ICON (Ionospheric Connection Explorer) Mission
In April 2013, NASA selected a new satellite mission and a new space-based instrument to begin development as part of the agency's Heliophysics Explorer Program. The projects will provide space observations to study Earth's ionosphere and thermosphere. NASA Selects Explorer Projects To Probe Earth's Upper Atmosphere 1) 2) 3) 4)
NASA awarded UCB (University of California, Berkeley) a contract to build a satellite to determine how Earth's weather affects weather at the edge of space, in hopes of improving forecasts of extreme "space weather”. ICON will be designed, built and operated by scientists at UCB/SSL (Space Sciences Laboratory). 5) 6) 7)
Science: ICON will explore the boundary between Earth and space – the ionosphere – to understand the physical connection between our world and the immediate space environment around us. This region, where ionized plasma and neutral gas collide and react exhibits dramatic variability that affects space-based technological systems like GPS. The ionosphere has long been known to respond to “space weather” drivers from the sun, but recent NASA missions have surprised us in showing this variability often occurs in concert with weather on our planet. ICON will compare the impacts of these two drivers as they exert change on the space environment that surrounds us (Ref. 2).
We've known for many years that the Sun has a significant effect on this region of the upper atmosphere. As the Earth rotates, a new portion of its atmosphere is exposed to solar ultraviolet radiation, which heats and partially ionizes the neutral atoms and molecules, creating the ionosphere. At night, when that portion of the atmosphere is not exposed to the Sun, this ionized gas, or “plasma” tends to recombine, dramatically reducing its density as it converts back to an electrically neutral state. In between, near sunset, the low-latitude ionosphere is dominated by a “plasma fountain”, which results in a dramatic upwelling of ionized gas. This results in a sharp increase of the density of ionized gas in narrow bands on either side of the magnetic equator. Basic models of the ionosphere predict that there should be a regular rise and fall of this portion of the atmosphere as the Earth rotates from day to night, independent of longitude. 8)
Figure 1: Left: The predicted distribution of plasma around the magnetic equator after sunset. Right: The observed distribution of plasma around the magnetic equator, made by NASA's TIMED spacecraft. Notice the large, unexplained enhancements over the continents (image credit: UCB/SSL)
Understanding what is causing the variations in the atmosphere is very important in the technological era we live in today. As a society, we are very dependent on communication and navigation networks around the globe – both space based, and ground based. We have also recently developed a strong reliance on broadcast navigation signals such as those provided by GPS satellites. The radio signals used for communication and navigation must propagate through the ionosphere, and non-uniform distributions of plasma in the ionosphere can act like bubbles in a lens or scratches in a mirror, distorting the signal, sometimes to the point of unintelligibility or unusability.
Though the solar inputs are now well quantified, the drivers of ionospheric variability originating from lower atmospheric regions are not. ICON is the first space mission to simultaneously retrieve all of the properties of the system that both influence and result from the dynamical and chemical coupling of the atmosphere and ionosphere. ICON achieves this through an innovative measurement technique that combines remote optical imaging and in situ measurements of the plasma. With this approach, ICON provides the ability to:
1) separate the drivers and pinpoint the real cause of ionospheric variability
2) explain how energy and momentum from the lower atmosphere propagate into the space environment
3) explain how these drivers set the stage for the extreme conditions of solar-driven magnetic storms.
ICON’s imaging capability combined with its in-situ measurements on the same spacecraft (Figure 2) gives a perspective of the coupled system that would otherwise require two or more orbiting observatories.
ICON targets the low-latitude ionosphere because recent global-scale observations of this region show remarkable spatial and temporal variability that contravene the conventional view of ion-neutral coupling in space, and evince strong forcing by lower atmosphere drivers. The coupling of the atmosphere to space is strongest at these latitudes because the atmospheric waves are largest and so is the density of the space plasma, produced in abundance by the sun overhead and confined by the magnetic field.
The ICON mission is lead by Thomas J. Immel (PI) of UCB/SSL. The NASA Heliophysics Explorer mission, ICON, follows the PI-management mode of operation. The mission management is performed by UCB/SSL under the NASA/GSFC (Goddard Space Flight Center) Explorers program.
Orbital will design, manufacture and integrate the ICON spacecraft under contract from UCB/SSL. The LEOStar-2 platform of Orbital ATK was selected for the ICON spacecraft which is a flexible, high-performance spacecraft for space and Earth science, remote sensing and other applications. The LEOStar-2 series spacecraft have supported multiple missions for commercial and government customers over the past 15 years. ICON will be the ninth LEOStar-2-based spacecraft built by Orbital. 9) 10) 11)
ICON is a relatively straightforward application of the existing LEOStar-2 line. Attitude control is provided with a zero momentum bias system that includes actuators sized to perform all of ICON’s attitude maneuvers for science. Two star trackers and an inertial measurement unit provide the information necessary to keep the instrument suite continually pointing at the Earth’s limb as the observatory orbit the planet. This highly agile spacecraft generates ample electrical power using a two wing solar array design. A passive thermal control system maintains the spacecraft and instruments in their operational ranges.
