Minimize EO-1

EO-1 (Earth Observing-1)

EO-1 is a technology demonstration mission in NASA's NMP (New Millennium Program) Earth science program (a Landsat-7 follow-up) with the overall objectives to perform Landsat-like measurements and to explore new remote sensing technologies (instruments, spacecraft, ground segment) that advance and enhance capabilities [evaluation/validation of technologies and performance, intersatellite calibration (comparison of data), evaluation of lunar calibration, autonomous navigation/instrument operation, use of proven new technologies for other missions]. A number of baseline validation scenarios are defined (generation of test scenes, ground proofs in conjunction with major field campaigns, airborne underflights, tandem flights with Landsat-7, etc.) in support of the overall objectives. S/C design life = 1 year. The paradigm of “faster, better and cheaper” applies to all aspects of the mission. 1) 2) 3) 4)


Figure 1: Artist's view of the EO-1 spacecraft (image credit: NASA/GSFC, MIT/LL, Ref. 9)


The EO-1 spacecraft, managed and operated by GSFC, was developed, built and integrated by ATK Spacecraft Systems and Services (HQ Minneapolis, MN), formerly Swales Aerospace of Beltsville, MD as prime contractor (ATK acquired Swales in 2007). The S/C bus (of MIDEX heritage) features an Al structure, hexagonal in shape with 1.25 m diameter (across flats) and 0.73 m high (bus mass of 370 kg, payload mass of up to 110 kg). The S/C is three-axis stabilized for inertial and nadir pointing, an AST (Autonomous Star Tracker) provides three-axis attitude knowledge. Four 1 N thrusters of a hydrazine propulsion system provide a pointing accuracy of 0.03º, the jitter is < 5 arcseconds. The PPT, a secondary propulsion system, represents a new technology which may eventually replace the reaction wheels. 5)

S/C orbit and attitude adjustment is based on GPS receivers and ACS (Attitude Control System) sensors (demonstration of autonomous maneuver capabilities initiated by GPS). ACS consists of the following major components: reaction wheel assembly, magnetic torquer bars, a three-axis magnetometer, IRU (Inertial Reference Unit), an autonomous star tracker, a GPS receiver (Space Systems Loral GPS Tensor with 4 antennas), and coarse sun sensors. A versatile ACS software permits S/C slewing to celestial bodies (it provides instrument calibration capability of ALI sun and lunar calibrations); this provides also a cross-track pointing capability to observe adjacent tracks. The EO-1 ACS software includes the functions of attitude determination and closed-loop control modes for magnetic de-spin following separation from the Delta II launch vehicle, initial stabilization and sun acquisition, nadir pointed science data collection and downlink, thruster maneuvers for delta-V, and solar/lunar slew/scan maneuvers for instrument calibrations.


Figure 2: Line drawing of the EO-1 spacecraft (image credit: NASA/GSFC, Swales Aerospace)

The solar array consists a single wing, comprised of three panels. The array is canted at 30º, it utilizes a single-axis drive assembly to clock the array at an orbital rate, essentially keeping the array normal to the sun, minimizing cosine losses. A cable wrap mechanism is implemented between the rotating array and the S/C, which requires the array to “rewind” during eclipse periods, preparing for the next sunlit period in which it moves forward at orbital rate. Each panel is of size 1.26 m x 1.43 m. Two types of solar cells are used: Si cells of 15% efficiency and GaAs (cascade) cells of 22% efficiency (due to the later addition of the Hyperion instrument). A hydrazine propulsion system with 22.3 kg propellant mass is used for orbit maintenance and formation flying. An on-board solid-state recorder provides science data storage of up to 40 Gbit. S/C total mass =572 kg (dry mass = 370 kg, payload mass = 90 kg), power= 600 W EOL (300 W orbit average) and 50 Ah super NiCd battery (28±7 V DC power).

The OBC uses a Mongoose V (M5) processor with 12 MHz and 1.8 Gbit of TT&C data storage. A lightweight fiber optic data bus with ATM (Asynchronous Transfer Mode) protocol is used to transfer on-board data. 6) 7)

Note: The EO-1 spacecraft has two Mongoose M5 processors onboard. The first M5 is used for the EO-1 TT&C functions. The other M5 is part of the WARP (Wideband Advanced Recorder Processor) system, a large mass storage device (see below). Both M5's run the VxWorks operating system.


Figure 3: Schematic view of the avionics architecture (image credit: Swales Aerospace)

Orbit: Sun-synchronous circular polar orbit, altitude = 705 km, inclination = 98.7º, period = 99 minutes, descending nodal crossing time of 10:15 AM. The orbit supports a 16 day repeat cycle. - A coordinated tandem orbit within a minute of Landsat-7 is prescribed for reasons of data calibration and synergetic use of data.

Spacecraft launch mass

588 kg

Spacecraft bus dry mass

410 kg (including WARP and X-band phased array antenna (PAA)

Spacecraft bus size

1.4 m x 1.4 m x 2 m high

Spacecraft structure

Hexagonal; aluminum honeycomb

Spacecraft power

- 350 W (average), nominal voltage = 28 V,
- Batteries: super NiCd with 50 Ah capacity
- Array: 3 panel/Si with GaAs/articulating/5.25 m

Spacecraft stabilization

Zero momentum system, 3-axis stabilized,


- Actuators: reaction wheel assembly, magnetic torquer bars,
- Sensors: three-axis magnetometer, IRU, autonomous star tracker, coarse sun sensors
- GPS receiver (attitude measurements)
- Roll/pitch/yaw axis pointing accuracy < 0.02º, 3σ
- Pointing stability (jitter): 0.3 arcsec/s
- Slew rate: 15º / minute

Navigation (location) accuracy

60 m, each direction (3σ)

C&DH bus architecture

- Mongoose V processor, Rad Hard at 12 MHz, RISC architecture
- MIL-STD-1773 fiber optic data bus for serial command/telemetry communications at 1 Mbit/s

Hydrazine populsion system

- 1 tank, 4 thrusters
- propellant capacity: 23 kg
- Maximum ΔV = 85 m/s

RF communications

- Science data storage capability = 48 Gbit within WARP
- Science data downlink capacity = 105 Mbit/s
- Downlink formats/network: CCSDS / STDN, DSN, TDRS
- Downlink band: S-band (variable to 2 Mbit/s), X-band (105 Mbit/s)
- Uplink band: S-band (2 kbit/s)

Mission design life

1.5 years, (nominal life = 1 year)

Table 1: Overview of spacecraft parameters


Figure 4: Overview of EO-1 technology allocations on the EO-1 spacecraft (image credit: NASA)


Launch: A launch of EO-1 from VAFB on a Delta 7320-10 launch vehicle took place on Nov. 21, 2000 (along with SAC-C (CONAE, Argentina), and Munin (Sweden) as secondary payloads).

RF communications: The data system uses the MIL-STD-1773 fiber optic data bus for serial command/telemetry communications at 1 Mbit/s. The EO-1 antenna complement is comprised of two transmit/receive S-band omni antennas and a transmit-only electronically steerable XPAA (X-band Phased Array Antenna). The 64-element XPAA is nadir pointing. The XPAA is LHCP polarized with a 3 dB beam width which varies from 18º-30º depending on scan angle. It scans a 360º azimuth angle within a 65º half cone angle from boresight. Each of the 64 XPAA transmit elements contains its own solid-state power amplifier. One of the omni antennas is zenith facing, the other is nadir facing. Use of CCSDS protocols in all S/C communication links. The science data downlink is in X-band: frequency = 8225 MHz, data rate=105 Mbit/s, minimum EIRP = 22.0 dBW, encoding scheme= modulo-4 gray code differential encoding, modulation=QPSK. The TT&C links are in S-band, either direct or via TDRSS. The downlink data rate is selectable from 2 kbit/s to 2 Mbit/s (2, 32, 1000, or 2000 kbit/s), the uplink data rate is 2 kbit/s.

The ground network consists of the following stations: SGS (Spitzbergen Ground Station) at Svalbard, Norway; AGS (Alaska Ground Station) at Fairbanks, AK; WOTS (Wallops Orbital Tracking Station) at Wallops Island, VA; and MSG (McMurdo Ground Station) at Antarctica. SGS is the primary station. In addition, there is the TDRSS White Sands Station for links via TDRSS. EO-1 mission operations are provided at NASA/GSFC. 8)


Figure 5: Overview of the EO-1 ground system (image credit: NASA)



Sensor Complement (ALI, Hyperion, LAC)


Figure 6: Overview of EO-1 technologies (image credit: NASA/GSFC)


ALI (Advanced Land Imager):

ALI was designed/built and tested at MIT/LL as prime contractor within the New Millennium Program of NASA [focal plane system: Raytheon/Santa Barbara Remote Sensing (SBRS), optical system: Sensor Systems Group, Inc. (SSG), Wilmington, MA]. ALI is a hybrid multisensor system, a technology verification instrument, for the measurement of Earth surface reflectance. A prime objective is to provide continuity with data from earlier Landsat missions at reduced sensor/operating costs (introduction of a new measurement approach).

