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Satellite Missions Catalogue

MightySat Program

Last updated:Jun 13, 2012



Mission complete


Hyperspectral imagers




Quick facts


Mission typeEO
Mission statusMission complete
Launch date14 Dec 1998
End of life date12 Nov 2002
Instrument typeHyperspectral imagers, In situ
CEOS EO HandbookSee MightySat Program summary

MightySat Program

MightySat is a long-term, multi-mission, small satellite program (started in 1994) of the Space Experiments Directorate of Phillips Laboratory (USAF/PL) at Kirtland Air Force Base, Albuquerque, NM (as of 1998 designated as AFRL (Air Force Research Laboratory). The overall program objectives are to provide an environment for frequent, inexpensive, on-orbit demonstrations of emerging space system technologies and to accelerate their transition into operational use. The MightySat spacecraft are modular and functionally standardized platforms and buses capable of supporting a wide range of experimental payloads. Launches are considered from STS (Shuttle) or with multi-service launch systems. The emphasis is on low-cost projects, with fast building periods from contract to launch, with high-risk and high-payoff technologies. 1) 2) 3)


MightySat I Spacecraft

MightySAT I is a single mission project which uses an all-composite bus (similar in configuration to its aluminum XSAT predecessor of NASA) with significant mass reduction. The S/C was built by OSC (Orbital Sciences Corporation) of McLean, VA (formerly CTA Space Systems). The S/C structure is characterized by three decks supported by six structural frames which make up a hexagonal prism body.

The S/C mass is 63 kg, the payload mass is 17 kg (average power of 12 W), orbit average power = 14-27 W. The S/C is spin-stabilized (3 rpm) with the spin axis oriented normal to the orbit plane. Attitude sensing and actuation is provided by a three-axis magnetometer, two coarse sun sensors, and three torque coils. A coarse attitude knowledge of ±5º is required. S/C design life of one year. 4) 5)

Figure 1: Illustration of the MightySat I microsatellite (image credit: J. Heyman)
Figure 1: Illustration of the MightySat I microsatellite (image credit: J. Heyman)

The S/C has four subsystems: C&DH (Command & Data Handling) made up of nine electronic boards, EPS (Electrical Power Subsystem) consisting of seven solar panels and a single 21-cell NiCd battery (4 Ah), RF (Radio Frequency - communications) consisting of a receiver, transmitter, transmit/receive switch, and a four-blade antenna, and ADACS (Attitude Determination and Control Subsystem). Data transmission is in UHF-band (306.775 MHz) at 9600 bit/s (downlink, BPSK modulated) and at 2400 bit/s (uplink, FSK modulated). The link margin is about 20 dB for uplink and downlink above a 5º elevation angle.



MightySat I was launched by ejection from the Space Shuttle using HES (Hitchhiker Ejection System) on Dec. 14, 1998. The entire satellite, HES, and avionics were stowed away into a standard Hitchhiker canister of 0.14 m3 in volume prior to launch. Deployment was initiated by the Shuttle crew on Dec. 14, 1998. The entire MightySat I deployment sequence took about 10 minutes. The launch of Shuttle flight STS-88 (Endeavour) took place on Dec. 4, 1998 (landing of STS-88 on Dec. 15 at KSC). 6) 7)

Note: The main payload on the STS-88 assembly flight mission to the ISS was the US module Unity with a mass of 11,800 kg. Secondary payloads on the STS-88 mission included the IMAX Cargo Bay Camera (ICBC), the Argentinean Scientific Applications Satellite-S (SAC-A), the MightySat I Hitchhiker payload, the Space Experiment Module (SEM-07), and the GAS (Getaway Special) G-093 sponsored by the University of Michigan.

MightySat mission operations are conducted by SMC/TEO at Kirtland AFB. All communications are via AFSCN (Air Force Satellite Control Network). The operations concept is largely based upon experience from similar satellite programs, most notably RADCAL.

Orbit: altitude = 385 km (initial altitude), inclination = 51.6º.


Mission Status

The MightySat I spacecraft reentered the Earth's atmosphere on Nov. 16, 1999 due to its relatively low orbital altitude. All of the mission objectives were accomplished.


Payload Complement

MightySat I has five AFRL-developed advanced technology demonstration experiments. Two of the demonstrations are tests of experimental bus components.