Orbital’s flight experience with single-string spacecraft such as ICON has been excellent, with most previous missions operating past their design lifetimes. The spacecraft also features selective component redundancy and an autonomous, flight proven fault management system to enhance reliability. The impressive past performance of the LEOStar-2 combined with high design margins early in development really helps ICON start with a low and manageable risk profile.
Figure 3: Illustration of the deployed ICON minisatellite (image credit: Oribital ATK)
Table 1: Overview of spacecraft parameters
Project development status:
• On Oct. 4, 2018, technicians at Vandenberg Air Force Base in California, installed the payload fairing on the Northrop Grumman Pegasus XL rocket that will launch NASA’s Ionospheric Connection Explorer, or ICON, satellite. 12)
Figure 4: Installation of the payload fairing on the Northrop Grumman Pegasus XL rocket (image credit: NASA)
- ICON is being prepared for launch on a Pegasus XL rocket which will be carried aloft by Northrop Grumman’s L-1011 Stargazer aircraft. The Stargazer with the Pegasus XL attached is scheduled to fly from Vandenberg, where it was processed, to the Skid Strip at Cape Canaveral Air Force Station in Florida on Oct. 19, 2018.
- Launch of the Pegasus XL rocket is scheduled for Oct. 26, 2018. The Stargazer jet will take off from the Skid Strip at the Cape. About 50 miles offshore of Daytona Beach, Florida, the Pegasus XL will be dropped with the engine igniting five seconds later boosting ICON to orbit. The Stargazer is a mobile launch platform and the only one of its kind in the world.
Figure 5: This illustration depicts NASA's ICON in space (image credit: NASA)
• On June 6, 2018, Northrop Grumman Corporation announced it has closed the acquisition of Orbital ATK Inc. (“Orbital ATK”), a global leader in aerospace and defense technologies. Orbital ATK is now Northrop Grumman Innovation Systems, a new, fourth business sector. 13)
• November 3, 2017: NASA is postponing launch of the Ionospheric Connection Explorer (ICON) until 2018. The mission was previously planned to launch Dec. 8, 2017, on an Orbital ATK Pegasus XL rocket from the Reagan Test Site on Kwajalein Atoll in the Marshall Islands. NASA and Orbital ATK need additional time to assess a separation component of the rocket. More information on a revised launch date will be provided once it becomes available. 14)
• April 7, 2016: All instruments were delivered in March to the SDL (Space Dynamics Laboratory) at USU (Utah State University) in Logan, where the instruments will be connected to the Payload Integration Plate — the common body of the spacecraft that will house the science payload. The integration and testing procedures are a vital part of preparing for flight, ensuring that they will function together as planned. 15)
- The I&T (Integration and Testing) processes for the ICON payload at USU/SDL are a vital part of preparing for flight. In order to ensure that the instruments will be able to function together as planned, they are brought together on the PIP (Payload Integration Plate). The PIP, the instruments, and the Instrument Control Package together become the science payload, which will then undergo a series of thorough vibration and thermal tests at SDL over the coming months. 16)
- Following I&T, the payload will be shipped from Utah to Orbital ATK in Virginia, where it will be integrated onto the main spacecraft “bus” - the guts of the satellite that controls communication, attitude, and other overall controls. This will happen towards the end of 2016, in preparation for launch in summer 2017.
• In November 2014, NASA has officially confirmed the ICON mission, clearing it to move forward into the development phase. 17)
• In July 2014, the ICON project passed its PDR (Preliminary Design Review). 18)
• On January 14, 2014, the ICON project passed the SRR (System Requirements Review). 19)
• In April 2013, UCB/SSL was selected to build NASA's next space weather satellite. 20)
Figure 6: Artist's rendition of the deployed ICON spacecraft. Charged particles in the ionosphere create bands of color called airglow (image credit: NASA)
Launch: A launch of the ICON mission is scheduled for 26 October 2018 aboard a Northrop Grumman Pegasus XL launch vehicle. The launch site is Cape Canaveral Air Force Station in Florida. ICON will be launching off the coast of Daytona from former Orbital's "Stargazer" L-1011 aircraft at 39,000 ft. (11.9 km) at a heading of 105.0 degrees. 21)
Already in November 2014, NASA had selected Orbital Sciences Corporation of Dulles, Virginia, to provide launch services for the Ionospheric Connection Explorer (ICON) mission. ICON is targeted to launch in June 2017 from the Reagan Test Site on Kwajalein Atoll in the Republic of the Marshall Islands aboard a Pegasus XL launch vehicle from Orbital's "Stargazer" L-1011 aircraft. 24)
Orbit: Circular orbit, altitude = 575 km, inclination = 27º.