The ALI design features a WFOV (Wide Field of View) telescope and a highly integrated multispectral and panchromatic instrument. The telescope employs a reflective Cooke triplet mirror design with a 12.5 cm unobscured aperture diameter and a FOV (Field of View) of 15º (cross-track) x 1.26º (along-track) and is providing a ground swath of 185 km, nominally centered along the nadir ground-track. The design uses four mirrors; the primary is an off-axis asphere, the secondary is an ellipsoid, and the tertiary is a concave sphere; the fourth mirror is a flat folding mirror. This technology enables the use of large arrays of detectors at the focal plane for covering an entire swath of 185 km. The optical design features a flat focal plane and telecentric performance, which greatly simplifies the placement of the filter and detector array assemblies. The design uses lightweight silicon carbide (SiC) mirrors and an Invar truss structure with appropriate mounting and attachment fittings. The detector array assembly has 10 spectral bands in VNIR and SWIR. Note: SiC offers the advantage of a very high stiffness-to-density ratio and a very high conductivity to heat capacity ratio. These characteristics are superior to currently used materials for reflective optical systems. 9) 10) 11)

The focal plane assembly (FPA) for ALI is populated with four sensor chip assemblies (SCA) providing a continuous FOV of 3º x 1.625º for a fully populated focal plane, each SCA providing a cross-track coverage of 37 km. There are four SCAs per module. Each module covers a ground swath swath of about 37 km. A fully populated focal plane would have five populated modules, i.e., 20 SCAs in all. ALI has only one module as a low-cost technology demonstration.

Unlike the Landsat TM and ETM+ whiskbroom scanners, the ALI instrument operates in pushbroom fashion providing panchromatic and multispectral imagery. A single hybridized SCA is employed for PAN/MS (panchromatic multispectral) detection in the VNIR/SWIR spectral range (0.4 - 2.4 µm). This includes a PAN band, six VNIR bands (identical with Landsat ETM+), plus two additional VNIR bands (primed in Table 2), and three SWIR bands. The focal plane is only partially populated with SCAs resulting in only partial coverage by the various SCAs.

• Four MS SCAs are used side by side. Each SCA contains nine rows consisting of 320 MS detector cells in the cross-track direction. The dimension of each MS cell is 39.6 µm x 39.6 µm, providing a GSD (Ground Sample Distance) of 30 m

• The PAN detector cells are integrated directly into each MS SCA as a 960-element row, providing overlapping coverage to the MS bands, but at a 10 m GSD.


Figure 7: Schematic layout of ALI FPA and optics (image credit: MIT/LL)

The HgCdTe SWIR detectors provide high performance over the 0.9 to 2.5 µm wavelength region at temperatures which can be reached by passive or thermoelectric cooling. The nominal focal plane temperature is 220 K and is maintained by the use of a radiator. Application of detectors of different materials to a single readout integrated circuit (ROIC) enables a large number of arrays covering a broad spectral range to be placed close by together. This technology is extremely effective when combined with the WFOV optical design being used on ALI. The VNIR bands (MS/Pan) use silicon-diode detectors fabricated in the silicon substrate of a ROIC. The SWIR detectors are HgCdTe photodiodes that are Indium-bump bonded onto the ROIC that services the VNIR detectors. 12)

The nominal integration times are 4.05 ms for the MS detectors and 1.35 ms for the Pan. The frame rate can be adjusted in 312.5 ns increments to synchronize frame rate with ground scan velocity variations due to altitude and velocity variations during orbit. The focal plane electronics samples the output of each detector with a 12 bit converter.


Figure 8: Cross-sectional view showing the interior elements and optical path of the ALI (image credit: MIT/LL)


Figure 9: FPA of and layout of the detector chip assembly of ALI (image credit: MIT/LL)


Figure 10: Illustration of the main FPA (Focal Plane Assembly), image credit: MIT/LL

ALI (EO-1)

ETM+ (Landsat-7)

Band No.

Wavelength (µm)

GSD (m)

Band No.

Wavelength (µm)

GSD (m)


0.480 - 0.690 VNIR



0.50 - 0.90 VNIR

13 x 15


0.433 - 0.453 VNIR






0.450 - 0.515 VNIR



0.450 - 0.515 VNIR



0.525 - 0.605 VNIR



0.525 - 0.605 VNIR



0.630 - 0.690 VNIR



0.630 - 0.690 VNIR



0.775 - 0.805 VNIR



0.760 - 0.890 VNIR



0.845 - 0.890 VNIR






1.200 - 1.300 SWIR






1.550 - 1.750 SWIR



1.550 - 1.750 SWIR



2.080 - 2.350 SWIR



2.080 - 2.350 SWIR






10.40 - 12.50 TIR


Table 2: Spectral parameter comparison of ALI and ETM+ instruments


Figure 11: Illustration of the ALI instrument (image credit: MIT/LL)

ALI is the prime instrument of EO-1. It is mounted to and aligned to an interface plate, bolted to the S/C. The instrument-to-S/C alignment is within 20 arcseconds. A comparison of ALI/ETM+instruments (see also Tables 2, 3, and 6) reveals the following: ALI can gather higher-quality imagery than ETM+ but is smaller and less expensive (one-forth); its mass is one-forth, with one-half its size it requires only one-fifth the power of ETM+.

Calibration: ALI instrument calibration is provided by several means: solar and/or lunar calibration, internal reference lamp calibration, and ground truth calibration. 13) 14) 15) 16) 17)

• On-orbit solar instrument calibration of ALI is conducted in 2-week intervals. This involves pointing the instrument into the sun with the aperture cover closed. A motor-driven aperture selector in the aperture cover assembly moves an opaque slide over a row of small to increasingly larger slit openings and then reverses the slide motion to block all sunlight. Just prior to solar calibration, a space grade Spectralon® diffuser plate is swung over the secondary mirror by a motor-driven mechanism. The diffuser reflectively scatters the sunlight that would otherwise impinge on the secondary mirror. The scattered sunlight exposes the FPA to irradiance levels equivalent to Earth-reflected sunlight for albedos ranging from 0 to 100%.

• Lunar calibration involves observing the moon with the instrument and comparing the measured lunar irradiance with a predicted lunar irradiance for the observation period. Lunar calibration is used to test the optical MTF of the system.

• Another calibration method involves reflectance-based ground truth measurements of selected target sites (imaging of a stable, high-altitude, flat, diffuse ground target using the ALI while ground teams simultaneously measure the reflective properties of the target region and local atmospheric conditions). Several ground truth campaigns were conducted during the first year of the EO-1 mission.

• Internal calibration: The internal reference lamps are activated during two data collection events per day, when the ALI aperture cover is closed.


Figure 12: Spectral collimator used during spectral calibration of ALI (image credit: MIT/LL)


ALI (EO-1)

ETM+ (Landsat)

Instrument mass

106 kg

425 kg

Instrument power

100 W

545 W

Instrument size

0.2 m3

1.4 m3

Nr. of VNIR/SWIR bands



Detectors per band



Thermal bands



Data rate

300 Mbit/s

150 Mbit/s

Pan resolution

10 m

15 m

Relative SNR

4 x

1 x

Instrument observation technique


Whiskbroom scanner

Swath width

37 km

185 km

Table 3: EO-1/Landsat instrument parameter comparison


Figure 13: Photo of the ALI instrument (image credit: NASA)



The pushbroom Hyperion instrument is of HSI (Hyperspectral Imager) heritage on the Lewis S/C (launch Aug. 23, 1997, control of the S/C was lost Aug. 26, the S/C was lost and reentered the atmosphere Sept. 28, 1997) and was built by Northrop Grumman (formerly TRW) of Redondo Beach, CA. Hyperion is a pushbroom instrument with a grating imaging spectrometer (slit design); the design includes a telescope, two grating spectrometers with the supporting focal plane electronics, and the cooling system. 18) 19) 20)

The fore-optics design is based on the EOC (Electro-Optical Camera) instrument of the KOMPSAT mission (KARI of Korea). The telescope has an aperture diameter of 12 cm, it images the Earth onto a slit that defines the instantaneous FOV which is 0.624º wide (7.5 km swath) x 42.55 µrad in along-track. This slit image of the Earth is relayed at a magnification of 1.38:1 to two focal planes in the two grating imaging spectrometers. A dichroic filter in the system reflects the band from 400 to 1,000nm to one spectrometer (VNIR) and transmits the band from 900-2,500 nm to the other spectrometer (SWIR). The SNR is improved by the two separate spectrometers.


