Minimize IMAGE

IMAGE (Imager for Magnetopause-to-Aurora Global Exploration)

Spacecraft   Launch   Mission Status   Sensor Complement   References

IMAGE is the first MIDEX (Medium-class Explorer) mission of NASA/GSFC with the overall objective to study the global response of the Earth's magnetosphere to changes in the solar wind. Observations of high spatial and temporal resolution 3-D imagery of magnetospheric plasma motions. IMAGE uses ENA (Energetic Neutral Atom), ultraviolet, and radio imaging techniques to: 1) 2) 3)

• Identify the dominant mechanisms for injecting plasma into the magnetosphere on substorm and magnetic storm time scales

• Determine the directly driven response of the magnetosphere to solar wind changes

• Discover how and where magnetospheric plasmas are energized, transported, and subsequently lost during substorms and magnetic storms.

SwRI (Southwest Research Institute, PI: James L. Burch) of San Antonio, Texas, is the prime partner of NASA in this project.


The minisatellite, built by LMMS (Lockheed Martin Missiles & Space) of Sunnyvale, CA, employs a spin-stabilized platform. It has the form of a regular octagon and measures 2.25 m in diameter and 1.52 m in height. Surface-mounted solar cells (high-efficiency, dual-junction GaInP2/GaAs/Ge cells) provide an efficiency of 20-21.5% with average power of 250 W (21 Ah NiCd batteries for eclipse operation). S/C mass = 496 kg.

IMAGE has a nominal spin period of 2 minutes (or a spin rate of 0.5 ± 0.01 rpm); its spin axis is perpendicular to the orbital plane. The AD&C (Attitude Determination & Control) subsystem consists of the following actuators/sensors: a magnetic torque rod, a passive nutation damper, a three-axis magnetometer for magnetic aspect information, an enhanced sun sensor, and AST (Autonomous Star Tracker) developed by LMMS. AST is mounted with its boresight 10º from the spin axis. The two-axis sun sensor provides spin rate and sun aspect angle information. MCS (Magnetic Control System) controls both spin axis orientation and spin rate. Attitude knowledge is accurate to within 0.1º.

The C&DH (Command & Data Handling) subsystem design employs the MIL-STD-1553B bus which communicates with the instrument controllers over RS-422 interfaces. The CCSDS protocol is used for all internal and external communication. The 1553/CCSDS protocol design provides substantial advantages in terms of spacecraft/payload decoupling (the S/C serves mainly as a "bent pipe" for science data produced by the payload). All time synchronization between S/C and payload is accomplished exclusively over the 1553 bus. 4) 5) 6)


Figure 1: Isometric drawing of the IMAGE spacecraft (image credit: NASA, SwRI)

Mechanical subsystem

Aluminum honeycomb side panels (8), forward and aft panels, payload deck, and interior shear walls; two central load-bearing aluminum cylinders (forward and aft)


Spacecraft: 340 kg, Instrument Payload: 196 kg, Total: 536 kg

Thermal subsystem

Heat pipes in payload deck connected to radiators on S/Ct side panels; MLI blankets; thermostat- and CIDP/PDU-controlled electrical resistance heaters for payload and spacecraft operations and survival

Command & data handling subsystem

System Control Unit (SCU) consisting of RAD6000 flight computer plus modules for I/O, mass memory, command & telemetry, communications & memory, and power supply; VME backplane; 4.096 Gbit DRAM in mass memory module; Mil-Std-1553B interfaces to other observatory systems

Communications subsystem

S-band transponder; 1 medium-gain helix antenna; 2 low-gain omni-directional antennas; uplink data rate: 2 kbit/s; downlink data rate: 38 kbit/s (real-time mode); 2.28 Mbit/s (stored data playback mode)

Attitude determination & control subsystem

Spin-stabilized; closed-loop spin-rate control
Sensors: Adcole 44690 sun sensor; Lockheed Martin ATC AST-201R autonomous star tracker; MEDA TAM-2A 3-axis magnetometer
Actuator: 1 Ithaco magnetic torque rod

Solar arrays

Body-mounted dual-junction gallium indium phosphide/gallium-arsenide/germanium solar cell arrays


