orbits per sidereal month.
This geometry allows mapping the southern polar region at a lower altitude.
Orbit shaping will be performed, as required, to adjust periselene and orbital
period to meet mission requirements.
On 25 March 1994, after approximately 32 days, mapping operations will be
temporarily suspended for approximately one day to allow the mapping orbit to
be changed so periselene will occur over +30! to +27! lunar latitude. This
maneuver will allow mapping of the northern polar region at a lower altitude
for the second month of mapping.
On 29 April 1994, two orbit adjust burns will be made to change the periselene
so it occurs over +48! latitude. This orbit will be the departure orbit. On 3
May 1994, a lunar orbit departure RCS burn will be performed to leave the Lunar
orbit and to begin the trans-Geographos trajectory.
Lighting Conditions
The lunar mapping effort will consist of topographic imaging, altimetry and
multispectral imaging for mineral identification. The best data for the lunar
mineral mapping mission presently considered will be obtained if the solar
phase angle is less than 30!. The solar phase angle is defined as the angle
between the vector to the Sun and the vector to the
spacecraft from a point on the Moon's surface. To maximize the time period in
which the solar phase angle is within 30!, the plane of the lunar orbit will
contain the Moon-Sun line half way through the two-month lunar mapping period.
Therefore, insertion into the lunar orbit has been selected so that, as the
Moon-Sun line changes with Earth's motion about the Sun, the Moon-Sun line will
initially close on the orbital plane, and then lie in the orbital plane
half-way through the mapping mission. Assuming a two-month mapping mission and
several days checkout time in lunar orbit prior to beginning mapping, the angle
between the Moon-Sun line and the orbital plane would close for approximately
five weeks before becoming zero.
The current lunar planning trajectory has this angle set at 28.3! at insertion
into lunar orbit, so that at the beginning of mapping (5 days after insertion)
it will be 23.6!. Half way through mapping it will be
-2.4! and at the completion of mapping, it will be -30.3!. At lunar departure
(70 days after insertion) this angle will be -40.3!.
Spacecraft Activities and Tests
A more detailed description of specific experiment activities is contained in
section 6.4.5. The detailed experiment operations are described in the
Appendices. The main spacecraft activities and tests to be performed during
the 70 day lunar mission include:
Pre-Lunar Mapping Checkout Activities:
The initial 5 days will be devoted to establishing the proper mapping orbit,
autonomous position estimation tests, observing special areas of interest on
the lunar surface, and preparations for lunar mapping. These preparations
include:
- Verify the gain settings and exposure times (integration time
settings) for each camera and assessing the ability of the on-board
software to control these settings;
- Determine the range of altitudes for which the laser ranging system
can measure the range to the lunar surface;
- Assess jitter effects from spacecraft pointing activities, filter
wheel movement, solar array movement, and cryocoolers in a lunar
mapping scenario;
- Assess orbit determination accuracy and error analysis;
- Perform Autonomous Position Estimation Experiment activities. This
will involve imaging the lunar limb from the dark side of the moon.
- Perform Autonomous Operations and Scheduling including testing of
the auto-sequence code;
- Assess on board maintained pointing vectors to the center of the
Moon, Sun, Earth, and each of the ground stations;
- Rehearse lunar mapping operational readiness :
-- Command script generation, verification, uplink, and execution;
-- Collection of imaging data;
-- Real-time downlink of image data during data collection;
-- Downlink of image data stored on the Solid State Data Recorder
(SSDR);
-- Data transfer, processing, and display.
- Collect high quality images and altimetry and downlink to the
ground.
Lunar Mapping Activities:
The formal collection of lunar mapping data is planned to take
approximately 55 days or two lunar sidereal days (1 lunar sidereal day is
approximately 27.3 earth days)]. The UV/Visible (UV/Vis) camera with 5 filters
and the near infrared (NIR) camera with 6 filters will each collect overlapping
images. Each filter is used along the satellite's ground track such that
complete mapping of the lunar surface for each of the filters of each camera
will be accomplished during the mapping operation. The High Resolution (HiRes)
camera will make a continuous string of overlapping images with TBD filters
along the track of the satellite on the sunlit side of the Moon. The laser
ranging system will be used when the spacecraft altitude is below 500 km (TBD)
at a pulse rate of 1 Hz (TBD) to preclude exceeding its thermal constraints.
The LWIR camera will be used to make images of areas of interest along the
spacecraft's path, especially near the terminator on the sunlit and dark sides
to collect temperature change information. Since this requires terminator
motion along the lunar surface, off-nadir pointing will be required. Complete
mapping of the lunar surface will not be accomplished with the HiRes camera,
the laser ranging system or the LWIR camera.