Figure 7: The ICON satellite will orbit above the upper atmosphere – higher than the International Space Station – to observe how interactions between terrestrial weather and the ionosphere create bright, shimmering patches of color called airglow, as well as other changes in the space environment (image credit: NASA/GSFC)
Sensor complement: (MIGHTI, FUV, EUV, IVM)
The sensor complement consists of four instruments: MIGHTI of NRL, FUV and EUV, built at UC Berkeley, and IVM built at the University of Texas in Dalles.
The instruments all are mounted on the PIP (Payload Interface Plate). It provides a single structural interface to the spacecraft bus (main body of the satellite), through three bipod legs. The ICP (Interface Control Package), which provides a common source of control over the instruments' power, heaters, detector modes, and all mechanical operations. The PIP mates to the Spacecraft, and provides a mounting area for ICON's two Star Trackers, for continuous determination of the observatory attitude, a 3-axis magnetometer, which also is used for attitude determination, and the S-band antenna providing a bi-directional communication path to the ICON ground network.
Figure 8: CAD drawing of the ICON sensor complement (image credit: UCB/SSL, Orbital)
The ICP is the primary interface between the Spacecraft avionics systems and the instrument payload. The ICP manages the distribution of power to each of the ICON instruments, and it manages the flow of data from the instruments that is routed to the spacecraft solid state recorder. Commands are sent to the ICP to control the mode of the instruments. For example, the ICP controls the apertures in the MIGHTI optical systems, and it controls the FUV turret position. All of the payload heaters are controlled by the ICP. The FSW (Flight Software) for the payload resides on the DCB (Data Control Board) which is one board in the ICP. The FSW is programmed to manage all facets of instrument operation, whether it's related to control, or data processing. 25)
Each of the ICP's four electronics boards is assembled at the SSL (Space Sciences Laboratory), in accordance with NASA approved procedures and inspection criteria.The ICP integrates all the capabilities necessary to run all 4 instruments into one unit. ICP contains all the following boards:
• Digital Control Board DCB with (2 MB MRAM)
• Provides all capability to command instruments and accept all sensor data
• Power Control Board PCB
• Provides all power to instruments, instrument heaters, and motors
• TPS (TEC Power Supply) for MIGHTI
• Power supply specifically to drive the TEC cooler for MIGHTI
• LCPS (Low Voltage Power Supply)
• Provides secondary power to instrument cameras and survival heaters
Figure 9: Schematic block diagram of the ICP (image credit: UCB/SSL)
Figure 10: Photo of the ICP (Instrument Control Package), image credit: UCB/SSL
MIGHTI (Michelson Interferometer for Global High-resolution Thermospheric Imaging)
The MIGHTI instrument is provided by NRL (Naval Research Laboratory) with Christoph Englert as PI. The primary objective for MIGHTI is to determine thermospheric winds and temperatures. Wind measurements will be made between 90-300 km altitude (day) and between 90-105 km and 200-300 km (night) with a horizontal resolution of <500 km. Thermospheric temperature measurements are required for the 90-105 km range. These mission requirements drive the design of the instrument. 26) 27) 28)
The wind information is determined from the Doppler shift of atomic oxygen airglow emission lines and the temperature is determined from the spectral shape of the O2 A-band emission. These naturally occurring emissions are present day and night and MIGHTI observes the red-line (O(1D)@630.0 nm) at heights near the peak in the ionospheric density (200-300 km altitude) resulting mainly from electron recombination and impact reactions with O2+ and O, and the green-line (O(1S) @ 557.7 nm) at the base of the thermosphere (~100 km altitude) originating mainly from three-body collisions (O+O+M) and subsequent energy transfer. Furthermore, bright, temperature sensitive near-infrared band emissions (0-0 band of O2(b1Sg+) @762.0 nm) in the 100 km altitude region are used to retrieve the temperature. Recent imaging with a handheld Nikon camera show how bright these emissions are at night (Figure 11). MIGHTI employs a high performance baffle so that these emissions can be observed during daytime as well.
Legend to Figure 11: The 762 nm O2 band emission (false color in image) provides temperatures, while the Doppler shift of the 557 nm (green) and 630 nm (red) OI emission lines provides wind profiles.