21 Ah super NiCd battery; operating range: 22-34 VDC

Table 1: Overview of S/C design features


Figure 2: Block diagram of the IMAGE electrical power subsystem (image credit: SwRI)

RF communications: Three antennas are used for S-band communication with the ground. A medium-gain helix antenna and two low-gain omni-directional antennas. One of the omni antennas is mounted on the aft end panel of the spacecraft; the other is mounted together with the helix antenna on the forward panel. The helix antenna is used to transmit data from the spacecraft to the ground; the co-mounted omni antenna is used to receive uplinked commands and data. Uplink data rate = 2 kbit/s. Downlink of stored science, engineering, and housekeeping data at a rate of 2.28 Mbit/s. In addition to the playback of stored data, the S/C also continuously transmits real-time data at a nominal rate of 38 kbit/s. The real-time data is mainly for CRL of Tokyo, Japan and for NOAA (space environment weather forecasts).


Figure 3: Photo of the IMAGE spacecraft (image credit: NASA)


Launch: The IMAGE spacecraft was launched on March 25, 2000 on a Delta II 7623-9.5 ELV vehicle from VAFB, CA.

Orbit: HEO (Highly-elliptical Earth Orbit), a polar orbit of 90º inclination, perigee = 1000 km, apogee = 7.2 RE (45,922 km). The location of the apogee changes during the course of the two-year mission, both in latitude and, because of the Earth's revolution about the sun, in local time.

At the beginning of the mission, apogee is at approximately 40º north geographic latitude and at dusk local time. As the Earth moves around the sun, the plane of the orbit shifts relative to the Earth-sun line (by 30º of longitude each month). During IMAGE's two-year nominal mission, the line of apsides will precess over the pole and return to 40º north geographic latitude. This type of orbit permits the IMAGE instruments to image the inner magnetosphere on nearly a continuous basis. 7)


Figure 4: Illustration of IMAGE orbit rotation over mission time (image credit: NASA)



Operational status of the mission:

• April 10, 2018: IMAGE's signal remains too weak to achieve frame lock, which is necessary to retrieve data from the spacecraft. But important steps have been taken this week to be prepared in case of re-established contact. 8)

- Last week, the engineers at NASA's Goddard Space Flight Center in Greenbelt, Maryland, successfully established network connections with both the antennas at NASA's Wallops Flight Facility in Virginia and at the agency's White Sands Test Facility in Las Cruces, New Mexico. These antennae are now prepared to both command and receive telemetry data from IMAGE, if the spacecraft is re-contacted.

- These preparations are necessary for the team, led by former IMAGE mission director Richard Burley, to attempt to command IMAGE to switch from its current medium gain antennae to its omnidirectional antennae, which has a weaker signal but a wider footprint. The team hopes to be able to lock onto this broader signal, which would lay the groundwork for reestablishing contact, retrieving data and attempting to restore IMAGE to full working capacity.

• On Feb. 22, 2018, the signal from IMAGE began to break up and has been silent since Feb. 24. The team continues to assess what may be the issue, but it is known that this episode does not mimic the sudden silence that occurred in 2005 when contact was originally lost with the spacecraft. The team continues to make preparations to attempt to bring the attitude determination and control systems back online should communications with IMAGE be re-established. 9)

• February 5, 2018: Current information from the IMAGE spacecraft shows that the battery is fully charged, and that overall, the satellite itself seems to be in good shape. The next step is to attempt to turn on the science instruments – but this could take some time as the 12-year-old software to do so must be recreated. Additionally, as computers have evolved greatly in that time, work is being done to find a machine that can run the instrument commanding software. 10)

- During this process of inspecting the spacecraft, there are several puzzles that the team is investigating to better understand the spacecraft's health and how best to communicate with it, including:

a) What caused the spacecraft to reboot and begin sending signal again?

b) Why is one side of the on-board electronics working and not the other? We are currently communicating with IMAGE through the original A side of the on-board electronics. The A side was thought to have failed in 2004, when the communications were switched to the redundant B side. How and why the A side is now working is something we are looking at.