ACS (Advanced Composite Structure)

Objective: technology test. The S/C structure (a hexagon with a diameter of 50 cm and a height of 53 cm with a mass of about 8 kg) consists of a composite frame, three decks, and seven solar panel substrates. The composite material is graphite fiber. However, ACS has no data interfaces to the S/C (the structure was ground-tested). The S/C frames were fabricated by using the so-called “SnapSat” approach (developed by Composite Optics Inc.), in which the elements are cut from cured flatstock layups and fitted together using a mortise&tenon technique.


ASCE (Advanced Solar Cell Experiment)

Objective: advance of in-space power generation technology. Use of dual-junction solar cells (average efficiency of 21%) which provide a 15% performance gain over conventional GaAs cells. The cells are covered with a GaInP (gallium indium phosphate) layer which captures and converts the short wavelength spectrum. ASCE consists of 13 strings of 40 GaAs cells (2cm x 4 cm) and six strings of 18 GaInP cells (2 cm x 2 cm) for comparison of cell performance. The cells are bonded directly to the solar panel substrates. An objective is also to determine aging effects in the space environment.


MAPLE-1 (Microsystem and Packaging for Low Power Electronics)

The objective is to provide an on-orbit demonstration of advances in low-power electronics including performance (emerging electronics and packaging technologies in space). The MAPLE-1 experiment suite is a collection of five sub-experiments, brokered by a central controller that supplies power and communications through interfaces to the MightySat I host satellite. MAPLE sub-experiments explore flight-worthiness, issues associated with low-power, commercial and radiation tolerant microelectronics, advanced microelectronics packaging, MEMS (Micro-Electromechanical Systems) devices, circuits, associated subsystems, and components. Particular experiments are:

• In-situ operation of MEMS commercial cantilever beam accelerometers. Two accelerometers are used, a local 5 gram full-scale unit and a remote 2 gram full-scale unit.

• Monitoring of advanced packaging structures through reliability monitoring integrated circuit die (monitoring of moisture, dust, and material property changes in electronic components), developed by SNL. The SSRB (Solid State Recorder Board) was designed to evaluate the performance of HCSM (High Capacity Spaceborne Memory) SRAM module in the space environment. The EMB (Environment Monitor Board) was designed to measure the temperature inside the Maple-1 enclosure.

• Total ionizing dose dosimetry

• Advanced thin film high density interconnect multi-chip modules

• Comparison of military grade commercial and radiation hardened bulk silicon anti-fuse-based field-programmable gate arrays.

Figure 2: Photo of MAPLE-1 modules (image credit: AFRL)
Figure 2: Photo of MAPLE-1 modules (image credit: AFRL)


SMARD (Shape-Memory Actuated Release Device)

Objective: demonstration of a new class of low-shock release devices. SMARD devices are based upon a shape-memory alloy (Nitinol) which is used as the driving force to actuate the release of a fastener. - The SMARD payload consists of four release devices mounted on a common, instrumented deck (a conventional pyrotechnic device, a linkwire device, and two shape-memory actuated devices are used). The SMARD experiment involved measuring the response of a three-axis MEMS accelerometer to the shock wave generated when a test “separation bolt” was fired.


MPID (Micro-Particle Impact Detector)

MPID is an instrument of NASA/LaRC. The objective is to collect information on orbital debris. Measurement of direction and time of impact of spaceborne micro-particles with time of impact resolution of 0.1 s. The primary element in this experiment consists of two MOS (Metal Oxide Semiconductor) discharge capacitor detectors that discharge upon hypervelocity particle impact. MPID consists of two small plates (4 cm x 8 cm, each the size of a credit card), mounted on the bottom outside surface of the satellite, providing indications of micro-particle impacts (recording of time of impact). Each particle impact causes an impact event record that is stored in the S/C control unit for later downlink. Each impact event record stores time of impact and output from two coarse sun sensors. Data from the coarse sun sensors is used to help determine attitude of the spacecraft. 8)

Figure 3: Illustration of the MPID electronic board within the S/C system (image credit: NASA/LaRC)
Figure 3: Illustration of the MPID electronic board within the S/C system (image credit: NASA/LaRC)

MightySat II.1 (Sindri P99-1) Mission

MightySat II.1 is a technology demonstration mission of the US Defense Space Test Program (test of high-risk, high-payoff space system technologies), a joint project of SMC and AFRL. The MightySat II program, initiated in March 1996, represents a series of up to five small satellite missions over a decade.


