Ionospheric Mappers

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Ionospheric Mappers
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IM Mission Profile
Launch Date April, 2009
Description A small constellation of satellites
that provide global coverage and characterization
of the ionosphere at all latitudes and local
times Instruments Five in-situ instrument
packages per spacecraft for combined particles,
fields, and gas properties measurements as well
as a GPS tomography instrument and an ionospheric
sounder Spacecraft Eight identical, three-axis
stabilized, spacecraft including a propulsion
system for orbit maneuvers, station-keeping, and
disposal with six spacecraft each in a different
polar orbit plane and two spacecraft at low
inclination in the same orbit plane
Mission Life 2 years with an optional 3-year
extension Orbits 450 km circular--high
inclination (850 and 870) and low inclination
(150 to 450) Space Access One launch on a
Medium Class ELV
Key Technologies Smaller instruments and
enhancing technologies at the subsystem or
component level
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IM Mission Time Line
The following serial time spans are assumed for
mission planning with July 1, 2000 as the initial
reference date

• 3 years for mission unique technology development
• 2 years for studies, project formulation, and
  definitization
• 4 years from approval to launch readiness
• April 2009 launch
• 2 years for baseline mission operations
• 3 year mission extension (option for evaluation)

G O D D A R D S P A C E F L I
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IM Mission Objectives
The IM mission employs a constellation of small
three-axis stabilized satellites distributed in
latitude and local time around the Earth to
gather simultaneous, global data of the
ionosphere. Specific mission objectives are as
follows

• Global characterization and understanding of the
  Earths ionosphere / upper atmosphere (100 to
  1000 km) and its connection to the Sun, solar
  wind, and magnetosphere
• Major improvements of ionosphere and thermosphere
  specification models
• Improvement of forecast and now-cast accuracy
• Establishment of a quantitative baseline for
  Sun-climate studies

G O D D A R D S P A C E F L I
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IM Instrument Complement
The baseline IM instrument complement has been
arranged into five distinct measurement packages
listed below

• (1) Combined Particles Instrument (includes High
  Energy Electron Sensors, High Energy Ion Sensors,
  Low Energy Electron Sensors, and an Electronics
  Box)
• (2) Combined Fields Instrument (includes Langmuir
  Probe, including Boom Magnetometer and 2 m Boom,
  E-Field Booms and Spheres, and an Electronics
  Box)
• (3) Combined Gas Properties Instrument (includes
  Ion Velocity Meter Sensor, Neutral Wind Sensor,
  Ion/Neutral Mass Spectrometer Sensor, and an
  Electronics Box)
• (4) GPS Tomography Instrument (includes Dishes, a
  Patch Antenna, and an Electronics Box)
• (5) Ionospheric Sounder

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IM Instrument Parameters
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IM Systems Synopsis

• This mission requires the use multiple
  satellites, distributed in latitude and local
  time, to globally characterize the Earths
  ionosphere.
• Mission design is based on single string with a
  two year mission design life, but a 5 year goal
  redundancy is inherent in having multiple,
  identical satellites where some failure can be
  tolerated.
• Having spacecraft at two different inclinations
  requires the use of two geographically separate
  ground stations (Alaska, S.Florida).
• The need for multiple spacecraft and multiple
  instrument packages will require the use of
  innovative designs to simplify hardware assembly,
  integration and testing new techniques my have
  to be developed to accommodate building and
  testing these multiple systems within the
  available resources.

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IM Mass Summary
Values are best estimates and do not include
contingency.
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IM Power Summary
Values are best estimates and do not include
contingency.
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IM Delta II 7920 Launch Profile
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IM Orbital Pictorial

• 6 polar orbit planes, 30 deg apart
• Exploiting nodal regression to deploy from one
  initial orbit.
• Launch vehicle delivers three spacecraft each to
  two near polar orbits, inclinations 85 and 89
  degrees.
• Using the difference in nodal regression rates
  for each inclination, allow the two orbit planes
  to separate.
• At selected separations, maneuver spacecraft to
  inclination 87 degrees.
• Repeat until desired configuration is attained.