- As we move forward, NASA is starting to recreate a small control center that can generate the commanding needed to better understand and control the satellite. This will then allow us to gain insights into the state of various science instruments, and see whether any are still functional. Should any of the instruments be functional, NASA will convene a panel of external scientists to assess the science potential in the context of constrained budgets for operating spacecraft.

• As of 29 January, observations from all five sites were consistent with the radio frequency characteristics expected of IMAGE. Specifically, the radio frequency showed a spike at the expected center frequency, as well as side bands where they should be for IMAGE. Oscillation of the signal was also consistent with the last known spin rate for IMAGE. 11)

- To confirm beyond doubt that the satellite is IMAGE, NASA will next attempt to capture and analyze data from the signal. The challenge to decoding the signal is primarily technical. The types of hardware and operating systems used in the IMAGE Mission Operations Center no longer exist, and other systems have been updated several versions beyond what they were at the time, requiring significant reverse-engineering.

- If data decoding is successful, NASA will seek to turn on the science payload - currently turned off - to understand the status of the various science instruments. Pending the outcome of these activities, NASA will decide on how to proceed.

January 29, 2018: NASA's Long Dead ‘IMAGE' Satellite is Alive! 12) 13)

- Some 12 years since it was thought lost because of a systems failure, NASA's IMAGE (Imager for Magnetopause-to-Aurora Global Exploration) satellite has been discovered, still broadcasting, by an amateur astronomer. The find, which Scott Tilley reported in a blog post this week, presents the possibility that NASA could revive the mission, which once provided unparalleled views of Earth's magnetosphere.

- "Upon reviewing the data from January 20, 2018, I noticed a curve consistent with an satellite in High Earth Orbit (HEO) on 2275.905 MHz, darn not ZUMA ... This is not uncommon during these searches. So I set to work to identify the source." A quick identity scan using ‘strf' (sat tools rf) revealed the signal to come from 2000-017A, 26113, called IMAGE.

- Meanwhile, Richard Burley of NASA/GSFC confirmed in an email to Tilly that the radio source is compatible with that of IMAGE. "Engineers at GSFC have acquired the suspect S-band source using the 4 m CTA (Compatibility Test Antenna) at GSFC. They acquired the signal while the target was on ascent at about 2RE. The center frequency (CF) was between 2272.478 and 2273.418. The difference between IMAGE documented CF of 2272.5 MHz can be attributed to expected Doppler. Subcarriers are visible as well 1.7 MHz from CF as expected. The signal strength was oscillating. Plots will be forthcoming. The oscillation is not unexpected given IMAGE's loss of spin balance. All indications so far suggest that this is, in fact, IMAGE."

- Stay tuned for further updates. If IMAGE is revived, its orbit will be well positioned to monitor Earth's northern auroral zone.

• The extensive archival database generated by IMAGE promises to yield new discoveries and will support investigations by other spacecraft and ground-based observatories for many years.

• In December 2005, the IMAGE spacecraft ended its operations, bringing to a close a successful mission of 5.8 years. IMAGE was the premier producer of new discoveries on the structure and dynamics of the Earth's external magnetic field (magnetosphere) and its contents.

- IMAGE was launched on March 25, 2000. It successfully completed its two-year primary mission and continued providing data into December 2005, when it stopped responding to commands from ground controllers. Analysis indicated the craft's power supply subsystems failed, rendering it lifeless.14) 15)



Sensor complement: (LENA, MENA,HENA, EUV, FUV, WIC, SI, GEO, RPI)


Major component


Team Lead

LENA (Low Energy Neutral Atom Imager)


NASA/GSFC, Greenbelt, MD

T. E. Moore

MENA (Medium Energy Neutral Atom Imager)


SwRI (Southwest Research Institute), San Antonio, TX

C. Pollock

HENA (High Energy Neutral Atom Imager)


JHU/APL, Laurel, MD

D. Mitchell

EUV (Extreme Ultraviolet Imager)


University of Arizona

W. Sandel

FUV (Far Ultraviolet Imager)

Wideband Imaging Camera

UCB (University of California at Berkeley), MSFC

S. Mende

FUV-SI (Far Ultraviolet Imager)

Spectrographic Imager

UCB, CSL Liege, Belgium

S. Mende

RPI (Radio Plasma Imager)

Antenna Deployers

Able Engineering Corp.