MightySat II.1 Spacecraft

The S/C structure is of modular design built by General Dynamics (formerly Spectrum Astro Inc.) of Gilbert, AZ. The spacecraft (SA-2008 bus) has dimensions of 68.6 cm x 89 cm x 89 cm. The size of the payload envelope is: 61 cm x 61 cm x 46 cm. The S/C is three-axis stabilized (zero momentum based, inertial pointing); attitude knowledge/control = 0.15º/0.18º.

The S/C has deployable solar arrays (Si) with 2-axis articulation. The total mass of the minisatellite is 123.7 kg (37 kg of payload mass); power = 330 W (EOL, 100 W average) with 28 V unregulated bus. In addition, there are three NiCd batteries, each with 4 Ah capacity for eclipse operations. A RAD6000 CPU is used (Quad TMS320C40 (QC40) Floating Point Digital Signal Processor). The S/C design life is 1 year + 3 months extension.

Figure 4: Photo of the MightySat-II.1 spacecraft without FTHSI instrument (image credit: General Dynamics)
Figure 4: Photo of the MightySat-II.1 spacecraft without FTHSI instrument (image credit: General Dynamics)

Attitude determination is provided with a star tracker and an interferometric fiber optic gyro, attitude control (actuation) uses three orthogonal reaction wheels. Secondary control may be achieved with three torque rods, the torque rods serve also to dissipate reaction wheel momentum. The ADCS (Attitude Determination and Control Subsystem) also autonomously controls solar array articulation. The command and data handling subsystem provides 380 MByte of solid state storage at data rates of 20 MByte/s.

Structure & Thermal subsystem

Composite primary bus structure
Paraffin wax deployment mechanisms
Passive, cold-biased system using local radiators
Thermostatically controlled heaters (contingency only)

ADCS (Attitude Determination and Control Subsystem)

3-axis stabilized with RWA's, Zero Momentum Biased (ZMB)
Sun sensor, star tracker, IMU
Pointing Accuracy (3σ): 648 arcsec
Pointing Knowledge (3σ): 540 arcsec
Attitude Jitter (3σ): 15.7 arcsec/sec

C&DH (Command & Data Handling) subsystem

RAD6000 CPU @ 20 MIPS, IEEE VME backplane
128 MByte CPU RAM, 21.6 MByte/s transfer rate
2 Gbit solid state recorder for science data

General spacecraft parameters

Dimensions, stowed: 0.67 m W x 0.83 m L x 0.86 m H
Propellant: none
Design reliability & life: 0.8 @1 year

Spacecraft mass & power

Launch mass: 123.7 kg, bus mass: 87.1 kg
Power Load (OAP): 90 W bus; 60 W PL
Solar Array: Si, 2-axis articulated, 330 W EOL
Battery: 12 Ah, NiCd

Communication links

SGLS compatible
1 Mbit/s downlink for payload/experiment data
2.0 kbit/s command uplink
20 kbit/s telemetry downlink (TT&C)

Table 1: Overview of the MightySat II.1 performance characteristics 9)



The MightySat II.1 (dubbed “Sindri”) spacecraft was launched on July 19, 2000 from VAFB, CA. The launch vehicle was a Minotaur of OSC (a converted Minuteman-2 missile motor combined with the upper stages of the Pegasus-XL booster).

Orbit: Sun-synchronous circular orbit, altitude = 556 km, inclination = 97.3º, local crossing at 11:15 on a descending node.