Figure 14: Illustration of the Hyperion instrument (image credit: NASA)


Figure 15: Schematic view of the Hyperion optical system (image credit: NASA)


Figure 16: The Hyperion imaging spectrometer (image credit: NASA)

A focal plane array provides separate VNIR and SWIR detectors (HSI heritage). The SWIR detectors are cooled by a cryocooler. A dichroic filter in the system reflects the band from 400 to 1,000 nm to one spectrometer and transmits the band from 900 to 2,500 nm to the other spectrometer. The SWIR overlap with the VNIR from 900 to 1000 nm permits cross calibration between the two spectrometers. Both spectrometers use a JPL convex grating design in a 3 reflector Offner configuration. The VNIR spectrometer uses a 60 (spectral) x 250 (spatial) pixel detector array, which provides a 10 nm spectral bandwidth over a range of 400-1000 nm. The SWIR spectrometer has HgCdTe detectors in an array of 160 (spectral) x 250 (spatial) channels, cooled to 120 K. The SWIR spectral bandwidth is 10 nm. The telescope is a three-mirror astigmatic design, the high resolution hyperspectral imager is capable of resolving 220 spectral bands (from 0.4 to 2.5 µm) with a 30 meter resolution. All of the mirrors in the system are constructed from coated aluminum; the structure holding the optical elements is also constructed from aluminum so that the mirrors and housing all expand and contract at the same rates.

Hyperion calibration: A common calibration system is provided for both the VNIR and SWIR spectrometers. Dual calibration lamps produce reference signals to monitor detector performance following image acquisition. Calibrations are also performed with the sun, ground target sites, and the moon. The solar calibration utilizes a diffuse reflector on the backside of the optical cover to provide uniform illumination across the focal plane arrays. Direct viewing lunar calibration is accomplished by scanning the instrument across the lunar surface. Solar and in-flight calibration data is used as the primary source for monitoring radiometric stability, with ground site and lunar calibration as secondary alternatives.

The Hyperion on-orbit calibration strategies fell into four general categories: solar; lunar, lamp-based, and opportunistic. Currently, solar calibrations are performed every two weeks. As EO-1 passes beyond the North Pole in its orbit, the spacecraft is maneuvered to point the solar baffle aperture at the sun. The internal lamps are exercised as part of this calibration procedure. 21)


Figure 17: Block diagram of the Hyperion instrument (image credit: NASA)

The instrument typically images surface scenes of 7.5 km (swath) by about 100 km in length, providing detailed spectral mapping across all 220 channels with high radiometric accuracy. The data is typically processed into cubes (19.8 km long by 7.5 km wide) to facilitate data handling in current desktop computers. Each image cube consists of 75 MByte of data. A typical acquisition consists of multiple image cubes.

Spectral coverage

400 - 2500 nm; 400 - 1000 nm (VNIR); 900 - 2500 nm (SWIR)


0.624º (i.e., 7.5 km swath width from a 705 km orbital altitude)

IFOV (spatial resolution)

42.55 µrad (30 m )

Imaging aperture

12.5 cm diameter


60 x 250 silicon CCD array in VNIR; 160 x 250 HgCdTe array

No of spactral bands


Spectral resolution

10 nm in VNIR and SWIR

Data quantization

12 bit

Frame rate

225 Hz

Cross-track spectral error

< 1.5 nm (VNIR), < 2.5 nm (SWIR) (TBR)

Spatial co-registration of spectral bands

< 20% of pixel (TBR)

Absolute radiometric accuracy

<6% (1σ)

Instrument mass, power

49 kg, 78 W (average)

Instrument size

39 cm x 75 cm x 66 cm

Table 4: Hyperion instrument parameters


Figure 18: Optical system of Hyperion (image credit: NASA)


Cover position

Data collect





Default mode for active state

Dark calibration


100 frames (min)

Performed as close as possible to imaging, before and after

Lamp calibration


100 frames (min)

Performed after second dark calibration: two radiance levels

Solar calibration

Open 37º

1 second (min)
nominal 1 cube

Performed over North Pole only to keep cover out of ALI keepout zone; yaw maneuvers required

Lunar calibration

Fully open (135º)

1 second (min)
nominal 1 cube

Performed on dark side of earth; off-track spacecraft pointing required

Ground calibration

Fully open (135º)

1 second (min)
nominal 1 cube

Ground target selected


Fully open (135º)

1 second (min)
nominal 9 cubes

Nominal data collect is equivalent to Landsat scene, and takes 27 seconds

Table 5: Nominal data modes of Hyperion


Figure 19: Hyperion data handling scheme (image credit: NASA)


LAC (LEISA Atmospheric Corrector):

The objective is to provide atmospheric correction data (water vapor) for ALI and ETM+ on Landsat-7. LAC is a high spectral, moderate spatial resolution hyperspectral imager based on wedge filter technology. LAC is of WIS and LEISA heritage. LEISA was originally developed for the Lewis satellite mission (launch Aug. 23. 1997) under the Advanced Technology Insertion Program (however, on Aug. 26, three days after launch, an inflight anomaly led to the loss of attitude control and a discharged battery - contact was lost when Lewis entered a spin that disrupted the satellite's power?generating capability. LEISA was also flown as an airborne instrument in the summers of 1997, 1998 and 1999 as part of the instrument complement in an agricultural sensor program.

The spectral coverage of LAC is from 0.89 - 1.6 µm; in addition there is a separate channel at 1.380 µm to detect cirrus clouds. Spatial resolution = 250 m. LAC is coaligned with the ALI WFT assembly to cover the full swath of 185 km.

LAC is comprised of two modules, the optics module and the electronics module. The optics module contains the lenses, focal planes and electronics necessary to operate the arrays and to transfer the digitized pixel data to the electronics module. It is mounted to the nadir deck of the spacecraft and bore-sighted with ALI. The electronics module contains the command and data interface to the spacecraft, the array timing and bias circuitry, the thermal electric cooler (TEC) control circuitry and the instrument power supply. It is mounted in a bay below the nadir deck.

LAC has a total mass of 10.5 kg, 4.4 kg for the electronics module, 3.9 kg for the optics module and 2.2 kg for the cable connecting them. It uses a maximum power of 48 W on start-up, which decreases to about 35 W for a nominal TEC temperature setting of 275 K, once the temperature is stabilized. At the nominal fame rate of 28 Hz, which slightly oversamples the spatial dimension in the along-track dimension, the data rate is 95 Mbit/s (12 bit A/D converters are used). A frame rate of 56 Hz allows double sampling in the along-track spatial dimension at the expense of reduced single-frame signal-to-noise ratios (SNR).


Figure 20: Photo of the optics module of the LAC instrument (image credit: NASA)

Legend of Figure 20: The optics module is shown with lenses facing upward and solar calibration tubes facing forward. The red object to the right is a cover for the optical alignment cube used to align the LAC with the ALI. The fitting in front is an N2 purge coupling.

The instrument uses three 256 x 256 pixel InGaAs infrared detector focal plane assemblies in a single module. Each array is placed behind a lens covering a FOV of 5º to obtain a swath width of 185 km (15º). A wedged dielectric film etalon filter [a linear variable etalon (LVE)] is placed in very close proximity to a 2-D detector array. This produces a 2-D spatial image that varies in wavelength along one dimension. The filter has a nearly linear dependence of wavenumber on position. It has a 0.45 cm section which covers the 1.2 to 1.6 mm spectral region at a resolution of about 35 cm-1, and a 0.55 cm section covering the 0.9 to 1.2 mm spectral range at a resolution of about 50 cm-1. The sections are bonded together to form a single filter assembly. This filter represents an advance in dielectric thin film technology. Reflective ¼-wave stacked layers placed on both sides of one, or more, ½-wave etalon cavity(s) provide the spectral resolution. Order-sorting of the etalon is accomplished with lower resolution filter layers. In operation, the 2-D spatial image is formed by a small, wide field of view (WFOV) lens. Unlike the grating spectrometer that captures the spectra at a point “instantaneously,” the spectrum for the LAC is obtained as the orbital motion of the spacecraft scans the image across the focal plane in wavelength, thereby creating a 3-D spectral map. The spatial resolution of a pixel is 360 mrad x 360 mrad, or about 250 m x 250 m at nadir. 22) 23)


Figure 21: Schematic view of the optics module (image credit: NASA)


Multispectral Instruments

Hyperspectral Instruments


ALI (EO-1)

Hyperion (EO-1)

LAC (EO-1)

Spectral range

0.4 - 2.4 µm
(excluding TIR)

0.4 - 2.4 µm

0.4 - 2.5 µm

0.89 - 1.60 µm

Spectral resolution

30 m

30 m

30 m

250 m

Swath width

185 km

37 km

7.5 km

185 km

Spectral resolution



10 nm

2-6 nm

Spectral coverage





PAN band resolution

13 m x 15 m

10 m



Nr of bands





Table 6: Comparison of some instrument parameters on EO-1 and Landsat-7


Figure 22: ALI and Hyperion scene dimensions (standard and extended)) and corresponding Landsat 7 WRS path and row of ETM+ (image credit: USGS)


Figure 23: Formation flight geometries of EO-1 with Landsat-7 as part of the morning constellation train (image credit: NASA)



Demonstration of new technologies (XPAA, CCR, LFSA, WARP, PPT, FODB, AutoCon)

One important purpose of the EO-1 mission is to, wherever possible, maximize the transfer and infusion of EO-1 validated technologies into other applications and missions. Described below are the technologies that have been transferred and infused or have strong potential for such. 24)


XPAA (X-band Phased Array Antenna):