G. Heinemann

RPI (Radio Plasma Imager)

Antenna Couplers

University of Paris, Meudon Observatory

R. Manning

Radio Plasma Imager


University of Mass. Lowell

B. Reinisch

Central Instrument Data Processor (CIDP)



M. Epperly

CIDP Flight Software



R. Killough

HCU (Heater Control Unit)



M. Epperly

Payload Wiring Harness



P. Jensen

Payload Purge System



W. Perry

Table 2: Overview of instrument developers

The IMAGE S/C carries three ENA (Energetic Neutral Atom) imagers whose combined energy coverage permits the detection of ENAs with energies ranging from 1 eV to 500 keV per atomic mass unit (amu). Each neutral atom instrument generates images showing the intensity and spatial distribution of ENA emissions produced in the inner magnetosphere through charge-exchange reactions between geocoronal neutral hydrogen and various magnetospheric ion populations. Neutral atom imaging of the ionosphere and magnetosphere is possible because the Earth's geocorona acts like an imaging screen for magnetospheric and ionospheric ions. The interpretation and quantification of observed ENA signals depend upon the knowledge of the energy dependence and magnitude of the appropriate charge-exchange cross sections. 16)


Figure 5: Schematic view of the instrument layout in the spacecraft spin axis direction (image credit: SwRI)


LENA (Low-Energy Neutral-Atom Imager):

PI: T. E. Moore of GSFC (collaboration of GSFC, University of Maryland, University of New Hampshire, University of Denver, University of Bern, Lockheed Martin Co.). The objective is to detect ENAs in the energy range from 10-500 eV. LENA's primary role is to image the outflow of low-energy ions from the polar ionosphere. The specific objectives are to: 17) 18)

• Measure neutrals without interference from electrons, ions, or UV

• Distinguish neutral protons from oxygen

• Determine ion outflow on five minute time scales over broad range of local times

• Measure energies as low as 10 eV with high counting statistics

The LENA instrument consists of a collimator, a conversion unit, an extraction lens, a dispersive energy analyzer, and time-of-flight mass analyzer with position-sensitive particle detection. LENA is specifically designed of looking at and in the direction of the sun. Principle of operation: Neutral particles enter the instrument through a collimator which filters charged particles. LENA converts neutrals to negative ions through a near specular glancing reflection from a tungsten surface. Negative ions from the surface are then collected by the extraction lens which focuses all negative ions with the same energy to a fixed location. In the extraction lens, the ions are accelerated by 20 kV prior to entering the electrostatic analyzer (they are detected as they pass through a thin foil and strike an image plane detector). Finally, the ions pass into a time-of-flight/position sensing section where ion mass, energy, and angle are determined.

LENA uses electrostatic optics techniques for energy (per charge) discrimination and carbon foil time-of-flight techniques for mass discrimination. The instrument has a 90º x 8º FOV in 12 pixels, each nominally 8º x 8º. The S/C spin provides a TFOV of 90º x 360º, comprised of 12 x 45 pixels.

Energy range (incident neutral atom)

15-1250 eV

Energy resolution

E/ΔE - 1 at FWHM

Mass range

1-20 amu (H+ and O+)

Mass resolution

M/ΔM = 4 at FWHM

Angular coverage

Sampling: 8º x 8º x 12 sectors per spin: 360º x 90º in 45 x 12 samples

Angular resolution/response

8º x 8º = 12 steradians (sr)

TFOV (Total Field of View)

2.8 π sr

Pixels per image

3.2 k = 3 energy x 45 Az x 12 polar x 2 mass

Pixel physical aperture

1.0 cm2 (Aeff <1 x 10-3 cm-2)

Pixel solid angle

0.02 sr per pixel x 12 pixels

Time resolution 1D polar x energy
Time resolution 2D (Az x polar) x energy
Time resolution 3D (Az x polar x energy)

2.7 s
120 s
1 spin period (120 s)