For the wind measurement, MIGHTI uses an interferometric technique that is based upon the proven approaches used for the WINDII (Wind Imaging Interferometer) flown on the NASA UARS satellite (launch on Sept. 12, 1991) and the SHIMMER (Spatial Heterodyne Imager for Mesospheric Radicals) instrument, flown on STPSat-1 (launch on Sept. 3, 2007). Figure 12 (left panel) shows the monolithic interferometer employed by SHIMMER. WINDII used the same airglow emission lines to measure Doppler wind in the same altitude range as required for ICON. The added capability based upon SHIMMER space flight heritage allows for a wind measurement virtually identical to WINDII and eliminates moving interferometer parts. The MIGHTI design utilizes a cubic beam splitter (Figure 12 : right panel) that allows for a smaller and lighter interferometer and and a more compact optical bench layout, while maintaining excellent optical performance.
The MIGHTI instrument consists of two sensor units with orthogonal fields of view, pointed 45° and 135° from the S/C velocity direction toward the port (northern) side of the spacecraft. With this viewing geometry, MIGHTI makes two perpendicular line-of-sight wind measurements of the same air volume less than 8 minutes apart. Each measurement represents a set of limb observations for tangent altitudes between 90 and 300 km. The vector combination from the two perpendicular lines of sight provides an altitude profile of wind vectors. Temperature measurements are performed coincident with the wind measurements using multi-color band shape measurements.
Conventional Michelson interferometers require mechanical stepping of a mirror to sample four or more path differences. The SHS ( Spatial Heterodyne Spectroscopy) approach for MIGHTI results in an improved, more rugged design, replacing the Michelson mirrors with fixed, tilted diffraction gratings. Fundamentally, each facet or groove of the tilted gratings can be regarded as a separate interferometer mirror, each representing a unique path difference, which permits the sampling of many path differences without moving interferometer parts. This eliminates the need for precision mirror stepping, which simplifies both the instrument development and on-orbit operation. The SHS technique was demonstrated by both SHIMMER and the REDDI (Redline DASH Demonstration Instrument), two of the heritage instruments. The basic design of one of the two MIGHTI optical units is shown in Figure 13.
Because the measured interferogram is comprised of straight fringes (rather than circular fringes produced by a Fabry-Perot Interferometer, for example), the MIGHTI field of view can cover the entire limb altitude range required for the ICON science without a scanning mechanism (Figure 13: right panel). This eliminates a scanning mechanism and minimizes complexity in development and during on-orbit operations.
The temperature of the lower thermosphere is derived from measurements of the band shape of the bright oxygen “Atmospheric Band” around 762 nm. Most recently, this technique was applied using data from the OSIRIS (Optical Spectrograph and Infrared Imaging System) instrument of CSA (Canadian Space Agency), flown on the Odin mission of Sweden (launch on Feb. 20, 2001 and operational in 2013). Similar measurements were also reported from the RAIDS (Remote Atmospheric and Ionospheric Detection System) instrument of NRL and The Aerospace Corporation, flown on the JEM/Kibo-EF (Exposed Facility) of the ISS (launch on Sept. 10, 2009).
Legend to Figure 13, left panel: Light enters though the baffle, and is modulated by a temperature stabilized interferometer. A cooled CCD camera detects the fringe patterns to retrieve the wind profiles and the multispectral band signals to retrieve the temperature profiles. Right panel: Modeled airglow and calibration source signals on the MIGHTI detectors.
Two discrete MIGHTI units, their views separated by 90º, are mounted on the ICON Payload Interface Plate. Because the spacecraft motion introduces a different offset to the Doppler shift of the emissions across the horizontal imaging plane, a correction for this velocity differential is an important part of the wind retrieval.
Table 2: MIGHTI instrument characteristics
Detectors: e2v will provide two 2048 x 4096 pixel CCDs (Charge Coupled Devices) image sensors to equip the instrument. The sensors will be back-illuminated, coated with an anti-reflective coating to improve their sensitivity and operated in the frame transfer mode. 29)
FUV (Far Ultraviolet Imager)
The FUV imager is provided by UCB/SSL (PI: Stephen Mende). The primary objective for the FUV imager is to determine daytime thermospheric composition and nighttime ion density through imaging the limb and sub-limb of the ionosphere-thermosphere. 30)
During the day, photoelectrons colliding with atmospheric neutrals, N2 and O, produce emissions and by observing the limb brightness of the N2 LBH (Lyman-Birge-Hopfield) bands and of OI (Oxygen Iodine) at 135.6 nm, the density ratio of the neutral N2 and O atmospheric constituents can be retrieved.
At night the recombination of O+ ions with ionospheric electrons also creates OI 135.6 nm emissions and the nighttime electron density can be estimated from the limb brightness of 135.6 nm. ICON FUV will measure the altitude distribution of the OI 135.6 nm and N2 155 nm LBH dayglow emissions and the altitude and spatial distribution of the OI 135.6 nm nightglow emissions. These quantities can then be used to determine dayside O and N2 densities and altitude profiles, and the nightside O+ densities in the F-region.