XPAA was developed and built by the Boeing Co. XPAA is a communication experiment with the objective to demonstrate link-pointing capability with the use of a body-fixed low-mass and low-cost phased array antenna. The antenna is mounted on the Earth-facing side of EO-1 to allow communications with ground stations (downlink of high-rate data from the EO-1 solid-state recorder). 25) 26)

XPAA is composed of a flat grid of 64 radiating elements whose transmitted signals are combined spatially to produce the desired antenna directivity. The phases of each of the radiating elements are varied by computer to point the beam in the desired direction. It has an Effective Isotropic Radiated Power (EIRP) of approximately 160 W and transmits data at 105 Mbit/s. An inherent advantage of the body-fixed design is to permit simultaneous capture and transmission of data, avoiding perturbations to instrument measurements. Some performance parameters are:

• Frequency: 8225 MHz

• Bandwidth: 400 MHz

• Scan Coverage: 60º half-angle cone

• Radiating Elements: 64

• Polarization: LHCP

• Command Interface / Controller: 1773 / RSN


Figure 24: EO-1 X-band system configuration diagram (image credit: NASA)


Figure 25: Photo of the XPAA assembly (image credit: NASA)

XPAA has an integral controller and power conditioner, communicates with the spacecraft over a MIL-STD-1773 fiber-optic data bus, and is fully space qualified. The antenna aperture consists of an 8 x 8 array of modules, each comprising a dielectrically loaded, circular waveguide, two orthogonal antenna feeds, a phase shifter, and a dual power amplifier. The 64 modules are mounted in a printed wiring board, which distributes radio frequency (RF) excitation, logic control signals, and power to each module. The array and remote service node (RSN) are located in a single enclosure (size: 30.5 cm x 33 cm x 7.4 cm) and a total mass of 5.5 kg (Figure 25).

The validation plan for the XPAA called for collecting data to meet the following objectives:

• Validate the communications link error performance

• Validate the antenna pattern scan performance of the phased array

• Validate the performance and reliability of the antenna's electronics and software in the space environment.

The on-orbit performance of the XPAA was validated, first and foremost, by the conduct of successful science downlinks. Throughout the first year of on-orbit operation, the XPAA has been operated at a rate more than five times the original requirement, without difficulty. All tests show a consistent performance throughout the life cycle of the antenna. 27)


CCR (Carbon Carbon Radiator):

CCR is a material science experiment with the objective to demonstrate various uses of carbon-carbon. The composite material is used in the primary structure of EO-1, serving as both an advanced thermal radiator and a load bearing structure. The advantages of carbon-carbon are: high thermal conductivity (including through thickness) and good strength and weight characteristics. The use of CCR technology may simplify thermal radiator design in some applications (e.g., no need for actively cooled radiators) and support increased science payload mass ratios for future missions. 28)

The Carbon-Carbon Radiator (CCR) shown in Figure 26 is a sandwich composite panel with facesheets made of carbon fibers in a carbon matrix. The EO-1 flight panel is coated with an epoxy encapsulant (Figure 26 left) to prevent particle contamination of sensitive instruments on board EO-1, and provide additional strength to the panel. The external surface (Figure 26 right) of the CCR panel is coated with silver Teflon as required by the EO-1 spacecraft thermal design.


Figure 26: Illustration of the CCR panels (image credit: NASA)

Carbon-Carbon (C-C) is a special class of composite materials in which both the reinforcing fibers and matrix materials are made of pure carbon. The use of high conductivity fibers in C-C fabrication yields composite materials that have high stiffness and high thermal conductivity. The primary thermal function of the EO-1 CCR is to radiate the 27.8 W generated by the EO-1 Power Supply Electronics (PSE) and the 16.3 W (peak power) generated by the LAC electronics boxes. The panel is also a structural member and must support the combined weight of the PSE (23 kg) and the LAC (5 kg) boxes and the dynamic and static loads during EO-1 integration, launch and orbit induced stresses.

The CCR is a 73 cm x 73 cm composite panel with two 0.56 mm thick C-C facesheets bonded to a 2.5 cm 5056 aluminum honeycomb core weighing ~ 3 kg.

The Carbon-Carbon Radiator panel met the mission requirements and its on-orbit performance has been flawless to date. The thermal model results correlate very well with the EO-1 flight data. The on-orbit thermal conductivity correlation for the CCR panel resulted in thermal conductivities close to the reported value of 230 W/m K.


LFSA (Lightweight Flexible Solar Array):

Photovoltaic (PV) solar arrays are the primary sources of electrical power for all Earth-orbiting satellites. The LFSA technology could, for some missions, provide higher power-to-mass ratios (specific energy) than conventional solar arrays, thus allowing a higher science payload mass fraction. The LFSA technology is a lightweight photovoltaic solar array system. 29)

The objective of this photovoltaic system is to demonstrate significant improvements in the power to weight ratios of solar arrays. The LFSA design features:

1) Solar cells of copper indium diselinide (CuInSe2) also referred to as the CIS techniique for an integrated thin?film solar cell technology.

2) SMA (Shape Memory Alloys) for the hinge and deployment systems.

The shockless deployment technique could improve the spacecraft dynamics during deployment. It is also much safer than conventional solar array systems using conventional pyrotechnics for deployment. The goal of the LFSA initiative is to achieve > 100 W/kg power efficiency ratios (conventional solar arrays of the late 1990s provide <40 W/kg).

CIS solar cells: The CIS thin film solar cell technology is vapor-deposited on a flexible substrate which is substantially lighter than cells bonded to a rigid panel. The LFSA solar cell modules are 10 cm x 10 cm and each consists of 15 monolithically-interconnected cells in series. The AM0 (Air-Mass-Zero) module efficiency achieved for this size was approximately 2%. Higher efficiencies have been achieved on smaller areas.


Figure 27: View of the LFSA flight unit demonstrator (image credit: NASA)


Figure 28: SMA hinge in stowed configuration (image credit: NASA)


Figure 29: SMA hinge in deployed configuration (image credit: NASA)

The SMA hinge technique offers substantial mass savings over conventional hinges, deployment systems, and solar array drives. Therefore, a combination of these technologies could provide significant improvement in the power-to-mass ratios. The shockless deployment could improve the spacecraft dynamics during deployment, and also is much safer to handle,integrate and test that conventional pyros. It is also electrically resettable so that the same device flies that is tested. The SMA deployment/hinge devices are significantly cheaper, simpler and therfore more reliable than current technology.

The dual flexure concept was developed for integration on the EO-1/LFSA flight experiment as shown in Figures 28 and 29. In this concept, the SMA strips are heat treated in the deployed (“hot”) configuration and joined at the ends by metallic structural fittings. In the martensitic (“cold”) state, the hinge is manually buckled and folded into the stowed configuration. Application of heat via internally bonded, flexible nichrome heaters transforms the SMA into the austenitic (“hot”) state and causes the hinge deploy. Once deployed, power is turned off and the SMA is allowed to cool back to the low temperature martensitic phase. Although the martensite phase is “softer” than the high temperature austenite phase, the very efficient section geometry in the deployed configuration allows the martensitic SMA hinge to support the lightweight solar array sections.

On-orbit validation: The LFSA was deployed shortly after launch. The indicator switches and the panel temperature profiles indicated that the deployment was nominal.


WARP (Wideband Advanced Recorder Processor):

WARP is a high-rate solid-state recorder on the EO-1 spacecraft. The objective is to demonstrate a number of high density electronic board advanced packaging techniques. WARP utilizes advanced integrated circuit packaging (3-D stacked memory devices) and “chip on board” bonding techniques to obtain very high density memory storage per board (24 Gbit/memory card). It also includes a Mongoose V processor which can perform on-orbit data collection, compression and processing of land image scenes. WARP consists of a high-rate (up to 840Mbit/s capability), high-density (48 Gbit storage), low mass (< 20 kg) solid-state recorder/processor with X-band modulation capability. It is in fact the highest rate solid-state recorder NASA has flown so far. 30) 31) 32) 33)

Data storage

48 Gbit

Data recording rate

> 1 Gbit/s burst mode, 900 Mbit/s continuous

Data playback rate

105 Mbit/s in X-band (with built-in RF modulator)
2 Mbit/s in S-band

Data processing

Post-record data processing capability

Instrument mass, size

18 kg, 25 cm x 39 cm x 37 cm


38 W orbital average, 87 W peak


15 - 40ºC minimum operating range

Mission life

1 year minimum, 1999 launch


15 krad minimum total dose, LET 35 MeV

Table 7: WARP instrument specifications


Figure 30: EO-1 flight data system architecture (image credit: NASA)


Figure 31: Hardware architecture of WARP (image credit: NASA)


Figure 32: Image of WARP assembly in the spacecraft (image credit: NASA)


PPT (Pulsed Plasma Thruster):

PPT was designed and built at General Dynamics (formerly Primex Aerospace) for NASA/GRC. The system is also referred to as PRS-101 (Plasma Rocket System-101), its model number. The objective is to demonstrate on-orbit electromagnetic propulsion technology and to provide a S/C precision-pointing capability. The EO-1 primary propulsion is provided by the hydrazine system for orbit maintenance and formation flying. The PPT is being used during limited periods of the mission to demonstrate the capability of a PPT system by performing the function of a momentum wheel. When in operation, the pitch axis momentum wheel is being idled and the PPT takes over control of the pitch axis. 34) 35) 36) 37)

The PPT consists of a coiled spring to feed the Teflon propellant, an igniter plug to initiate a small trigger discharge and an energy storage capacitor and electrodes (also: on-board electronics to convert the low voltage from the S/C to appropriate charging voltage for the capacitor, to provide energy to fire spark plugs, and to isolate PPT discharge from the S/C power bus). Plasma is created by the ablation of the Teflon propellant from discharge of the storage capacitor across the electrodes. The plasma is accelerated by the Lorentz force in the induced magnetic field to generate thrust. Impulse levels of 10-1000 µNs are achieved; the specific impulse is 880-1170 m/s; the average power consumption of PPT is between 1-100 W. PPT is used to maintain S/C attitude (fine pitch attitude control pointing requirements while meeting stringent electromagnetic and contamination constraints for the mission).