Dynamic range


Instrument mass, power, data rate

20.75 kg, 13.1 W (orbit averaged), 0.5 kbit/s

Table 3: Overview of LENA instrument parameters


Figure 6: Exterior assembly view of the LENA sensor and its Control and Data Handling System (image credit: NASA)


Figure 7: Functional block diagram of LENA (image credit: NASA)


Figure 8: The LENA instrument (image credit: NASA/GSFC)


MENA (Medium-Energy Neutral-Atom Imager):

PI: C. J. Pollock of Southwest Research Institute (collaboration of SwRI, LANL, RAL, USC, UCB, University of West Virginia). Objective: Detection of ENAs in the energy range 1-30 keV. Provision of images of the ring current, near-Earth plasma sheet, and the nightside injection boundary. In addition, MENA images the ion populations of the cusp. The instrument determines the time of flight and incidence angle of the incoming ENAs; from these raw data it calculates their trajectory and velocity and generates images of the magnetospheric regions from which they are emitted.

The MENA instrument (a slit-type imager) is composed of three identical sensors, mounted to a common DPU (Data Processing Unit) assembly. Each sensor is supported by two dedicated FFE (Front End Electronics) cards, time of flight (FEETOF), a pulse height analyzer (FEEPHA), and a triple-function (+10 kV, +4 kV, -1kV) high-voltage power supply. Each of the three sensors provides 1-D imaging of incident ENAs in the polar angle direction based on a spherical coordinate system. The second imaging dimension (azimuth) is obtained using collimation and S/C spin. The sensors are mounted to provide look directions that view a common azimuth (spin) angle, but are offset from one another in their polar angle field of view. The center of the sensor 2 look direction is perpendicular to the spin axis, at a polar angle of 90º. 19)

The MENA instrument has a mass of 13.9 kg, power consumption = 22.5 W, data rate = 4.3 kbit/s, size: 42.5 cm x 22.3 cm x 29.3 cm.


Figure 9: Block diagram of the MENA instrument (image credit: SwRI)


Figure 10: Isometric drawing of the MENA imager (image credit: SwRI)


Figure 11: The MENA flight instrument prepared for integration (image credit: SwRI)

ENAs, charged particles and photons incident from within the sensor's FOV enter through the collimator, where charged particles with energies up to 13 times the adjustable applied voltage are removed by electrostatic deflection. The remaining particles and photons must pass through a free-standing UV blocking grating, where the UV photons are removed around a very wide stop band by the optical properties of the grating. The grating structure eliminates the solar hydrogen Lyman-α (121.6 nm) light reflected from the geocorona. The detector system uses MCP detector arrays (The detector assembly consists of a Hamamatsu MCP and an anode that employs a novel "capacitive charge division" technique to determine the position of the ENA impacts on the MCP).

The TOF for each ENA detected is determined by the FEE from the start and stop pulses triggered in the MCP. This value, together with the positions of the start and stop pulses on the MCP, is processed by the look-up tables to compute the incidence angle of an incoming ENA, the length of its path through the detector, and its velocity. The amplitude of the start and stop pulses i.e. their "pulse height," is also measured. Knowledge of both the spin phase and the incidence angle is needed to determine the position in the sky from which the detected ENAs are emitted and to produce the image of the ENA emission region.


HENA (High-Energy Neutral-Atom Imager):

PI: D. G. Mitchell of JHU/APL [collaboration of APL, University of Maryland, University of Arizona, MPAe (Katlenburg-Lindau, Germany)]. Detection of ENAs in the 10-500 keV energy range. HENA imaging is focused principally on the ring current, inner plasma sheet, and substorm injection boundary. HENA is a modified version of the Cassini INCA instrument, it provides global images of ENA emissions from Saturn's magnetospheric ion populations. The two main HENA components are the sensor and the main electronics unit (MEU)

The HENA sensor consists of alternately charged deflection plates mounted in a fan configuration in front of the entrance slit; three microchannel plate (MCP) detectors; a solid-state detector (SSD); two carbon-silicon-polyimide foils, one at the entrance slit, the other placed just in front of the back MCP; and a series of wires and electrodes to steer secondary electrons ejected from the foils (or the SSD) to the MCPs. 20) 21) 22)


Figure 12: Schematic view of the HENA head (image credit: JHU/APL)


Figure 13: Block diagram of the HENA instrument (image credit: JHU/APL)


Figure 14: Illustration of the HENA instrument (image credit: SwRI)

HENA determines the velocity of the ENAs that it detects by measuring their time of flight (TOF) and trajectory through the sensor, i.e., from the entrance slit either to the back foil and 2-D imaging MCP detector or to the SSD.