RF communications are in S-band using a miniature SGLS (Space-to-Ground Link Subsystem) transponder of NRL (1.5 kg of SGLS mass - see NSX below). TT&C data rates of 2 kbit/s in uplink and 20 kbit/s in downlink, 1 Mbit/s downlink for payload data (encrypted uplink/downlink). S/C operations are conducted by SMC/TEO from Kirtland AFB. Obviously, the payload downlink was considerably undersized for a hyperspectral imager. However, the objective was simply to demonstrate the technology on a low-cost spacecraft - and not the production of large-scale observations. 10)

Figure 5: Artist's view of the deployed MightySat II.1 spacecraft (image credit: Spectrum Astro)
Figure 5: Artist's view of the deployed MightySat II.1 spacecraft (image credit: Spectrum Astro)


Mission Status

• The MightySat-II.1 spacecraft reentered the atmosphere on Nov. 12, 2002. 11)

Figure 6: Ground track of the last orbit of MightySat-II.1 (image credit: The Aerospace Corporation)
Figure 6: Ground track of the last orbit of MightySat-II.1 (image credit: The Aerospace Corporation)

• In September 2001, all 10 experiments were declared to have achieved 100% mission success, including FTHSI.

• For 14 months after launch, spacecraft mission operations were conducted by SMC/TEO at Kirtland AFB, then the operations were turned over to a another user (Oct. 2002). Cooperative image collections of FTHSI were performed with AVIRIS and with MWIR of TSX-5 (Tri-Service Experiments Mission 5).

• In August 2002 the spacecraft has been turned off due to orbit decay.

• On Aug. 1, 2000, within days of a picture-perfect launch on July 19, the Fourier Transform Hyperspectral sensor aboard MightySat II.1 sent back its first ”hypercube” from orbit.

Sensor Complement

MightySat II.1 carries a total of ten new technologies, including both experimental bus components and stand-alone experiments. The FTHSI is the prime payload of the mission.


FTHSI (Fourier Transform HyperSpectral Imager)

FTHSI was designed and built by Kestrel Corporation of Albuquerque, NM, and the Florida Institute of Technology, Melbourne, FL, heritage of airborne version of FTVHSI. The objective is to demonstrate spaceborne hyperspectral imaging technologies (FTHSI is the first functioning spaceborne hyperspectral imager). The FT-approach is considered to be promising for spaceborne hyperspectral concepts (FTHSI demonstrates the advantage of Fourier systems over dispersive hyperspectral imagers, in that it can record the full spectra without any time delay and can decouple the spatial and spectral signatures). 12) 13) 14) 15) 16)

The interferometer is an innovative solid block design (Figure 8) that supplies an optimal path offset via a design that is extremely stable and immune to damage or misalignment during launch and on-orbit operations. The FTHSI instrument contains three major optical subsystems: a monolithic Sagnac interferometer which produces the spatially modulated interferogram; a Fourier transform lens, which frees the spectral properties of dependence on aperture geometry and allows the wide FOV; and a cylindrical lens, which re-images one axis of the input aperture onto the detector array providing the one dimension of imaging. The 1-D image is passed through the interferometer where the rays are split, slightly separated, and recombined to create an interference pattern (interferogram) in one dimension. From the interferometer, a Fourier lens collimates the light and a cylindrical lens images the energy onto the detector array, preserving the one by n spatial dimension.

Figure 7: Exploded view of the FTHSI instrument (image credit: Kestrel Corp.)
Figure 7: Exploded view of the FTHSI instrument (image credit: Kestrel Corp.)
Figure 8: Schematic view of the FTHSI solid block interferometer (image credit: Kestrel Corp.)
Figure 8: Schematic view of the FTHSI solid block interferometer (image credit: Kestrel Corp.)

Spectral range

350 - 1050 nm

Spectral resolution

85.4 cm-1 (1.7 nm @ 450 nm)

Number of usable spectral bands


FOV (Field of View)

3.0 º

IFOV (Instantaneous Field of View)

0.0058º or 0.0029º

Spatial resolution (best cross/along track)

28 m x 30 m

Swath width

7-29 km

Scene length range of imagery

10 km to 473 km

Instrument pointing (control/knowledge)


Instrument mass, power

20.45 kg, 66/60 W (peak/average)

Source data rate

20 MByte/s

Table 2: Some performance characteristics of FTHSI

The optical system of FTHSI employs a Ritchey-Chretien telescope with a clear aperture of 165 mm. The camera is an adaptation of a commercial camera made by Silicon Mountain Devices. The system has an f/3.4 number in the spatial and an f/5.3 number in the spectral dimension. A solid-block design is used to maintain alignment of the Fourier optics and cylinder lens. The detector assembly consists of a large-format CCD array (1024 x 1024) with a 12 bit full-frame readout capability at up to 120 frames/s. Four operating modes are available to vary on command the spectral and spatial imagery resolutions, respectively: 512 x 512, 512 x 1024, and 1024 x 1024.