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IM Orbit Parameters

• The parameters for the final mission orbit are
  given below
• Altitude 450 km, circular orbits
• Inclination
• Six polar orbiting satellites three at 85,
  three at 87
• Two low inclination satellites TBD
• Simultaneous polar coverage, multiple
  mid-latitude coverage

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IM Shadow Periods
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IM Spacecraft Features

• The major IM features include
• A three-axis stabilized, mission-unique bus
• Spacecraft bus will be based on an adapted RBM
  structure for cost savings and to achieve
  economies of scale for both missions
• Subsystem upgrades as appropriate to meet IM
  performance and lifetime requirements

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IM Mechanical Subsystem

• The IM Mechanical Subsystem relies on standard
  aerospace materials and fabrication techniques
  for both spacecraft and instrument support
  structures. Aluminum and/or composites may be
  used to accommodate mass, thermal, or electrical
  constraints.
• A number of critical deployments are required for
  the following elements
• Magnetometer boom
• GPS antenna
• E-field booms
• Langmuir probe

G O D D A R D S P A C E F L I
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IM Launch Vehicle Evaluation
In an effort to minimize launch costs, the
capability of several classes of launch vehicles
capable of delivering a payload to a 500 Km
circular orbit was evaluated. The chart on the
left illustrates that the Pegasus XL, Taurus
2110 and Delta II 7420 do not have adequate
lift capability. Although multiple Pegasus or
Taurus launch vehicles could meet the requirement
for placing 6 satellites in orbit, the cost would
exceed that of a single Delta II 7920. The
Delta II 7920-10 was thus chosen for the concept
study.
Launch vehicle payload capability vs total IM
payload (6 s/c)
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IM Launch Configuration
6 IM Satellites 0.7 meter nominal height
4.2 m
5.3 m
Mission-Unique Transition Adapter
6306 Payload Adapter Fitting
Delta II 7920 with 10-meter Long Shroud
4.0 m
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IM Orbit Configuration
High Energy Particles
Magnetometer Boom
E-Field Booms
GPS Antenna
Low Energy Particles (Top Hat)
Solar cells on all available s/c surfaces
Velocity Vector
Gas Properties
Langmuir Probe
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IM Power Subsystem

• The Power Subsystem is a 28-volt direct energy
  transfer system that can
• support a load of 400 watts at the BOL. It
  consists of the following elements
• A solar array with a total triple junction GaAs
  cell area of 1.54 m2
• A single 13.1-Ah Li-ion battery sized to handle
  transfer orbit, worst-case shadow period, and
  peak power load conditions
• Power electronics
• Solar array degradation over the life of the
  mission due to UV exposure,
• ionizing radiation, thermal cycling and system
  losses have been taken into
• account in the array sizing

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IM Power Subsystem (cont.)
Solar Array Power Margin 196.7W Load
Abs(Beta)14 Deg 129.8W Load Abs(Beta)G O D D A R D S P A C E F L I
G H T C E N T E R
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IM Thermal Subsystem

• The scientific instruments and spacecraft thermal
  designs are thermally coupled.
• Passive thermal control system consisting of
  blankets, heaters, thermostats, thermistors,
  coatings, constant conductance heat pipes (CCHPs)
  and variable conductance heat pipes (VCHPs)
  and/or loop heat pipes (LHPs), and radiators can
  accommodate all instrument and spacecraft
  requirements.
• Solar array will function as thermal radiator
  using heat pipes.
• Instrument electronics will be maintained between
  0 and 20 C.
• Spacecraft components will be maintained between
  0 and 40 C.
• Propulsion system will be maintained between 10
  and 40 C.

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IM Attitude Control Subsystem

• The pointing and knowledge requirements for the
  IM mission are not difficult to achieve the
  Attitude Control Subsystem (ACS) proposed for IM
  can accommodate a number of in-situ instruments
  with the following general pointing requirements
• Pointing Accuracy 3 1 ? about Ram
  vector (driven by gas properties package)
• Attitude Knowledge 0.1 1 ?
• The ACS integrates the following complement of
  hardware for the IM mission
• Coarse Sun Sensor
• Gyro
• DSS
• Reaction Wheel
• Earth Sensor
• Magnetic Coil

Data from the magnetometers and GPS transceiver
from the science complement are also used as
required
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IM Propulsion Subsystem

• The IM propulsion subsystem is a liquid hydrazine
  mono propellant system for dispersion correction,
  station-keeping, and disposal.
• Total ? -V requirement for each spacecraft is 450
  m/s

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IM CDH / Communications Subsystem

• The IM flight Command and Data Handling subsystem
  is relatively straightforward, with the only
  significant requirement being for a large memory
  storage capability (10 Gbyte). This allows for a
  total of 48 hours of data to be recorded,
  providing the capability to miss a scheduled
  downlink opportunity without losing data.
  However, this would require a longer than normal
  contact time or would require two contacts on the
  following day.
• The Communications Subsystem consists of
• An X-band system for science dumps
  (1/day/spacecraft), utilizing a single X-band
  omni antenna on the nadir pointing axis
• An S-band for real time user science and for H/K
  contingency mode with 2 omni antennas