Rectangular, parallel plate, breech fed propellant

Trust axis

Two parallel and opposing

Impulse bit

90-860 µNs, continuously throttleable

Specific impulse

650-1400 s (estimated)

Total impulse

460 Ns (estimated)

Pulse frequency

Single pulse or 1 Hz

Stored energy

8.5 - 56 J, throttleable

Input power

70 W at 28 V, maximum energy, 1 pps

Overall efficiency

8% at maximum energy (estimated)

PPU efficiency

81% (estimated)


40.1 µF, 2 kV

Total thruster mass

4.95 kg

Fuel mass, fuel

0.07 kg per side, rectangular cross-section fluorcarbon polymer bars

Table 8: Characteristics of the PPT instrument


Figure 33: Illustration of the PPT instrument (image credit: NASA)


Figure 34: Schematic layout of the PPT instrument (image credit: NASA)

A series of fine pitch pointing maneuvers were performed after the end of the primary imaging mission. Flight operations of the EO-1 PPT began on January 4, 2002. As of June 15, 2002, a total of 26.9 hours of operation and almost 97,000 pulses have been logged, including several image acquisitions and continuous control of the pitch attitude of the spacecraft for over 9 hours for 5.5 orbits. This is the first flight demonstration of PPT throttling. Spacecraft pitch attitude was controlled to well within the 30 arcsec requirement during image acquisition and was generally within 10 arcsec. In addition, sensitive tests with images of the dark Earth have detected no evidence of electromagnetic interference from the discharge or light pollution from the plume even though the ALI instrument was known to have very sensitive electronic components. 38) 39)


FODB (Fiber Optic Data Bus):

The objective is to demonstrate high aggregate bandwidth data transfer capabilities onboard EO-1 with low mass and low power requirements. FODB was developed as a standard high-speed (Gbit/s) data interface for future spacecraft. It complies with the IEEE P1393 SFODB (Spaceborne Fiber Optic Data Bus) standard. The bus concept is based on a ring topology of 2 to 128 nodes which provides flexibility to meet differing payload requirements. The master node is a Controller Fiber Bus Interface Unit (CFBIU) and up to 127 slaves are implemented with a Fiber Bus Interface Unit (FBIU). The interface between nodes is implemented in fiber optics, it reduces weight and provides for an EMI/EMC problem-free system. 40)

The FODB high data transfer rate and software configurable ATM (Asynchronous Transfer Mode) based protocol provides users extraordinary flexibility with which to design their data handling architectures. The COTS-based 12-channel parallel fiber optic transmitter and receiver pair were developed specifically for parallel FODB through DoD and NASA funding [NASA SBIR (Small Business Innovation Research) contract by Optical Networks, Inc (ONI)].

The FODB on EO-1 is implemented with four nodes. In this configuration, the WARP data system contains the master controller node (CFBIU) and one slave node (FBIU). An external instrument terminal box, mounted on the spacecraft, contains two slave nodes, glue logic, power, and ALI instrument interface connectors. The FODB bus employs an ATM (Asynchronous Transfer Mode) with minimum overhead which simplifies the node interface design. Transfer rates between 200 Mbit/s and 1 Gbit/s are supported by FODB.


Figure 35: The parallel FODB ribbon cable (image credit: NASA)


AutoCon (Autonomous Control for enhanced formation flying)

NASA/GSFC teamed with a.i.-solutions, Inc. to fly AutoConTM onboard EO-1. The objective is to demonstrate an on-board software package capable of autonomously planning, executing, and calibrating routine spacecraft maneuvers to maintain satellites in their respective constellations and formations. The long-term goal of this technology is to enable many small and inexpensive spacecraft to fly in formation and gather concurrent science data in a “virtual satellite”. This ”virtual satellite or virtual platform” concept lowers total mission risk, increases science data collection and adds considerable flexibility to future Earth and space science missions. 41) 42) 43)

The AutoConTM technology automated EO-1's maneuver planning and formation control. AutoConTM performs orbit maintenance of single spacecraft or constellations and formations. It can be applied to a variety of orbits ranging from LEO to non-Keplerian trajectories such as libration orbits. EO-1 has as a principal mission requirement to successfully complete paired scene observations with Landsat-7 in order to validate the technologically advanced imagers on EO-1. A paired scene process requires EO-1 to fly over the current groundtrack of Landsat-7 a repeating groundtrack mission, within ±3 km. Also, in order to maintain a safety criterion, the nominal along-track separation is one-minute (450 km), ± 6 s (42.5 km). The six-second tolerance is derived from a ± 3 km groundtrack requirement. maintaining this 3-D separation requirement is referred to as formation flying.

The AutoCon features include an innovative use of fuzzy logic decision making capabilities and natural language to resolve multiple conflicting constraints; scripting environment enables algorithm updates without software changes; flight wrapper interfaces directly with the command and data handling subsystem for input and output; multiple operating modes allow execution control; generic closed-loop formation flying control algorithms applicable to many missions; modular architecture design is flexible enough to control execution of multiple and varying algorithms from several partners.

The EO-1 spacecraft is using a Tensor GPS receiver of Space Systems/Loral (SS/L) for active on-board navigation and attitude control. The flight interface connects directly with the C&DH (Command and Data Handling) system to retrieve all required data, including GPS position information, and to create command loads for computed burn times and durations. Only the objects and methods needed to support EO-1 formation flying are incorporated in the AutoConTM system conserving onboard resources. The AutoConTM flight control system is compatible with the various onboard navigation systems ((i.e. GPS, uploaded ground-based ephemeris, etc). The AutoConTM system onboard EO-1, NASA's first autonomous formation flying mission, demonstrated that autonomous control and formation flying technologies can be implemented.


Figure 36: Functional elements of AutoCon (image credit: NASA) 44)

There are two principal components, AutoCon-F (flight) and AutoCon-G (ground), which are used for in-flight operation and ground simulation respectively. This sharing of parts enables use of the same code on the ground and in flight, which reduces complexity and increases reliability.

Formation-keeping control of the EO-1 satellite was successfully handed over to the AutoCon-F software on May 17, 2002, demonstrating for the first time the autonomous formation-flying technology that will enable a host of future multi-spacecraft missions. EO-1 was able to stay within one second of its specified position using onboard algorithms to plan and execute maneuvers, allowing the EO-1 hyperspectral imager to analyze the same image as the Landsat-7 instruments without atmospheric distortion. Although initially planned as a technology experiment, AutoCon-F has performed so reliably that it now functions as the operational software for regularly planned formation control maneuvers.



EO-1 mission overview and operational phases:

The EO-1 baseline mission (funded by NASA) lasted from launch (Nov. 21, 2000) to the end of 2001, covering mainly instrument performance and instrument data validation and analysis. EO-1 has provided a testbed for refining specifications and expectations in the Landsat Data Continuity Mission (LDCM). As an example, over 5,000 data collection events have been successfully completed, against original success criteria of 1,000 data collection events.

The EO-1 baseline mission was complemented by several campaigns. Examples are:

• Extensive fieldwork, supported by the HyMap airborne spectral imager, was conducted in Australia from early launch throughout the first 120 days in orbit.

• Another extensive field campaign was conducted in Argentina from early January through the third week of February 2001 using the AVIRIS (Airborne Visible/Infrared Imaging Spectrometer) instrument of NASA/JPL.

ALI has undergone extensive on-orbit testing, characterization and calibration. The results indicated superior performance in resolution, image quality, SNR, dynamic range, radiometric accuracy, and repeatability.

All NMP mission objectives were accomplished within 13 months of launch as planned. EO-1 has performed beyond expectations and has demonstrated many new applications for both multispectral and hyperspectral remote sensing of the Earth. In particular, continuous improvements in EO-1 operations have resulted in a six-fold reduction in the cost of EO-1 imagery while achieving a four-fold increase in the data collection rate. These operational improvements are widely applicable to future missions.