FUV instrument: The ICON FUV instrument is based upon the IMAGE FUV-SI ( Spectrographic Imager), flown on the NASA IMAGE mission (launch on March 25, 2000).
Figure 14: The ICON FUV instrument was adapted from the highly successful FUV Spectrographic Imager (SI) flown on the NASA IMAGE satellite (image credit: UCB/SSL)
The FUV instrument is an off-axis Czerny-Turner spectrometer, tuned for observations at 135.6 and 155.0 nm. Upon entrance into the first slit, the scanning mirror sends the light toward a collimating first mirror, reflecting the light from the scene onto the FUV grating. The FUV light is separated at the grating and refocused onto the two exit slits by a secondary mirror. The two intensified CCD detectors capture the images of OI and N2 light. The image of the scene is preserved.
The project selected this grating type of spectrographic “filter” because it minimizes contamination by out-of-band light leaks that are typically a problem for non-grating type FUV cameras and the spectrograph also has the spectral resolution to reject unwanted contamination by the bright optically thick 130.4 nm O line. FUV is nominally pointed 20º down towards nadir from the -Y axis of ICON to image the northern limb and sub-limb. The ICON FUV instrument is illustrated in Figure 15.
A small (40 mm x 12 mm) motorized steering mirror is placed in front of the entrance slit to allow viewing approximately in the plane of the magnetic field. The mirror has several fixed positions selectable by the instrument processor. As in the IMAGE heritage instrument, a sun sensor covering the full steering mirror range will command the safing of the high voltages to prevent instrument damage due to accidental direct sun exposure. Light reflected by the steering mirror passes through a 6 mm x 32 mm entrance slit and is collimated by the first concave mirror. Mirror 1, besides acting as the spectrograph collimator, also focuses the scene viewed by the instrument near the grating, producing an intermediate image. The 4300 groves/mm grating is slightly convex with a 1.7 m radius of curvature separates the O line (illustrated in red) from the N2 bands (illustrated in blue). Mirror 2, in combination with the back imager camera mirrors, located behind each of the two exit slits, reproduce the intermediate image on each detector. There is an exit slit and the back imager camera consists of a de-centered inverse Cassegrain configuration for each of the two bands.
Table 3: FUV instrument characteristics
The two detectors for the ICON FUV instrument are MCP (Microchannel Plate) intensified CCD detectors with an active area of 25 mm (circular) coupled to a 1k x 1k CCD similar to IMAGE WIC. The CCD is a Dalsa FTT1010M, which is the same CCD used for the ISUAL (Imager of Sprites and Upper Atmospheric Lightning) instrument flown on the FormoSat-2 mission (launch on May 20, 2004). It will be binned 4 x4 to achieve the desired size of 256 x 256. Adjacent 4 pixels will be co-added to support the final 64 x256 format, easily meeting the detector spatial requirement. Incident FUV photons strike the CsI photocathode of the sealed Photek image tube which is a modification of the Photek tube flown on the NASA TIMED mission (launched December 07, 2001) in the GUVI (Global Ultraviolet Imager) and in the DMSP SSUSI (Special Sensor Ultraviolet Spectrographic Imager) instrument (launched October 18, 2003). The photoelectrons are multiplied by the 6 mm pore MCP with an overall gain set to a value controlled by the voltage applied across it. The gain is chosen to optimize the dynamic range and sensitivity of the detector, while conferring relative immunity to saturation when viewing bright dayglow. The MCP output electrons are proximity focused onto an aluminized phosphor screen deposited on the output fiber optic window producing visible emissions imaged by the CCD bonded to the fiber optic.
Using the TIMED GUVI limb data to evaluate the viewing conditions one may predict a scene in the FUV FOV with a mean brightness of 5 kR. With this target brightness and detector, the dual MCP electron gain can be set to a moderate level of ~5000, leading to a longer design lifetime (vs. the lifetime of a 106 gain system for XDL detectors). This approach provides operational flexibility by setting the electron gain to optimize dynamic range during day/night operations while the system operates close to the photoelectron shot noise limit.
The FUV sensor package includes the CCD drivers, the amplifiers, and the A/D converters. Regulated power is derived from the ICP low voltage power supply (LVPS) and serial data is produced for the DICB (Digital Imaging Control Board) in the ICP. Each detector includes two high voltage power supplies, one for the MCP (0-2500 V) and one for the phosphor (0-5000 V). The HV power supplies are mechanically packaged with the intensifiers. In operation, HV commands use a two-step protocol, that requires a digital key to enable the operation. A sun sensor provides direct override of the HV, in the event of sun exposure. During ground testing, an HV enable plug, only installed when HV operation is appropriate, provides an additional level of operator error protection. This system has been used in many similar HV applications, e.g. on several spacecrafts e.g. IMAGE, FormoSat-2 (ISUAL) and THEMIS.