Autonomous position estimation and orbit determination tests, and autonomous
operations scheduling tests flowing from the position
estimation and orbit determination tests, will be conducted during all portions
of the lunar orbit phase. A goal is to verify the accuracy of the autonomous
position estimation and orbit determination and operations scheduling functions
during the earlier portion of the lunar orbit phase and then let the spacecraft
schedule and execute operations for one or more lunar orbits during the later
portion of the lunar orbit phase.
Preliminary Operational Scenario:
The formal collection of lunar mapping data is planned to take
approximately 55 days or two lunar sidereal day (1 lunar sidereal day is
approximately 27.3 earth days)]. The UV/Visible (UV/Vis) camera with 5 filters
and the near infrared (NIR) camera with 6 filters will each collect overlapping
images. Each filter is used along the satellite's ground track such that
complete mapping of the lunar surface for each of the filters of each camera
will be accomplished during the mapping operation.
The UV/Vis and NIR cameras need to be turned on 10 minutes before imaging. The
NIR cryocooler needs to be on 30 minutes before imaging. In
successive orbits, the UV/Vis and NIR camera images will be taken alternatively
from -90! to +60! and -60! to +90! latitude, when
periselene is at -30! latitude, and -60! to +90! and -90! to +60! latitude,
when periselene is at +30! latitude. This mapping strategy is designed to save
power and storage space since there is no need to take images of polar regions
on consecutive orbits. The integration times and gains will be varied as a
function of latitude and may be computed on board. The time to change filters
and to dampen is assumed to be 200 -250 ms. The UV/Vis camera has a stray
light exclusion angle of 40! (full angle). The UV/Vis camera, along with the
star trackers will also be used to collect data for autonomous orbit
determination tests by obtaining lunar limb and star images. There is also a
requirement to obtain frequent sensor calibration data for all the mission
sensors. These tests will be taken for TBD minutes each orbit as the timeline
allows.
The High Resolution (HiRes) camera will make a continuous strip of overlapping
images with up to 4 filters along the track of the satellite on the sunlit side
of the moon. Images using the HiRes camera will be taken -90! to +90!
latitude. The camera integration time and gains will be varied with the
latitude and may be computed on board. The time to change filters and to
dampen is assumed to be 200 -250 ms.
The laser ranging instrument (LIDAR) will be used whenever the altitude is less
than TBD km. The maximum usable altitude will be determined while in Lunar
orbit, but will not exceed 640 km. During initial lunar orbit operations,
tests will be conducted to determine the maximum altitude from which effective
altimetry can be performed. The laser ranger will be turned on TBD minutes
prior to use. The LIDAR electronics will be turned on 10 minutes before use
and the heaters will be turned on 15 minutes prior to use to bring laser
transmitter temperature up to +25 degrees. The laser pulse rate will be limited
to 1 Hz to keep temperatures in range.
The Long Wavelength Infrared (LWIR) camera will be used to take images of
thermal gradients occurring at the terminators. The cryocooler for the LWIR
must be on 30 minutes prior to imaging. The LWIR camera electronics will be on
10 minutes prior to imaging. Images will be taken 110! of each terminator over
the poles and the camera will be turned off between usages, but the cryocooler
will remain on between North and South pole image sequences. As with the NIR,
UV/Vis and HiRes cameras the
integration time and gain will be varied with the latitude and may be computed
on board.
During lunar mapping, the star trackers will not be used to collect scientific
data, but will be used to establish the attitude of the spacecraft. To meet
the high accuracy pointing requirements, the spacecraft attitude will be
updated every 10 seconds during mapping. Star tracker images will be processed
on the R3000 image processing computer and used to update the attitude of the
spacecraft. The star trackers have a full angle solar exclusion angle of 63! x
75! in order to resolve stars. One star tracker will have solar exposure during
each lunar orbit and will be powered off for this time.
Solar panel auto tracking will not be inhibited during imaging. The exact
solar panel position management procedures to minimize jitter is TBD and will
be determined during simulations and verified during the first 5 days of lunar
orbit.