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IM CDH / Communications Subsystem (continued)
OMNI 1
RF S-band Command 2kbps
Command
Hybrid Divide

Low-Power Transceiver (LPT)

RF S-band Up/Down
S-band Diplexer
Telemetry
RF S-band Telemetry 500 -bps RT User Science or
4-kbps H/S Safehold
OMNI 2
On-Board Computer CDH Bus
Playback science _at_ 150 Mbps
X-band Xmitter
X-band Nadir Omni Antenna
Record
10-Gbyte Record
2 days recording
Playback _at_ 150 Mbps
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IM Ground System

• The proposed ground system accommodations take
  advantage of existing
• infrastructure and include the following
  features
• Existing 13-meter systems in Alaska and Florida
  will be used
• Lease vs build option was looked at,
  recommendation is to lease service
• The on-board GPS will be used for orbit
  determination (no ranging required)
• The IM mission operations center can be located
  at GSFC or any other feasible location
• A real-time, continuously broadcast data stream
  (responsible for receiving and decoding the data
  stream

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IM Ground System Concept
Two Lower inclination Spacecraft
Six Polar Orbiting Spacecraft
X-band Playback Science and Housekeeping Data
150 Mbps
S-band Commands 2 kbps
X-band Playback Science and Housekeeping Data 150
Mbps
S-band Commands 2 kbps
S-band 500 bps or 4 kbps House-keeping
Real-Time User S-band Science Data (500 bps)
S-band 500 bps or 4 kbps House-keeping
Omni
Omni
World Wide Real-Time Users
Alaska 13-M station
World Wide Real-Time Users
South Florida 13-M station
T3
Commands
ATM Data
T3
ATM Data
Mission Operations Center
Commands
Science Operations Center
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IM Mission Operations
A mission operations concept has been chosen that
encourages automation of routine spacecraft
functions and makes use of commercial
off-the-shelf (COTS) products. Salient features
include the following

• Combined Mission and Science Operations Center
  (MSOC) co-located with dedicated ground station
  at GSFC (or other location)
• Automated mission operations using COTS command
  and control system
• Science data processed to Level Zero and
  short-term archival at MSOC
• Data distribution to Principal Investigators
  (PIs) via ftp
• On-board recording of health and safety data to
  support anomaly resolution

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IM Mission Specific Technology

• The IM mission concept incorporates new
  technology that is expected to be
• available in the near term. Such items include
• High-efficiency, triple-junction, GaAs solar
  cells
• Li-ion battery
• Low power, lightweight GPS transceiver
• Miniature earth horizon sensor

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IM Study Options

• Although the type of orbit for the IM mission was
  readily determined, the best way to deliver the
  spacecraft to these orbits was more difficult to
  assess. The advantages and disadvantages of a
  number of ways to deliver six s/c to the desired
  orbits was assessed as part of the IM concept
  study. Delivery methods evaluated included direct
  insertion and providing sufficient on-board
  propulsion. It was concluded that taking
  advantage of nodal regression from two initial
  orbits would allow the desired orbit
  configuration to be achieved with a minimum ?-V
  requirement. Although this method requires
  several months before the final s/c are in the
  proper orbit, valuable science data can be taken
  during this transition period.

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IM Preliminary Risk Assessment

• During the course of the IM concept study, a
  number of risk areas were
• identified and are listed below. Further study
  will be required to fully assess
• these risks, their potential impact, and
  mitigation strategies.
• Low power transceiver is not yet flight qualified
  
• At the current mass, the c.g. of the six stacked
  s/c in the Delta II 7920 is above the allowable
  limit
• Use of heat pipes in thermal system has
  implications for integration and testing
• Availability of anticipated technology
  enhancements

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IM Study Recommendations

• Conduct a survey of instruments now under
  development or planned for future development to
  ensure adequacy of assumed IM instrument resource
  requirements.
• More detailed study of orbits is required
  (elliptical orbits should be looked at)
• Mission operations concept for 6-8 spacecraft
  must be developed in more detail
• Use of a star tracker for attitude determination
  must be investigated
• Further mass reduction must be looked at to
  resolve mass margin and c.g. issues
• Frequency authorization for continuous radiation
  (real-time data broadcast) must be investigated
• Additional study and perhaps technology and/or
  process development needs to be done on how to
  build and test multiple research grade flight
  instruments and spacecraft particular attention
  should be given to the areas of calibration and
  quality assurance.