The steering mirror has a horizontal freedom of motion allowing a ±20° movement of the 24° FOV. This allows the FOV to be directed toward the local magnetic field for nighttime imaging observations along the magnetic field. The mirror can be used to block the sun when the instrument is commanded to safe hold. The mirror is driven by a stepper motor drive system that was flown in the ISUAL instrument on FormoSat-2.
Figure 16: Photo of the FUV spectrograph during a test at Lockheed Martin in August 2015 (image credit: NASA, UCB)
EUV (Extreme Ultraviolet Imager)
The primary objective for the EUV profiler is to determine the altitude profile of the dayside ionospheric density through limb imaging in the extreme ultraviolet range. 31)
One of the most prominent terrestrial airglow features is the OII 83.4 nm triplet emission. The strongest “initial source” of photons is solar photoionization of neutral atomic oxygen around 160-175 km. The Sun also emits 83.4 nm photons, serving as a secondary source whose impact on the dayside airglow increases with altitude above the ionospheric peak. The 83.4 nm photons propagating upward from the initial source or downward from the Sun undergo resonant scattering by O+, the dominant positive ion within the F-region, and capture the shape of the F-region ion distribution. This is illustrated in Figure 17.
This EUV instrument follows the heritage of four UCB “EURD-class” astrophysical space flight instruments. These astrophysical instruments were designed for detecting the same wavelengths as the ICON-EUV instrument, but for signals that are an order of magnitude less bright than Earth’s airglow. The interpretation of the EUV dayglow builds on the recent success of the SSULI (Special Sensor Ultraviolet Limb Imager) and RAIDS (Remote Atmospheric and Ionospheric Detection System) missions.
The ICON EUV spectral profiler is illustrated in Figure 18 . The EUV instrument measures these emissions with a straightforward grating spectrograph that provides spectral information in one dimension (horizontal) on the 2D detector, and altitude distribution of the emission brightness in the other dimension (vertical).
The EUV uses a simple “vertical pushbroom” one-dimensional imaging spectrograph of UCB design, consisting of a single diffraction grating that directly views a wedge of sky through a fixed slit aperture. The large difference in the vertical and horizontal curvature of the grating (toroidal design) permits the simultaneous vertical focusing of the scene located at great distances outside of the instrument and the horizontal focusing of the input slit located at the instrument entrance on the detector. The view wedge is dispersed and imaged into a spectrum per field angle on a photon-counting detector as shown in Figure 18. The simplicity of a single-optic design results in a highly efficient instrument and eases alignment and calibration issues compared to multi-optic designs.
The ICON EUV features an enclosed aluminum structural cavity withstanding 1 atm. vacuum that acts as the optical bench and contains the diffraction grating and the open-faced MCP (Microchannel Plate) detector. The instrument cavity can be hermetically sealed with a one-time operable flap valve actuated in flight by a shape memory alloy mechanism. A FOV (Field of View) entrance baffle extends beyond the optical entrance to eliminate solar panel glints and minimize sun-pointing constraints. To reject entrance of low-energy ions into the optical cavity that could be accelerated into the MCP detector and cause excessive background noise, the hermetic optical cavity uses a low-voltage electric field applied to the slit and to fine metal grids placed over the entrance baffle, the space-venting aperture, and the detector face inside the cavity.
The horizontal field (12º) and the instrument scale set the ruled width of the grating to 40 mm. A 40 mm tall entrance slit, the imaging-direction pupil, and the imaging angle (16º) sets the 90 mm grating height. The EURD-type optical theory, given a 3000 l/mm ruling, estimates the toroidal figure of the grating (R=176 mm, ρ=336 mm), incident angle (13.7º), and the optimum slit-to-grating distance and detector-to-grating incident angles (177 and 171 mm, respectively). Numerical ray tracing shows that the spectral resolution requirements allow a slit width of 890 µm which yields the per pixel geometric factor to 6.3 x 10-4 cm2 sr while maintaining a clean separation between 58.4 nm and 61.7 nm.
The glass diffraction grating uses straight holographic rulings blazed for 61.7 nm. Similar to one of the previous EURD-class instruments, SPEAR, the grating surface will be coated with chromium for ruling acceptance testing, and then over coated with boron carbide (B4C). Like SPEAR, the grating is adhesively mounted to a titanium block that interfaces to the structure via a compliant three-point mount. The points use spherical bearings on threaded mount rods to allow for remote, mechanized, optical alignment in an UV vacuum test chamber during ground tests. A set of internal baffle plates are arranged to block zero order radiation and stray diffraction grating orders from bright airglow lines.