During lunar imaging, the spacecraft will use nadir pointing to meet sensor
requirements for imaging. The best method will be finalized during the initial
5 day operational readiness phase. Attitude changes and attitude updates must
be scheduled with the camera image sequences and positioning of the solar
panels. Attitude commands will be made to the guidance software, which will
generate attitude control parameters for the desired attitude and supply these
parameters to the ACS software. Attitude parameters will be generated based on
any attitude constraints, primary pointing requirements (nadir or earth
pointing), and secondary pointing requirements (maximizing solar incidence
angle on solar arrays). The ACS software will compare the guidance generated
attitude parameters with the current attitude and generate the required
commands to slew to the desired attitude. The reaction wheels or thrusters
will be used for the slew depending on the amount of slew and time available
for the slew. Slews using the reaction wheels will take up to 10 (TBD) minutes
maximum.
After lunar mapping sequences and autonomous position estimation and sensor
calibration tests are complete for an orbit, the spacecraft will dump data to
the ground using the high gain antenna. This dump will take between ~120 to
130 minutes to completely downlink the data from the SSDR. Engineering data
will be stored and dumped along with the image data using the high gain
antenna. Engineering and limited image downlink via the omni antennas during
lunar mapping will be possible when orbital geometry and ground station
visibility permit.
Post Mapping Activities:
During the 8 days following the completion of the lunar mapping there will be
time for special observations of targets of opportunity. These observations
will provide images of lunar surface features and man-made objects on the Moon
at the Apollo, Surveyor, and Russian landing sites. There will be targets that
will be imaged without changing the orbit. Other targets of opportunity may
involve a change of orbit for
observations which will be done if adequate fuel remains for the
Geographos transfer and flyby. See section 6.4.7 for further details.
Prior to lunar departure there will be two orbit adjust maneuvers to prepare
for the lunar deorbit maneuver. This will change the periselene so that it
occurs at +48! latitude. Pomonkey will supply range-rate and the DSN sites
will supply range and range-rate data for orbit
determination to GSFC who will compute the state vector and supply it to the
DMOC. The TAMP will verify parameters, develop command sequences, and verify
these sequences using the OTB spacecraft simulator. The Lunar Orbit Departure
will be fully rehearsed to acquaint operations personnel with the activities
and timing of the departure RCS burn and to verify the activities and sequences
can be smoothly executed.
Ultraviolet/Visible (UV/Vis) Camera
The UV/Vis sensor is a CCD video camera with an 8-bit digitization of the array
data. The actual ground resolution from 400 km (TBD) altitude is between 79 to
106 meters depending on the actual amount of jitter. Six bandpasses are
defined by filters in a six position filter wheel. One of the six filter wheel
positions will be required for a very wide bandpass filter, 400 to 950
nanometers (nm), so virtually the entire sensor bandpass can be used to allow
early identification of Geographos and closed-loop tracking control. The five
remaining filter bandpasses were specified by the NASA Science Advisory
Committee (SAC) and SDIO.
The UV/Vis camera will be mounted so the 4.2! dimension of its field of view is
in the spacecraft's velocity direction (X-axis - along track) for the lunar
mapping phase and the 5.6! dimension of its field of view is perpendicular to
the spacecraft's velocity (Y-axis - cross track).
For temperature stabilization, the UV/Vis camera electronics must be turned on
for 10 minutes before it can be used to collect image data. Radiation exposure
could damage the UV/Vis; however, the effects can be corrected on the ground.
Near Infrared (NIR) Camera
The NIR sensor is a cooled video camera with an 8-bit digitization of the array
data. The actual ground resolution from 400 km (TBD) altitude is between 112
to 128 meters depending on the actual amount of jitter. The NIR has a six
position filter wheel with the filters as shown in Table 6.2-2 selected by the
DSPSE SAC and SDIO. The NIR sensor has a
mechanical cooler (cryocooler) to bring the CCD array to a sufficiently low
temperature for operation.
For temperature stabilization, the NIR camera's cryocooler must be turned on
for 30 minutes and the camera electronics must be turned on for 10 minutes
before it can be used to collect image data.
Long Wavelength Infrared (LWIR) Camera
The LWIR sensor is a cooled video camera with an 8-bit digitization of the
array data. The actual ground resolution from 400 km (TBD) altitude is
approximately 43 meters and is fairly independent of the amount of jitter. The
target temperature range for the LWIR is 250 to 400 K. The sensor does not
have a filter wheel. The LWIR sensor has a mechanical cooler (cryocooler) to
bring the CCD array to a sufficiently low temperature for operation.
For temperature stabilization, the LWIR camera's cryocooler must be turned on
for 30 minutes and the camera electronics must be turned on for 10 minutes
before it can be used to collect image data.