EO-1 data became public in December of 2001 at the onset of the EO-1 extended mission (funded by USGS). The NASA/USGS arrangement calls for USGS/EDC to take over level 0 and ALI level 1 data processing at the Sioux Falls SD site. The extended mission opened opportunities for the public to commission EO-1 data collects. Data from both the base mission and extended mission are publicly available from the EROS Data Center (EDC). 45) 46)


Mission status:

• The EO-1 mission is operational in 2014, in its 14th year on orbit. The EO-1 mission was extended for the next two years as baselined at the very low requested support level and then closes out in FY16-17 (Ref. 49).

• December 2013: Volcanic activity along the western edge of the Pacific “Ring of Fire” gave rise to a tiny island in late November 2013 (Figure 37). Located in the Ogasawara Islands, part of the Volcano Islands arc, the new islet sits about 1,000 km south of Tokyo in waters considered part of Japanese territory. 47)


Figure 37: Natural color image of the new island of Niijima acquired by the ALI instrument on Dec. 8, 2013 (image credit: NASA)

Legend to Figure 37: The new island, named Niijima, rose up out of the sea during a volcanic eruption first reported on November 20, 2013. Niijima sits about 500 m from Nishino-shima, another volcanic island that last erupted and expanded in 1973–74. The two islands are located at approximately 27°14’ North latitude and 140°52’ East longitude, about 130 km from the nearest inhabited island.

In the first few days after the eruption, scientists speculated that Niijima might not last. New islets like those recently formed off the coast of Pakistan (Figure 38) and in the Red Sea can naturally sink back below the water line as they are eroded by wave action that carries away loose sediment, mud, and tephra (volcanic rock fragments). Some subsidence can also occur from the simple weight of gravity and the cooling of the hot rock.

But according to news reports, Niijima is still erupting and growing. Scientists from the Japan Meteorological Agency (JMA) think the island is large enough to survive for at least several years, if not permanently. By early December, Niijima had grown to 56,000 m2 (5.6 hectar), about three times its initial size. It stands 20 to 25 m above the sea level (Ref. 47).

• On Sept. 24, 2013, a major strike-slip earthquake rattled western Pakistan, killing at least 350 people and leaving more than 100,000 homeless. The 7.7 magnitude quake struck the Baluchistan province of northwestern Pakistan. Amidst the destruction, a new island was created offshore in the Paddi Zirr (West Bay) near Gwadar, Pakistan.

On September 26, 2013, ALI (Advanced Land Imager) on NASA’s EO-1 (Earth Observing-1) satellite captured the image of that new island, which sits roughly 1 km offshore. Likely a “mud volcano,” the island rose from the seafloor near Gwadar on September 24, shortly after the earthquake struck about 380 km inland. 48)


Figure 38: New island off the coast of Pakistan acquired by the ALI instrument on Sept. 26, 2013 (image credit: NASA)

Legend to Figure 38: In the satellite image, lighter shades of green and tan in the water reveal shallow seafloor or suspended sediment. The water depth around the new island is roughly 15 - 20 m, according to marine geologist Asif Inam of Pakistan’s National Institute of Oceanography. The image from ALI is also clear enough to show the parallel ripples of waves marching toward the shore.

• June 2013: The 2013 Senior Review evaluated 13 NASA satellite missions in extended operations: ACRIMSAT, Aqua, Aura, CALIPSO, CloudSat, EO-1, GRACE, Jason-1, OSTM, QuikSCAT, SORCE, Terra, and TRMM. The Senior Review was tasked with reviewing proposals submitted by each mission team for extended operations and funding for FY14-FY15, and FY16-FY17. Since CloudSat, GRACE, QuikSCAT and SORCE have shown evidence of aging issues, they received baseline funding for extension through 2015. 49)

- The 12 ½ year old EO-1 mission continues to make numerous valuable contributions to the Earth Science community. It serves as a model for advanced technology capabilities, including spacecraft agility, on-board intelligent processing, reliable support technologies, and unique passive optical imagery. The mission also delivers and tests new technologies and strategies for satellite acquisition, algorithms for terrestrial environmental monitoring, calibration and validation, data synergy, and continuing technology advancements for data volume throughput, autonomous operations, and on-board processing. The two instruments onboard EO-1, the Advanced Land Imager (ALI) and Hyperion, are high spatial resolution sensors capable of imaging any spot on earth up to 5 times every 16 days (plus 5 nighttime images over the same period).

- The EO-1 mission is central to NASA’s strategic Earth Science plan: to advance Earth system science related to climate and environmental change, and to characterize, understand, and predict how the earth is changing, with consequences for life on Earth. Though originally designed as a technology demonstration project (ALI to inform Landsat 8 and Hyperion as the first grating-based, hyperspectral, civilian sensor in orbit), the mission continues to serve that purpose but has also made significant progress as a contributor to science and applied science investigations as well as national and international disaster monitoring efforts. Despite the small team size (<10 part time) in the Mission Science Office and Missions Operations, the EO-1 mission supports an impressive array of activities.

- Over the past 2 years, significant progress has been made on improving communication to the broader community of how data should be tasked and acquired, and improving the Level 2 products, as well as on other fronts. These improvements have strengthened the relevance of EO-1 to NASA’s science goals. In particular, the EO-1 Mission Science Office (MSO), in partnership with the USGS EROS has made data processing improvements and continues to advocate for science-quality products. The panel also notes significant improvements to data tasking and acquisition and the mission is highly commended for these changes. The EO-1 website provides mission information, a reference list, user support tools, and links to data. In addition, through its web-based task management system (GeoBPMS), the user community can submit task requests, which could likely serve as a model for future sampling missions.

- The EO-1 mission was extended for the next two years as baselined at the very low requested support level and then closes out in FY16-17 (Ref. 49).

• The EO-1 mission is operational in 2013, in its 13th year on orbit. 50)


Figure 39: ALI natural color image of Plosky Tolbachik, Ostry Tolbachik, and Tolbachinksy Dol volcanos on the Kamchatka Peninsula,Russia, acquired on April 5, 2013 (image credit: NASA) 51)

Legend to Figure 39: The Tolbachik Volcano is not a single peak, but a complex of volcanic features superimposed on one another. The varied shapes result from differences in the chemistry, gas content, and temperature of lava. Over time the composition of magma feeding a volcano may change, generating volcanoes with complex shapes.

The current eruption, which started in late 2012, is on Tolbachinsky Dol, a lava plateau marked by small volcanic cones that formed during earlier eruptions. These cones stretch southwest from the summit of Plosky Tolbachick, a gently sloping shield volcano formed from layers of fluid lavas. Just to the west of Plosky Tolbachik lies Ostry Tolbachik, a steep-sided stratovolcano composed of layers of thick lava, ash, tephra, and other volcanic debris.


Figure 40: ALI image of Mount Etna observed on Feb. 20, 2013 (image credit: NASA, Ref. 50)

Legend to Figure 40: The false-color image combines shortwave infrared, near-infrared, and green light in the red, green, and blue channels of an RGB picture. This combination makes it easier to differentiate between fresh lava, snow, clouds, and forest. In the image, fresh lava is bright red, as the hot surface emits enough energy to saturate the instrument’s shortwave infrared detectors but is dark in near-infrared and green light. Snow is blue-green because it absorbs shortwave infrared light, but reflects near-infrared and green light. Clouds made of water droplets (not ice crystals) reflect all three wavelengths of light similarly and appear white. Forests and other vegetation reflect near-infrared more strongly than shortwave infrared and green, and so appear green. Dark gray areas are lightly vegetated lava flows, 30 to 350 years old.


Figure 41: ALI image of lava flow at Puyehue-Cordón Caulle, a volcano inChile, acquired on January 13, 2013 (image credit: NASA)

Legend to Figure 41: From June 2011 until April 2012, Puyehue Cordón Caulle, a Chilean volcano, erupted a massive obsidian lava flow. The flow covered roughly 16 km2 of land in lava about 30 m thick. This natural-color satellite image shows the flow on January 13, 2013. It was collected by ALI (Advanced Land Imager) on EO-1(Earth Observing-1). The dark gray lava stands out against the light gray ash and lava bombs that cover surrounding areas. 52)

A team of geologists visiting Puyehue in January 2013 discovered that the lava was still in motion even though the eruption had stopped. Unlike a crystalline rock, obsidian is not completely rigid: it can flow, even when solid. The higher the temperature, the faster a glass will deform, especially near its melting point. Volcanologist Hugh Tuffen described his experience approaching the flow: “The sound of advancing obsidian lava is quite fascinating and unlike anything I have ever heard—a succession of platey fracturing sounds, as if a bowl of rice crispies were made up of thousands of fragile plates that each broke, rather than the usual snap, crackle and pop.” The hot interior of the lava flow, insulated by a shell of solidified rock, allowed it to continue to ooze downhill (Ref. 52).

• The EO-1 mission is operational in 2012. On Nov. 21, 2011, EO-1 celebrated its eleventh anniversary on orbit. During its time on orbit, the satellite has accomplished far more than anyone envisioned, and its Earth-observing mission continues on.