The detector will use an open microchannel plate z-stack with a crossed delay line (XDL) detector (< 150 µm electronic resolution). We have flown these detectors on numerous missions including several rockets. The instrument electronics include a high voltage power supply attached to the cavity housing. A detector electronics module containing front end amplifiers and dual (X,Y) time delay to digital converters (TDCs) and a digital interface card are located in the XDL electronics stack. An existing electronics module is used that has been developed through the PKB-ALICE, LRO-ALICE, SSULI (DMSP), and JUNO-UVS flight programs.
Table 4: Summary of EUV instrument parameters
IVM (Ion Velocity Meter)
The IVM instrument is provided by UT (University of Texas), Dallas (PI: Rod Heelis). The IVM consists of two planar thermal ion sensors, a RPA (Retarding Potential Analyzer) and an IDM (Ion Drift Meter) that will provide measurements of the ion drift velocity in the spacecraft reference frame, the ion temperature and the total ion number density at the location of the spacecraft. Using a magnetic field model the drift velocity vector can be converted to a local electric field perpendicular to the magnetic field and expressed in a geographic and/or geomagnetic coordinate system. Above 250 km altitude the direction parallel to the Earth’s magnetic field is an electric equipotential for horizontal scales greater than 25 km. Thus the electric field at the spacecraft location can be translated to any other location along the magnetic field through the spacecraft. 32)
The IVM instrument is derived with little change from the instrument that is presently operating successfully as part of the CINDI (Coupled Ion-Neutral Dynamics Investigation) project on the C/NOFS (Communication/Navigation Outage Forecast System) satellite (launch on April 16, 2008). The IVM sensors and sensor electronics are constructed as a single package and are mounted to view approximately along the spacecraft velocity vector through a common aperture plane that is isolated from the spacecraft ground. Figure 19 shows the sensor configuration that is flying on the CINDI mission and and a schematic illustration of the ICON configuration.
The aperture plane serves as the reference ground for the sensors and ensures that all external fields are uniform and normal to the entrance aperture. A portion of the aperture plane, the SenPot reference surface, is isolated from the reference ground. The SenPot amplifier generates a potential with respect to spacecraft ground to maintain the IVM reference ground at the floating potential with respect to the plasma. The operation and data analysis techniques have extensive heritage from many previous missions starting with Atmosphere Explorer in the 1970’s and subsequently the DMSP (Defense Meteorological Satellite Program) and ROCSAT-1/FormoSat-1 before the CINDI mission in 2008.
The RPA (Retarding Potential Analyzer) is a planar sensor that presents a circular aperture to the incoming plasma stream that is intersected by a number of planar semi-transparent conducting grids and a large solid collector. Two schematic views of the RPA sensor are shown in Figure 20.
The sensor is mounted to view approximately along the satellite velocity through a large aperture plane that provides a uniform planar electrical ground reference potential. Grounded grids cover the entrance aperture to ensure that no internally applied potentials influence the ion beam trajectories prior to entering the sensor. Inside the sensor the ion beam traverses a series of semi-transparent grids before impacting the collector. The retarding grids are biased at potentials between 0 and 25 V and thus control the minimum energy that the ions must posses to reach the collector. A suppressor grid prior to the collector is biased at -12 V to reject ambient electrons and suppress photoelectrons ejected from the collector.
A current-voltage characteristic is obtained by measuring the ion current while the retarding voltage moves over a series of predetermined discrete levels. Retarding voltage sequences are chosen to optimize the sensor performance for the changing conditions expected in the ionosphere over a substantial part of a solar cycle. Figure 21 shows a simulated data curve obtained assuming a total ion concentration or 105 cm-3 with 20% H+ and 80% O+. This current-voltage characteristic has a well-known functional form that can be fitted to retrieve the ion drift component along the sensor look direction, termed , the ion temperature and the major constituent ions. The current at zero retarding voltage is used to derive the total ion number density.
The IDM (Ion Drift Meter) is a planar sensor that presents a square aperture to the incoming plasma stream that is intersected by a number of planar semi-transparent conducting grids and a solid segmented collector. The collector segments are arranged such that the cuts lie approximately along the satellite nadir and the orbit normal, which define the local vertical and horizontal directions perpendicular to the satellite motion. Two schematic views of the IDM sensor are shown in Figure 22.
The sensor is mounted to view approximately along the satellite velocity through a larger aperture plane that provides a uniform planar electrical ground reference potential. Two planar grids are placed prior to the entrance aperture to prevent the passage of light ion species. To ensure that no internally applied potentials influence the ambient external plasma the outermost grid, which is coplanar with the aperture plane, is always tied to reference ground, as is the aperture plane. This suppressor grid is biased at -12 V to prevent access of ambient thermal electrons to the collector and to suppress electrons liberated from the collector by solar euv radiation. Figure 23 shows a schematic cross-section and the O+ beam trajectory, in the plane of the satellite motion and the local vertical. Also shown is the electrical configuration indicating the current measured by logarithmic electrometers that provide the inputs to a linear difference amplifier.