High Resolution (HiRes) Camera
The HiRes camera is the imaging portion of the LIDAR system. The HiRes sensor
uses a frame transfer CCD with a micro-channel image intensifier and uses an
8-bit digitization of the array data. With the current jitter constraints, the
HiRes camera is expected to give a maximum ground resolution at lunar
periselene of 13 m to 30 m (short to long integration time.) Six bandpasses are
defined by filters in a six position filter wheel. One of the six filter wheel
positions will be required for a very wide bandpass filter, 400 to 750 nm, so
virtually the entire sensor bandpass can be used to allow early identification
of Geographos and closed-loop tracking control. Another of the six filter
wheel positions will be opaque to protect the intensified camera from high
light input when it is not in use. The four remaining filter bandpasses were
specified by the DSPSE SAC and SDIO.
Radiation exposure could damage the HiRes camera, however, the effects can be
corrected on the ground. Solar illumination could also damage the HiRes
camera, so imaging should not take place if the Sun is within 2.38! from its
optical axis, and the opaque filter should be selected if the sensor is not
being used for active imaging. For temperature
stabilization, the high resolution camera electronics must be turned on for 10
minutes before it can be used to collect image data.
Laser Ranging System (LIDAR)
The laser ranging system comprises a Nd:YAG laser and a ranging receiver. The
laser emits 180 millijoules per pulse at 1064 nanometers with a beam divergence
of 500 microradians. The ranging receiver is an avalanche photo diode with a 1
milliradian full-angle field of view. The laser ranging system is able to
resolve range measurements to 40 meters. Ranging activity is desirable while
the kick motor recedes from the spacecraft during the lunar transfer
trajectory, during the flyby of the asteroid, and during any low altitude
portions of the lunar orbit phase. The laser ranging system design has been
modified to permit range measurements to ranges of 500 km above the lunar
surface. The laser has a maximum pulse repetition rate of 8 pulses per second,
but is not expected to sustain a pulse repetition rate greater than 1 pulse per
second without thermal problems.
For temperature stabilization, the laser heater must be turned on for 15
minutes before it can be used for ranging.
Star Trackers
The star trackers are star imaging sensors used for 3-axis Rlost in spaceS
attitude determination. The star trackers consist of a S-20 photocathode and a
full frame CCD video camera with an 8-bit digitization of the array data.
Associated with the star tracker is a set of algorithms and a catalog of
selected stars which can be used with data from the star tracker to determine
its attitude to an accuracy of 150 microradians in pitch and yaw and 450
microradians in roll. With a 150 millisecond integration time, the star
trackers can detect stars down to magnitude 4.5.
For reliable operation, sources such as the Sun, reflected sunlight, or bright
limbs must be excluded from a 63! by 75! region about the star trackerUs
optical axis. Staring at the Sun may damage the star trackers, therefore, the
star tracker optics must not be exposed to direct sunlight (sun within the 28!
by 42! FOV) for a period exceeding 3 (TBD) minutes, operating or non-operating.
The two star trackers will each have a different line of sight selected to
allow at least one of the star trackers to determine the spacecraft's
orientation while lunar mapping is being conducted and during all other phases
of the mission. One star tracker will have solar exposure during each lunar
orbit and will be covered and powered off during this time. There is also a
potential for radiation exposure damage to the star trackers which may require
onboard processing to correct.
For temperature stabilization, the star tracker electronics should be turned on
10 minutes before images are obtained for downlink. Star tracker imaging for
attitude determination does not have to wait 10 minutes and can be used right
away.
Science Objectives:
The science objective of the mission is to obtain data useful for scientific
investigations. These investigations involve using the DSPSE mission sensors
to image the lunar surface and the asteroid Geographos. The operational
experiments for these events were discussed above. This section discusses the
scientific reasons for performing these
observations.
The DSPSE mission will provide an abundance of information about the surface
morphology, topography, and composition of both the Moon and Geographos,
providing an insight to their history and processes that have shaped that
history. This information will be used to address fundamental questions in
lunar science and will contribute to significant advances toward deciphering
the complex story of the Moon. The DSPSE mission will also permit a
first-order global assessment of the resources of the Moon and provide a
strategic base of knowledge upon which future missions to the Moon can build.
Lunar Mapping
The pressing need for global mapping of the Moon, by a variety of
remote-sensing techniques, has been stressed for the last 20 years. The DSPSE
mission begins this task allowing a global digital image model (DIM) of the
Moon to be developed. DSPSE lunar mapping will provide improved resolution
allowing for more detailed geologic mapping and will obtain improved spectral
coverage as well as improved spectral resolution over any previous lunar
observations. While the Galileo spacecraft provided spectacular multi-spectral
images of the lunar surface before leaving the Earth/Moon system, the DSPSE
mission offers many significant advances relative to the Galileo data.
o The pixel resolution will be at least 10 times better than
Galileo's, providing improved resolution for unit mapping.
o The HiRes camera images offers up to 100 times better resolution,
allowing for detailed geologic mapping.
o Using the various sensors, improved spectral coverage (up to 2.8
microns) will be obtained.
o With the increased number of filters on the sensors, improved
spectral resolution will be possible allowing improved distinction
between olivine, pyroxenes, and plagioclase.