- Since 2009, EO-1 has expanded its science role in disaster monitoring by taking advantage of its platform pointing capability, SensorWeb network, and the ability for user-level tasking. ALI provides NASA with the capacity to gap-fill between Landsat-7 and LDCM (Landsat-8), gap-fill for ASTER SWIR bands, and provide spaceborne prototyping for HyspIRI. - All components of EO-1 are predicted to function through 2015, although the orbit degradation will shift the equatorial crossing-time earlier by one hour.

- In June 2011, the NASA Earth Science Senior Review recommended an extension of the EO-1 mission up to 2013 as baseline and a further baseline to 2015. 53)

• Detailed Views of Erupting Nabro Volcano in Eritrea taken by ALI (Advanced Land Imager) on June 24, 2011. The Nabro volcano has been erupting since June 12, 2011. It sits in an isolated region on the border between Eritrea and Ethiopia and satellite remote sensing is currently the only reliable way to monitor the ongoing eruption. 54)

Since it began erupting on June 12, 2011, emissions from Eritrea’s Nabro Volcano have drifted over much of East Africa and the Middle East. Ash has displaced residents living near the volcano and disrupted flights in the region. Despite the volcano’s widespread effects, little is known about the eruption. Nabro is located in an isolated region along the border between Eritrea and Ethiopia, and few English-language reports have been published. Satellite remote sensing is currently the only reliable way to monitor the ongoing eruption.

The bright red portions of the false-color image (Figure 42) indicate hot surfaces. Hot volcanic ash glows above the vent, located in the center of Nabro’s caldera. To the west of the vent, portions of an active lava flow (particularly the front of the flow) are also hot. The speckled pattern on upstream portions of the flow are likely due to the cool, hardened crust splitting and exposing fluid lava as the flow advances. The bulbous blue-white cloud near the vent is likely composed largely of escaping water vapor that condensed as the plume rose and cooled. The whispy, cyan clouds above the lava flow are evidence of degassing from the lava.


Figure 42: This false color satellite image shows active lava flows of the Nabro volcano in Eritrea on June 24, 2011 (image credit: NASA)

The natural-color image (Figure 43) shows a close-up view of the volcanic plume and eruption site. A dark ash plume rises directly above the vent, and a short, inactive (cool) lava flow partially fills the crater to the north. A gas plume, rich in water and sulfur dioxide (which contributes a blue tint to the edges of the plume) obscures the upper reaches of the active lava flow. Black ash covers the landscape south and west of Nabro.


Figure 43: This natural-color image shows a close-up view of the volcanic plume and eruption site of the Nabro volcano (image credit: NASA)

• The EO-1 spacecraft and its payload are operating nominally in 2011.


Figure 44: ALI image of the volcanic island Ostrov Shikotan in the Pacific Archipelago of the Kuril Chain (image credit: NASA) 55)

Legend to Figure 44: The image was acquired on Feb. 14, 2011. Ostrov Shikotan is a volcanic island at the southern end of the Kuril chain. At about 43º North—more than halfway to the Equator—Shikotan lies along the extreme southern edge of winter sea ice in the Northern Hemisphere. The island is surrounded by sea ice — swirling shapes of ghostly blue - gray. Although sea ice often forms around Shikotan, the extent varies widely from year to year, and even day to day.

• Scheduled to fly for a year, designed to last a year and a half, EO-1 celebrated its tenth anniversary on November 21, 2010. During its decade in space, the satellite has accomplished far more than anyone dreamed of. 56)

The EO-1 is currently the only satellite that simultaneously acquires high spatial resolution (30 m) data for terrestrial and aquatic monitoring with two unique spectral instruments - the Hyperion and the ALI. It also has a 10 m pan-sharpening capability. As the only civilian satellite sensor now acquiring continuous spectrum hyperspectral imagery, Hyperion is greatly assisting investigators with the identification of ecosystem components and enabling the separation of various vegetated land covers. 57)

Because of this range of capabilities, the Hyperion has become the primary data source for numerous investigations which have demonstrated the utility of imaging spectroscopy for a broad range of ecological and geophysical studies around the world. These include examinations of the benefit of spectroscopy in applications relating to forestry, agriculture, species discrimination, invasive species, desertification, land-use, vulcanization, fire management, homeland security, as well as natural and anthropogenic hazards and disaster assessments. In addition, EO-1 has provided sensor information to support characterization for a number of instruments on existing orbital platforms (e.g., MODIS on Terra and Aqua, Landsat-5 & -7), and on future platforms (e.g., LDCM for Landsat-8; EnMAP, and HyspIRI). The EO-1 participates in numerous activities of the Committee on Earth Observation Satellites (CEOS) designed to characterize calibration/validation of ground targets (Ref. 57).


Figure 45: The ALI instrument on the EO-1 satellite captured a volcanic plume from Krakatau on November 17, 2010 (image credit: NASA)

Legend to Figure 45: The active volcano on Indonesia's island Krakatau was automatically monitored by the SensorWeb. ALI took this image on November 17, 2010 using the ASE (Autonomous Science Experinet), an onboard intelligent scheduling tool that allows the satellite to decide for itself which images Hyperion and ALI should take.

• The spacecraft and its payload are operating nominally as of 2010. NASA granted a mission extension to the end of 2011. In 2009 the Senior Review Panel recommended to NASA that an extended mission status will lead to provide important data for the LDCM (Landsat Data Continuity Mission), national operations in response to critical events, preparation for Decadal Survey mission(s), and science questions articulated in the NASA Science Plan.. There is great value in having advanced instruments with capabilities like Hyperion and ALI up and running in orbit for testing and development. 58) 59)

The unique status of EO-1 among NASA missions is notable. Although it was designed as a technology mission, EO-1 has successfully pioneered new techniques and now (since 2003) is in extended mode. It now collects data for special event imagery at high spectral and spatial resolution, contributes to long-term data sets and interoperability and contributes images and new techniques for coordinated synergistic data collections, especially for natural disasters and coastal monitoring efforts.

The ALI sensor has proven useful as a first look resource to assist in assessing disasters and events of a critical nature. This has been justified through the use of several thousands of images processed for disaster/weather monitoring and in the development of automated first response systems which automate data collection such as the innovative SensorWeb 2.0 development.

• The EO-1 spacecraft is operational - with the ALI and Hyperion instruments fully functional as of 2009 - more than seven years after its original retirement date. NASA/HQ has approved funding for operational services for the years 2008-2009. Toward the end of year 2009, the EO-1 mission will be re-evaluated and considered for an additional two years of operational funding. Aside from funding, it is projected that it would be necessary to terminate the mission at the end of year 2011 due to the onboard propellant having been exhausted.

Starting in late 2005, the EO-1 satellite orbit was channged and taken out of the constellation flying with Landsat 7.

- The new orbit parameters of EO-1 are: inclination = 98.2º, apogee = 710 km, perigee = 694 km. SMA (Semi Major Axis) = 7063 km, MLT = 9.996.

- The remaining fuel is being used to perform periodic inclination burns to achieve and maintain a descending node mean local time of slightly above 10:00 hours. 60)

• In early 2005, ASE is the primary mission planning and control system onboard EO-1. The use of automated planning onboard EO-1 has enabled a new system-of-systems capability - thus creating a sensor-web.

• The software package ASE (Autonomous Sciencecraft Experiment) has been flying on EO-1 since January 2004 (some initial onboard testing was conducted in late 2003). ASE started as a technology experiment. The technology was declared fully validated in May 2004 after all 20 onboard autonomy experiments were fully tested. The overall system performed as expected and was considered a success. 61)

• In 2004, the EO-1 satellite had first demonstrated its abilities by detecting and self-directing observations of an eruption of remote Erebus Volcano in Antarctica and delivering data to scientists within hours. - On May 7, 2004, EO-1 detected heat emission from the Erebus Volcano lava lake, alerted scientists on the ground, and automatically rescheduled the satellite to obtain more observations.

• After the technology validation, ASE has been running on EO-1 in support of over 4000 autonomous data acquisitions. In addition, over 400 closed-loop executions have been run where ASE autonomously analyzed science data onboard and triggered subsequent observations.



Extended mission phases of EO-1:

Following the first year of operations (having more than achieved its original technology validation goals), the EO-1 mission entered its extended mission phase, flying in an on-orbit testbed mode, in which additional validations were being performed which centered on the theme of enabling Sensor Webs. The EO-1 Sensor Web Enablement (SWE) experiment is implemented on OGC (Open Geospatial Consortium) architecture standards. 62) 63) 64) 65)

SWE is divided into a number of components, each of which contains models, services or XML encodings of various aspects of the Sensor Web. For example, the SensorML (Sensor Model Language) contains models and encodings for sensors, and the Observations and Measurements component contains the same for sensor observations and measurements.


ASE (Autonomous Sciencecraft Experiment):

In the fall of 2003, the ASE software package of JPL and GSFC was uplinked and installed on EO-1 (as a testbed). [Note: Initially, the ASE package was planned to be flown on the ST-6 (Space Technology-6) mission of NASA, but this was changed when a flight opportunity on EO-1 arose]. The ASE onboard flight software package architecture includes three autonomy software components: 66) 67)

1) CASPER (Continuous Activity Scheduling Planning Execution and Replanning). The CASPER software replan activities, including downlink, based on science observations in the previous orbit cycles.