Wavelength, bandpass

30.4 nm, 5 nm


0.2 counts/s per Rayleigh per pixel

Field of View (FOV)

30º by 84º (instantaneous), 84º by 360º (total)

Spatial resolution

640 km x 640 km at apogee (8 Earth radii)

Imaging frequency

1 image accumulated every 10 min (= 5 S/C rolls)

Instrument mass, power

15.6 kg, 15.5 W

Thermal range

-25 to +40º C (operating), -50º to +60º C (non-operating)

Table 4: EUV instrument parameters


EUV (Extreme Ultraviolet Imager):

PI: B. R. Sandel, University of Arizona. The objective is to detect solar EUV photons with a wavelength of 30.4 nm that are resonantly scattered by singly ionized helium in the plasmasphere, the torus of cold dense plasma surrounding the Earth in the inner magnetosphere. A sophisticated computer deconvolution technique is used to translate the EUV photon counts registered by the instrument into images of the plasmasphere. detect solar photons. 23) 24) 25)

The imager consists of three identical sensor heads mounted one above the other in a common bracket; a common electronics module. Each sensor head has a FOV of 30º, the three sensors are tilted so that their FOVs overlap, giving the imager a fan-shaped IFOV of 30º by 84º. With each rotation of the S/C, the imager completes a 360º sweep of the sky, resulting in a TFOV of 84º by 360º. Spatial resolution is about 0.6º at apogee (EUV can distinguish plasma-spheric features with scale sizes down to about 640 km). This resolving power enables the study of fine-scale density structures.

The EUV instrument has a mass of 15.5 kg, power consumption of 9.0 W, size of 49.7 cm x 23.3 cm x 49.5 cm.


Figure 15: Block diagram of the EUV Imager (image credit: University of Arizona)


Figure 16: Front view of the EUV Imager with sensor heads (image credit: University of Arizona)


Figure 17: Photograph of the EUV curved surface flight MCP sensor (image credit: University of Arizona)

The detector design uses a novel photon counting detector scheme with a spherically curved MCP (Microchannel Plate) stack and wedge and strip readout for the EUV instrument. The system includes two main elements, the detector and the detector electronics. The detector consists of a Kovar-alumina brazed body assembly containing the MCPs and readout anode. The detector electronics consists of three boards, amplifier, ADC and interface electronics, necessary to encode photon event data. Incoming photons strike the bare surface of the MCP, where they may eject photo-electrons, which are multiplied and deposited onto a wedge and strip anode. The charge is proportionately divided among the wedge, strip and zigzag electrodes, and then the signals are transferred to the detector electronics. Each of the three signals is amplified and converted to digital form before being sent to the EUV controller and subsequently converted to X,Y photon positions.


FUV (Far Ultraviolet Imager):

FUV is designed and built by UCB/SSL (University of California at Berkeley/Space Sciences Laboratory). The FUV instrument employs three detectors: 26) 27) 28)

• The Wideband Imaging Camera (WIC) to image the aurora in broad band for maximum spatial resolution day and night

• The Spectrographic Imager (SI) to measure different types of aurora and separating them by wavelength and to measure proton induced aurora by removing the bright geocorona emissions

• Geocorona photometers (GEO) to observe the distribution of the geocorona emissions to derive the magnetospheric hydrogen content responsible for neutral atom generation in the magnetosphere.