The altimetry obtained by the laser ranger, will augment the DIM by providing a
set of topographic profiles for the mid-latitude band of the Moon. The DSPSE
data when tied to the Apollo data, will permit knowledge of the true positions
of lunar surface features to within a few hundred meters. Maps of the Moon
made from DSPSE data will enable studies of regional history and permit the
processes of volcanism, tectonism, and impacts that have shaped lunar history
to be deciphered.
The DSPSE data will allow several key unresolved scientific issues to be
addressed:
o Character and evolution of the primitive lunar crust.
o Thermal evolution of the moon and lunar volcanism
o The impact record and redistribution of crust and mantle materials
o Distribution of potential resources.
From the combined UV/Vis and NIR camera images, a global color map will be
formed that can be interpreted in terms of rock types. At a minimum, it will
be possible to recognize and discriminate between the absence of mafic minerals
(pure feldspar) and the presence of orthopyroxene, clinopyroxene, and olivine,
as has been done for the near side of the Moon from Earth-based data. Thus on
a global basis, the distribution of anorthosite, RnoriticS rocks,
olivine-bearing rocks (dunites and
troctolites), and gabbros will be able to be distinguished. For mare deposits,
visible color mapping can classify the mare in terms of titanium abundance, and
element that can be used to estimate the distribution of solar wind hydrogen,
an important lunar resource.
Combined with the knowledge of cratering and the use of basins as probes of the
crust, these data will permit the composition and petrologic structure of the
crust to be reconstructed in three dimensions. The question of the existence
of a magma ocean, the nature of Mg-suite magnetism, the history and extent of
ancient KREEP and mare volcanism, the compositional diversity of mare units,
and the effects of cratering on the composition of the lunar surface can be
addressed. Topographic data from the laser ranger combined with spectral
information will allow the dynamics of large impacts ,e.g., the problem of
depth of excavation for basin-sized impacts, to be modeled.
The high-resolution images from the HiRes camera will allow the surface
processes and compositions to be studied in greater detail. Many mare units
display significant heterogeneity, and color imaging from the DSPSE HiRes
camera images can map different color units, some of which are perhaps related
to individual mare flows. Images of crater walls and central peaks can not
only provide high-resolution compositional data, but permit a better
understanding of the geological setting and processes that have affected given
regions, information that may prove critical to the proper interpretation of
the regional compositional information. Finally, the high-resolution imaging
can be used to make detailed geological studies of the areas of high scientific
interest.
Finally, the tracking information received as part of the normal orbit
determination tasks during the period in lunar orbit will also be used to
provide a more accurate model of the moonUs gravitational potential.
Asteroid Flyby Objectives:
Asteroid flybys are necessary for the characterization of multiple targets to
address issues of asteroid diversity. Flybys cannot address
fundamental questions of elemental composition (for link to meteorites) nor of
internal structure - both of these will require rendezvous missions. The DSPSE
mission will set the stage for NASAUs Discovery Program Near-Earth Asteroid
Rendezvous (NEAR) Program. The NASA science objectives for the Geographos
flyby are exploration and mapping.
Exploration is the key characteristic - no Near Earth Asteroid (NEA) has ever
been investigated close-up before. The mapping will provide insight into the
geological and thermal processes characteristic of this type of asteroid.
Other information to be gained by the flyby and resultant images will be:
o Volume/shape determination
o Spin state determination
o Multi-spectral imaging to infer compositional heterogeneity
o Investigation of regolith with thermal imaging of the dark side.
The DSPSE data will allow several key unresolved scientific issues to be
addressed:
o Relationship of NEAs to main belt asteroids, comets, meteorites,
and the planetesimals that were the building blocks of the
terrestrial
planets (need diversity).
o Characteristics of NEAs which have influenced EarthUs
geological/biological evolution
o Potential of NEAs for sample return and resource utilization.
o Surface processes of small objects with weak gravity fields.
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Lou Wheatcraft, Phone: (713)280-1892; Fax: (713)283-7903
E-Mail: lsw@bonnie.jsc.nasa.gov
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E-Mail Fredric L. Rice / The Skeptic Tank