2) SCL (Spacecraft Command Language) package to enable event-driven processing and low-level autonomy. SCL acts as a robust middleware execution engine.

3) Onboard science processing algorithms that analyze the image data to detect trigger conditions such as science events, “interesting” features, changes relative to previous observations, and cloud detection for onboard image masking

There is also a complementary ground piece of software called ASPEN (Automated Scheduling and Planning Environment), which works collaboratively with CASPER to process triggers and imaging requests for use by CASPER. In particular, triggers and requests are automatically sorted and prioritized and then converted into a list of high-level goals that can be worked on by CASPER.

The overall ASE objective is to demonstrate several integrated autonomy technologies to enable autonomous onboard science applications. Several science algorithms including: onboard event detection, feature detection, change detection, and unusualness detection are being used to analyze science data, in particular those of Hyperion. The instrument images a 7.5 km by 42 km target area per image and provides detailed spectral mapping across all 220 channels with high radiometric accuracy. The first autonomy experiment on EO-1 used data from Hyperion, to demonstrate the ability to automatically discriminate between clouds, ice/snow, and other high reflection land features onboard the satellite. 68) 69) 70)

These capabilities were also linked to other satellites and ground instruments to form various ad hoc sensor webs enabling various operations concepts that would be useful for future NASA missions. The sensor web and autonomy capabilities developed thus far have been retained and used operationally on EO-1 and have served to reduce the cost of operations by over 50%.

ASE on EO-1 is demonstrating an integrated autonomous mission using onboard science analysis, replanning, and robust execution. The ASE performs intelligent science data selection that leads to a reduction in data downlink. In addition, the ASE increases the science return through autonomous retargeting.

The Livingstone software package version 2 (LV2) is being used in EO-1 on-orbit testbed demonstrations. The AI (Artificial Intelligence) software package “Livingstone” was developed by a team at NASA/Ames Research Center (S. Hayden) and uplinked to EO-1 in the summer of 2004. The LV2 automatically detects and diagnoses simulated failures in the EO-1 payload instruments. The software package has the ability to find and analyze errors in the spacecraft's systems. A key feature of the software is its “reasoner” function, which enables it to compare the predicted performance of a system based on readings from onboard monitors. Contradictions between the predicted and actual performance are used to identify the failures. The autonomous diagnostic tool of LV2 can help controllers in identifying and detecting a potential problem in sufficient time to make repairs or to find a work-around solution. 71) 72) 73)


Figure 46: The operations flow in the extended mission period (image credit: NASA/JPL)

ASE on EO-1 demonstrates an integrated autonomous mission using onboard science analysis, replanning, and robust execution. The ASE performs intelligent science data selection that will lead to a reduction in data downlink. In addition, the ASE will increase science return through autonomous retargeting. Demonstration of these capabilities onboard EO-1 will enable radically different missions with significant onboard decision-making leading to novel science opportunities. The paradigm shift toward highly autonomous spacecraft will enable future NASA missions to achieve significantly greater science returns with reduced risk and reduced operations cost.



EO-1 demonstrations in Sensor Web applications:

In 1999, ESTO (Earth Science Technology Office) of NASA initiated its AIST (Advanced Information Systems Technology) program with the objective to identify, develop and (where appropriate) demonstrate advanced information system technologies. The goal of NASA's Sensor Web approach is to employ new data acquisition strategies and systems for integrated Earth sensing that are responsive to environmental events for both application and scientific purposes. Sensor Webs can achieve science objectives beyond the abilities of a single platform by:

1) Reducing response time (where events unfold rapidly or where time is otherwise constrained)

2) Increasing the scientific value, quantity, or quality of the observation (where unique science criteria are met, or when co-incident observations are possible) by enabling collaboration among sensing and analysis assets.

The EO-1 spacecraft has been used and continues to be used as the core satellite to demonstrate various mission autonomy technology. In this context, a “Sensor Web” is defined as a coherent collection of spaceborne and/or ground-based sensors and computation nodes linked by a communication fabric that collectively act as a single, dynamically adaptive, observing system. An example sensor web might be one satellite observing a target event and triggering another satellite autonomously. 74) 75) 76)

The key elements of the sensor web include autonomous detection of events, autonomous monitoring of detection notifications, autonomous generation of observation requests, and autonomous rescheduling of observations to acquire data of higher temporal, spatial, and spectral resolution.

The sensor web architecture has the following design objectives:

• Enable autonomous tasking of the EO-1 spacecraft in response to ground-based science events

• Support a diverse number of sensor sources and science event types

• Enable reaction observations to be serviced and uploaded promptly in order to maximize the responsiveness of the sensor web

• Provide autonomous detection algorithms onboard the satellite

• Deliver detection notifications from the onboard algorithms to the ground

• Allow for detailed tracking of changes made by the sensor web to the mission operations schedule

• Minimize the impact on the EO-1 operations staff and procedures.


Figure 47: Sensor-web detection and response architecture (image credit: NASA/JPL)

The EO-1 sensor-web architecture consists of a number of components which operate in the following sequence of steps: 77) 78)

1) Asset1 acquires data (usually global coverage at low resolution)

2) Data from Asset1 is downlinked

3) This data is automatically processed to detect science events

4) Science event detections are forwarded to a re-tasking system. This system generates an observation request which is forwarded to an automated planning system.

5) The automated planning system then generates a command sequence to acquire the new observation.

6) This new command sequence is uplinked to Asset2 (EO-1) which then acquires the high resolution data.

7) The data is then downlinked, processed, and forwarded to the interested science team.

In this operational setup, Asset2 functions as the EO-1 spacecraft. The automated re-tasking element of the sensor-web consists of several components working together as follows:

• Science tracking systems for each of the science disciplines automatically acquire and process satellite and ground network data to track science phenomena of interest. These science tracking systems publish their data automatically to the internet each in their own format. In some cases this is via the http or ftp protocol, in some other cases via email subscription and alert protocols.

• Science agents either poll these sites (http or ftp) to pull science data or simply receive emails to receive notifications of ongoing science events. These science agents then produce « science event notifications » in a standard XML (eXtensible Markup Language) format which are then logged into a « science event » database.

• The science event manager processes these science event notifications and matches them up with « science campaigns ». When a match occurs, an observation request is generated.

• The observation requests are processed by the ASPEN (Automated Scheduling and Planning ENvironment) automated mission planning system. ASPEN integrates these requests and schedules observations according to priorities and mission constraints.

• For observations that are feasible, an observation request is uplinked to the spacecraft.

• Onboard EO-1 the ASE software will accommodate the observation request if feasible. In some cases onboard software may have additional knowledge of spacecraft resources or may have triggered additional observations so some uplinked requests may not be feasible.

• Later, the science data is downlinked, processed, and delivered to the requesting scientist.

The EO-1 spacecraft has been networked with other satellites and ground sensors. This network is linked by software and the Internet to an autonomous satellite observation response capability. The system is designed with a flexible, modular, architecture to facilitate expansion in sensors, customization of trigger conditions, and customization of responses.

The EO-1 sensor-web has been used to implement a global surveillance program of science phenomena including: volcanoes, flooding, cryosphere events, and atmospheric phenomena.


Figure 48: Conceptual view of the Sensor Web architecture (image credit: NASA)


VSW (Volcano Sensor Web):

In the timeframe 2004/5, JPL implemented VSW (Volcano Sensor Web) for a period of 18 months in which data from ground-based and space-based sensors, that detect current volcanic activity, are used to automatically trigger the EO-1 spacecraft, to make high spatial-resolution observations of these volcanoes.

The fully-automated process allows for rapid acquisition and transmission - typically within 48 hours, though theoretically possible within 2-3 hours - of data products containing the most useful data content, namely the numbers, locations, and spectra of hot pixels. This information allows scientists to evaluate the instantaneous eruption extent and intensity. Prior to VSW, this process took weeks. In the future, the sensor web could become an integrated network of ground, airborne, and orbiting sensors that will enable seamless, rapid, autonomous reactions to the detection of volcanic activity. 79)


Figure 49: The Volcano Sensor Web at the end of 2007, including two-way triggering between spacecraft and in situ sensors (image credit: NASA)

MSW (Model-based Volcano Sensor Web). The MSW represents an advance beyond the simple detection-response operation mode, where an alert of activity generates a request for a spacecraft observation with, generally, no deeper understanding of the magnitude or extent of the eruption that was taking place. 80) 81)

The goal of the MSW is to have asset operations based on determining what additional information and data are needed to understand the state of a volcanic eruption. The required information flow between sensor web assets is performed through web services. The web service interfaces are defined by the Open Geospatial Consortium (OGC) Sensor Web Enablement (SWE) and allow for describing a process flow, enabling extraction of higher-level information from datasets, exchanging of metadata, determining the quality of data, acquiring instrument and data information, communicating between sensors, and discovery of assets, data, and data products and observation requests.


Figure 50: Schematic layout of the Model-driven Volcano Sensor Web (image credit: NASA)

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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.