1) WIC (Wideband Imaging Camera). The WIC instrument is designed to image the whole Earth and the auroral oval from satellite distances greater than 4 Earth radii to the center of the Earth. A curved image intensifier is optically coupled to a CCD and the optics provides a FOV of 17º x 17º. Spectral range: 140-160 nm; resolution elements of < 0.1º; temporal resolution of 120 s; the size of the final images is 256 x 256 pixel elements, corresponding to spatial resolutions of < 100 km at apogee distances. 29)


Figure 18: Schematic view of the WIC instrument (image credit: UCB/SSL)


Figure 19: Illustration of the WIC flight unit (image credit: UCB/SSL)

2) SI (Spectrographic Imager). SI was largely designed, tested and calibrated at CSL (Centre Spatial de Liège) and LPAP (Laboratoire de Physique Atmosphérique et Planétaire) of Liege, Belgium in the frame of the PRODEX program of ESA. The objective is to image the whole Earth proton aurora from satellite distances greater than 4 Earth radii to the center of the Earth. It uses a reverse Wadsworth design to select the Doppler shifted Lyman H-alpha line at 121.82 nm in the ultraviolet part of the optical spectrum and to reject the non-Doppler shifted Lyman H-alpha from the geocorona at 121.567 nm. The FOV is 15º x 15º. The temporal resolution between two images is 120 s and the size of the final images is 128 x 128 pixel elements, corresponding to spatial resolutions of < 100 km at apogee distances. 30)

SI detector system: A set of MCP (Microchannel Plate) detectors was developed utilizing a cross delay-line readout system for the FUV Spectrographic Imager. Two detectors are required for the two pass bands. Both detectors are nearly identical, the only difference being the position of the input window on the detector cover plate. Each detector, optimized for operation in the far ultraviolet with a KBr photocathode, provides high spatial resolution and good linearity over a 20 mm2 format.

Active area (photosensitive area)

20 mm x 20 mm

Sensitivity (QE)

20%-121.6 nm, 10%-135.6 nm (w/o window)

Response uniformity


Digital pixels at the time delay converter

1024 x 1024 super samples for TDI scanning


±50 µm

Detector resolution

60 µm

Maximum count rate (system)

105 counts/s

Maximum single pore rate

1 count/s

Background state

2 counts/s/cm2

Table 5: Performance requirements of the SI detector


Figure 20: View of the SI instrument with its cover off (image credit: UCB/SSL)

3) GEO (Geocorona Photometer) is designed to measure the hydrogen Lyman-alpha emission of the neutral atmosphere. The three photometers have a field of view of 1º x 1º and look into three different directions perpendicular to the spin axis of the satellite and tilted by ±28º. They are ultraviolet photon counters with filters giving the desired bandpass (121.6 nm). An additional O2 gas cell provides an excellent throughput of the hydrogen radiation and simultaneously rejects the photons at 130.4 nm from excited neutral oxygen. The temporal resolution between two complete measurements around 360º is 120 s.


Figure 21: View of the GEO instrument with its side cover removed to show the proton tube (image credit: UCB/SSL)


Figure 22: Single GEO detector tube with built in MgF2 lens and aperture to limit the FOV to 1º (image credit: UCB/SSL)


RPI (Radio Plasma Imaging):

PI: B. W. Reinisch, University of Massachusetts at Lowell. The objective is to characterize plasma in the Earth's magnetosphere utilizing imaging in the radio frequency range. The RPI instrument is a low-power radar which operates in the radio frequency bands which contain the plasma resonance frequencies characteristic of the Earth's magnetosphere (3 kHz to 3 MHz). RPI can locate regions of various plasma densities by observing radar echoes from the plasma that are reflected where the radio frequency is equal to the plasma frequency. By stepping through various frequencies for the transmitted signal, features of various plasma densities can be observed and, by fitting contours and/or magnetospheric models to the features, a 3-D specification of the shape of the magnetosphere can be created. 31) 32)


Figure 23: Block diagram of the RPI (image credit: SwRI)

The RPI instrument consists of an electronics enclosure, four 250 m wire antennae with deployers (including switches and couplers), and a z-axis boom canister containing two 10 m lattice boom antennae and two preamplifiers. Instrument mass = 49.8 kg, power = 133.98 W (peak) and 30.8 W (average). 33) 34)


Electron density range (cm-3)

Plasma frequency range (kHz)

Magnetopause boundary layer



Polar cusp













8000-1 x 105


Table 6: Representative RPI targets


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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: "Observation of the Earth and Its Environment: Survey of Missions and Sensors" (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates(

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