The RPA utilizes an automatic ranging linear electrometer to measure the ion current for discrete values of the retarding potential that may specified by ground command and stepped at a 32 Hz rate. A nominal sweep sequence will comprise 16 points, recorded with 14-bit accuracy and executed in 0.5 seconds. The IDM utilizes two logarithmic electrometers that may be engaged to measure currents from collector segments that are alternately horizontal and vertical in the transverse direction. These signals provide the inputs to a linear difference amplifier that directly provides a signal proportional to the tangent of the ion arrival angle. The difference amplifier is sampled at 32 Hz with 14-bit accuracy, providing alternating horizontal and vertical arrival angles that provide a measure of the dynamic stability of the ionosphere and may be averaged to provide the transverse ion drift components that accompany the ram drift component derived by the RPA. The instrument performance capabilities are shown in Table 5.
The MOC (Mission Control Center) at UCB/SSL is tasked with operating the ICON mission. The MOC is currently operating THEMIS, ARTEMIS, RHESSI and NuSTAR, all NASA Explorer missions, recently operating the FAST Explorer as well.
For ICON, operations begin early with the integration of the spacecraft at Orbital ATK, where Berkeley operations engineers will work with the spacecraft engineers in all the initial tests and checkout of the bus. This continues with the arrival of the scientific payload, through integration of the full ICON observatory with the launch vehicle. This is exactly how Berkeley has worked with Orbital on NuSTAR, a NASA Small Explorer (PI F. Harrison – Caltech) which UCB is currently operating (quite successfully!).
With launch, ICON will be placed into a 27º inclination orbit and communications will immediately be established between ICON and the UCB MOC using TDRSS, the orbiting NASA communications network. Ground contacts are planned and used from then on to command and operate ICON. Ground contacts with ICON are performed mainly from the Berkeley Ground Station, with backup contacts out of Wallops and Santiago. 33)
Figure 24: Photos of the Berkeley ground station and the MOC (image credit: UCB/SSL)
Dwayne Brown, “NASA Selects Explorer Projects To Probe Earth's
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8) ”ICON Science Overview,” ICON Science Overview,” UCB/SSL, URL: http://dev.icon.ssl.berkeley.edu/Science/Science-Overview
9) ”ICON - Studying the Earth-Sun Connection in the Ionosphere,” Orbital ATK, Factsheet, URL:
11) “Orbital Selected By NASA For $50 Million Contract To Build Icon Space Weather Satellite,” Orbital Press Release, April 29, 2013, URL: https://www.orbital.com/NewsInfo/release.asp?prid=856
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15) Robert Sanders, ”Space weather satellite ICON on course for summer 2017 launch,” Berkeley News, April 7, 2016, URL: http://news.berkeley.edu/2016/04/07
17) Karen Fox, “ICON Cleared for Next Development Phase,” NASA, Nov. 12, 2014, URL: http://www.nasa.gov/content/goddard/icon-cleared-for-next-development-phase/
19) Thomas Immel, “ICON Passes System Requirements Review,” Jan. 14, 2014, URL: http://icon.ssl.berkeley.edu/icon-passes-system-requirements-review/
20) Robert Sanders, “UC Berkeley selected to build NASA’s next space weather satellite,” April 16, 2013, URL: http://newscenter.berkeley.edu/2013/04/16
21) ”ICON Launch Briefings and Events,” NASA, 17 October 2018, URL:
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25) ”The Instrument Control Package - ICP,” UCB/SSL, URL: http://dev.icon.ssl.berkeley.edu/Observatory/Instrument-Control-Package
27) “NRL's MIGHTI Slated for Launch on ICON Mission,” NRL, May 13, 2013, URL: http://www.nrl.navy.mil/media/news-releases/2013
28) C. R. Englert, J. M. Harlander, C. M. Brown, A. W. Stephan, J. J. Makela, K. D. Marr, T. J. Immel, “The Michelson Interferometer for Global High-resolution Thermospheric Imaging (MIGHTI): Wind and Temperature Observations from the Ionospheric Connection Explorer (ICON),” OSA Optics and Photonics Congress, Imaging and Applied Optics, Technical Digest, FW1D.3, ISBN 978-1-55752-975-6, 2013
29) “e2v image sensors chosen for NASA’s MIGHTI mission,” ev2, Feb. 5, 2014, URL: http://www.e2v.com/news/e2v-image-sensors-chosen-for-nasas-mighti-mission/
33) ”Mission Operations,” UCB/SSL, URL:
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 (firstname.lastname@example.org).