Figure 9: Block diagram of the FTHSI data-handling system (image credit: Kestrel Corp.)
Figure 9: Block diagram of the FTHSI data-handling system (image credit: Kestrel Corp.)

The instrument is operated with a maximum of 137 usable bands out of the 256 available, required by the limiting Nyquist sampling rate (to improve SNR). Note: an FT-type sensor creates data from zero wavenumbers to the Nyquist cutoff. Since the sensor cannot see wavenumbers lower than about 9523 cm-1 (1050 nm), there is information lost due to the sensor detectivity. The higher spectral resolution mode that operates the detector in the completely unbinned mode is not being used.

Data management: 17) The FTHSI instrument takes advantage of the VME satellite bus services by storing source data directly in an allocated RAM of 134 MByte. The data transfers are provided by the custom-built HII (Hyperspectral Imager Interface) card, which handles packing and delivery to the S/C memory, it serves also as the RS 232 controller interface to the camera. A typical data collect (observation) lasts from 6-30 s. Total scene sizes are limited by the resolution selected and the storage capacity of 134 MByte RAM.


Selections available

Default mode

Camera frame rate (frames/s)

15, 30, 60, 110


Camera gating time (s-1)

1, 1/125, 1/250, 1/500, 1/1000


Data bit depth

12 or 8 bit

12 bit

Nr. of imaging tiles per record

1, 2 or 4


Camera binning modes

1 x 1, 1 x 2, 2 x 1, 2 x 2

2 x 2

Data block length

0 to 134 MByte

134 MByte

Camera gain

1 or 4


Table 3: FTHSI camera operating modes and performances

FTHSI illustrated the technical advantages of Fourier systems over dispersive hyperspectral imagers in that it recorded the full spectra without any time delay and decoupled the spatial and spectral signatures. The spectral and spatial resolution of the device could be varied electronically through one of four modes, and the imager could ”see” the Earth in hundreds of spectral bands ranging from the visible to the far-infrared. For over 16 months, Kestrel's FTHSI observed the Earth from the MightySat II.1 satellite, an AFRL program for testing emerging technologies. 18)

Figure 10: View of the FTHSI instrument on top of the MightySat II.1 spacecraft during integration (image credit: Kestrel Corp.)
Figure 10: View of the FTHSI instrument on top of the MightySat II.1 spacecraft during integration (image credit: Kestrel Corp.)


SAC (Solar Array Concentrator)

SAC focuses more light on each solar cell, increasing the solar energy available and reducing the number of cells needed to produce the same amount of power. The objective is to test the new solar panel array design. The steerable SAC is capable to focus more light on each solar cell, increasing the solar energy available. SAC covers one-third of one of the four solar panels. The cells are made of GaAs material (22-24% efficiency). The SAC provides a concentration ratio of 3:1 and has a 20º pointing tolerance toward the sun. SAC generates 37 W (BOL, 27.5 W EOL), its mass is 0.74 kg. SAC was designed and developed at COI (Composite Optics, Inc.) of San Diego, CA.


MFCBS (Multi-Functional Composite Bus Structure)

The spacecraft design incorporates new materials to be flight-tested as a lighter, more flexible alternative to the traditional aluminum bus structure. The MFCBS design includes an integrated thermal management system, robust structural integrity, high-attitude control accuracy, and precision three-axis stabilization. MFCBS was developed at COI.


NSX (NRL SGLS Transponder)

NSX with SGLS is referring to “Space-to-Ground Link System.” NSX is a miniaturized satellite communications unit, about 70% smaller and lighter than the current industry product, the objective is to test its functionality. NSX has a mass of 1.5 kg and a size of 14.5 cm x 14.5 cm x 7 cm. It consists of a COMSEC (Communications Security) transmitter and receiver segment and is connected to the VME card. An encryptor/decryptor feature is provided in COMSEC. NSX requires 6 W in receive mode and 23.2 W in transmit/receive mode.


QC40 (Quad-TMS320C40 Processor)

Objective: test of the space radiation susceptibility of critical new forms of microelectronic components through an advanced, and high-speed processor. In particular, QC40 performs on-board processing and compression of the FTHSI raw data (also interferogram conversion), increasing the number of images that may be downlinked to a ground station. Feature extraction involves comparing and matching collected interferograms with on-board interferograms of known materials. Real-time center burst extraction allows ground operators to obtain a quicklook of the collected image to assess its further handling. - QC40 consists of two VME cards. One board is comprised of four COTS TMS320C40 microprocessors (120 Megaflop), while the other board provides the interface between the processor board and the S/C electronics. QC40 has a mass of 1.4 kg, the power is 17 W.


SMATTE (Shape Memory Alloy Thermal Tailoring Experiment)

Objective: an experiment involving actuation of a bi-modal composite sheet with respect to performing vibration isolation, structural control, deployment, and separation functions. SMATTE can change its physical properties, such as stiffness, damping, and shape, as a function of tailored thermal input signals. However, when SMATTE is heated above its transition temperature, it returns to its memorized shape. SMATTE consists of a layer of polymer matrix composite with a thin strip of shape memory alloy on each side; it is interfaced with the VME card and uses 5.3 W of power. Stress induced from thermal warping of the composite is automatically relieved by opposing stress in the shape memory alloy film. Optical fiber strain gauges monitor the performance of SMATTE.


SAFI (Solar Array Flexible Interconnect)

SAFI incorporates copper leads embedded in a flexible, composite film reducing the weight and complexity of traditional hard wiring (SAFI connects the solar cells and routes the current off the panel). The objective is to test a technology which might lead toward multi-functional bus technology. SAFI is bonded to the solar panel, it has a thickness of <0.25 mm.



MightySat II.1 flies the Aerospace Corporation's designed and built launcher assembly and two tethered sub-satellites. They are referred to as PICO20 and PICO22, according to their nodal positions in a prescribed array of picosats. Both PICOSATs are of the same hardware design as the PICOSAT1.0 assembly, flown on OPAL (launch of OPAL on Jan. 26, 2000). The software of PICOSAT1.1 was updated to reflect the lessons learned by the ground operations from the OPAL flight. The overall objectives are:

• To extend the PICOSAT1.1 mission life (to about a week) by accurate ejection knowledge and improved coordination with the USAF SSN (Space Surveillance Network)

• Test of improved flight software, designed to reduce power usage - to extend mission life

• To demonstrate the system functionality after long-term storage on a host S/C. The PICOSAT1.1 system is planned to be ejected after one year of the MightySat II.1 launch, in July 2001.

• To demonstrate repeatability and robustness of a low-cost platform for advancing MEMS in space systems applications

• To demonstrate radically new operating concepts for future systems. The PICOSAT1.1 experience will serve as a low-cost risk mitigator for the release and flight of small bodies from a primary S/C, in support AFRL and DARPA small-satellite programs like TechSat 21, XSS (Experimental Spacecraft System), and MEPSI (MEMS-based PicoSat Inspector), the latter one was also flown on STS-113 (Nov. 24 - Dec. 4, 2002).

Each PICOSAT, a box of size: 25 mm x 75 mm x 100 mm and a mass of 0.275 kg, uses a small, battery operated, very low power radio with a ground-link and a cross-link (PICOSAT to PICOSAT) capability. The tether, of 32 m length, contains a gold wire, serving as radar 'target' to facilitate ground based radar tracking (with SSN). A sensor board in each PICOSAT has one chip containing 4 MEMS RF switches (a 100 V charge pump on the sensor board is needed by the switches) that are operated in space after deployment. 19)

Figure 11: Illustration of the PICOSAT launcher concept (image credit: SSDL of Stanford University)
Figure 11: Illustration of the PICOSAT launcher concept (image credit: SSDL of Stanford University)

The picosat pair (PICO20 and PICO22) were stored onboard the host spacecraft for a period of 13 months to demonstrate proper delay deployment functioning. Successful deployment occurred in Sept. 2002 (immediate reception of the beacons from each PICOSAT - subsequently power was lost terminating the collection of data). The event demonstrated the ideas of onboard and on-call operations. 20)


SOR (Starfire Optical Reflectors)

The Starfire Optical Reflectors are utilized by the Starfire Optical Range, an optical research facility at Kirtland Air Force Base, to develop optical sensing, imaging, and propagation technologies to support Air Force aerospace missions. Two COTS optical reflector mirrors are affixed to the coarse sensor boom on the S/C top deck (the reflectors are passive with a mass of 1.76 kg). The Starfire Optical Range at Kirtland performs active ranging of satellite position via ground-based laser.


1) J. Freeman, C. Rudder, P. Thomas, “MightySat II: On-orbit Lab Bench for Air Force Research Laboratory,” Proceedings of the 14th Annual AIAA/USU Conference on Small Satellites, SSC00-I-2, Aug. 21-24, 2000

2) R. J. Davis, J. F. Monahan, T. J. Itchkawich, “MightySAT I: Technology in Space for about a Nickel,” Proceedings of the 10th Annual AIAA/Utah State University Conference on Small Satellites, Sept. 16-19, 1996

3) B. Braun, R. Davis, T. Itchkawich, T. Goforth, “MightySat-I: In Space,” Proceedings of 13th Annual AIAA/USU Conference on Small Satellites, Logan Utah, Aug. 23-26, 1999, SSC-99-I-3


5) R.J. Davis, J. F. Monahan, T. J. Itchkawich, “MightySat 1: Technology in space for about a nickel,” URL:

6) B. Braun, R. Davis, T. Itchkawich, T. Goforth, “MightySat-I: Transitioning Space Technology to the Warfighter,” AIAA-99-4484, 19999


8) P. J. Serna, G. H. Liechty, C. L.Neslen, R. Del Frate, E. Draper, “Micro-Particle Impact Detector Experiment On MightySat I,” Final report, 15 Jul 96-1 Oct 97, URL:

9) MightySat II.1 - a standard-interface demonstration smallsat,” General Dynamics, 2005, URL:

10) “NRL Miniature SGLS Transponder Launched on MIGHTYSAT II.1, “ URL:


12) L. J. Otten III, A. D. Meigs, B. A. Jones, P. Prinzing, D. S. Fronterhouse, R. G. Sellar, J. B. Rafert, C. Hodge, “The engineering model for the MightySat II.1 hyperspectral imager,” Proceedings of the Sensors, Systems and Next Generation Satellites, SPIE Vol. 3221-54, Sept. 1, 1997, London, UK, pp. 412-420

13) Summer Yarbrough, Thomas R. Caudill, E. T. Kouba, V.ictor Osweiler, James Arnold, Rojan Quarles, Jim Russell, L. John Otten, B. Al Jones, Ana Edwards, Joshua Lane, Andrew D. Meigs, Ronald Lockwood, Pete Armstrong, “MightySat II.1 hyperspectral imager: summary of on-orbit performance (Invited Paper),” Proceedings of the Conference on Imaging Spectrometry VII, San Diego, CA, Aug. 2001, SPIE Vol 4480, pp. 186-197, URL:

14) J. B. Rafert, L. J. Otten, E. W. Butler, A. D. Meigs, “Satellite sends hyperspectral images from space,” May 1, 2001, URL:

15) A. Barducci, P. Marcoionni, I. Pippi, “Recent advances in Earth remote sensing: Fourier Transform Stationary HyperSpectral Imagers,” Annals of Geophysics, Vol. 49, No 1, Feb. 2006, URL:

16) L. J. Otten, A. D. Meigs, B. A. Jones, P. Prinzing, D. S. Fronterhouse, “Payload qualification and optical performance test results for the MightySat II.1 hyperspectral imager,” Proceedings of SPIE, 'Sensors, Systems, and Next-Generation Satellites II,' Hiroyuki Fujisada; Ed., Vol. 3498, 1998, pp. 231-238

17) Courtesy of Leonard John Otten III of Kestrel Corporation, Albuquerque, NM

18) C. S. L. Chun, L. J. Otten, K. Norsworthy, “MDA (Missile Defense Agency) Technology Program - TechUpdate,” Newsletter Winter 2001,

19) B. Iannotta, “SWARM,” Smithsonian Air & Space, August/September, 2000, pp. 44-49

20) Information and confirmation provided by Ernest Y. Robinson of The Aerospace Corporation, El Segundo, CA

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 (