Minimize STARE

STARE (Space-based Telescopes for the Actionable Refinement of Ephemeris)

STARE is a SSA (Space Situational Awareness) nanosatellite project of NRO, a collaboration led by LLNL (Lawrence Livermore National Laboratory) in Livermore, CA, including NPS (Naval Postgraduate School) in Monterey, CA and TAMU (Texas A&M University) in College Station, TX. LLNL is providing the payload, an optical telescope for capturing satellite streaks, to be integrated by NPS and TAMU into Boeing’s C2B (Colony 2 Bus).

The objectives of the program include: observe objects that are predicted to pass close to a valuable space asset based on conjunction analysis using the AFSPC (Air Force Space Command ) catalog; transmit images and positions of observations to the ground; and refinement of orbital parameters of space objects to reduce uncertainty in position estimation and improve accuracy of conjunction analysis. 1) 2) 3)


Figure 1: STARE mission concept of operations (image credit: NPS)

With the added flexibility and maneuverability of the STARE satellite, conjunction analysis can be refined to a higher level of confidence, ensuring that a possible accident could be prevented. With the additional capability to take pictures of orbital debris, the STARE satellite, and eventual constellation, could be a valuable asset to the Space Surveillance Network and potentially the Joint Space Operations Center (JSpOC) for conjunction analysis. Other tools, such as LLNL’s TESSA (Test-bed Environment for Space Situational Awareness) super-computers have been developed to provide a higher fidelity model of the orbital debris that exist in the LEO (Low Earth Orbit) environment.

STARE is a pathfinder mission for a constellation of satellites that will provide refined orbital information for satellites and debris in Earth orbit. These "space traffic cams" will drastically lower the number of false collision warnings, allowing satellite operators to take action when their assets are in certain danger.



The STARE project uses the Boeing-built C2B (Colony 2 Bus) developed under a contract with NRO (National Reconnaissance Office). A C2B is a nanosatellite (3U form factor CubeSat) containing a radio, batteries, an attitude control unit, solar panels, and other parts with an empty space inside, roughly half the volume for the payload. The C2B configuration employs double-deployed solar panels in cross-axis direction and two single deployed solar panels in axial direction (Figure 4).

In February 2010, NRO contracted with Boeing Phantom Works for as many as 50 triple-unit CubeSats for use in technology demonstrations. The inexpensive satellite platforms will be used for the follow-on to an NRO research program called Colony 2. The Boeing-built Colony 2 platforms will be more powerful than their predecessors and feature better pointing accuracy. The Colony Program Office, nestled within the NRO’s AS&T (Advanced Systems and Technology) division, was established in 2008 to demonstrate new technologies. 4)

In December 2010, NPS received one of only two C2B “preliminary” engineering models (EM) built so far. Three dimensional (3D) polycarbonate models of the C2B and payload are shown in Figure 2. NPS students work with these models to understand the C2B bus modes of operation to include the software protocol and payload integration (Ref. 2). 5) 6)


Figure 2: Early 3D model of C2B with partial payload (image credit: NPS)

Integration of the flight unit at Boeing with NPS participation was completed in late November 2011. The flight unit has the flight software NanoSat GSS version 7.1.0 and the upgraded Nano View. Unlike the previous versions of the NanoSat GSS, the flight version is easier to program and it is more versatile.


Figure 3: Photo of the C2B flight unit in transport structure (image credit: NPS)



Figure 4: Illustration of the deployed STARE nanosatellite in orbit (image credit: NRL, NPS)

To perform the SSA (Space Situational Awareness) mission, STARE will use the same sidereal track mode of operation used on SBV (Space Based Visible) telescope on MSX (Midcourse Space Experiment) mission and on the SBSS (Space Based Space Surveillance) mission of DoD. The CubeSat will point at the stars using them as reference points. The stars appear stationary in the frame while the satellites or orbital debris are streaks as shown in Figure 5. To get useful data, it is essential to maintain 1º sensor pointing accuracy or better. However, the goal is to have 0.31º pointing accuracy. Along with the pointing accuracy, imager pointing stability is critical (Ref. 2).


Figure 5: Schematic view of a sidereal track (image credit: NPS)

The ADCS (Attitude Determination and Control Subsystem) is responsible for ensuring a reasonable level of pointing accuracy. The pointing or orientation time should be less than 24 hours, assuming the ground is sending the next pointing information after exposure, with self-position accuracy, determined by the on-board GPS, having a value of < 50 m. The maximum slew rate of the satellite is 3º/s, which occurs in less than a minute, and the ideal slew rate is 0.25º/s in a time span less than ten minutes. The ideal STARE pointing stability is 0.052º/s smear rate and 0.02º/s jitter rate.

C&DH (Command & Data Handling) subsystem: The C&DH of the spacecraft is supported by the C2B, to include both uplink and downlink paths. It includes hardware and software used in spacecraft control and the interface between the bus and payload. The characteristics of the C&DH module are outlined in Table 1.




C&DH module

Power of 0.65 W (average)

0.33 § (min), 0.97 W (max), 50% duty cycle

Data storage

Effective shared capacity = 2GB

Data is written redundantly to two Flash memory devices (2 x 2 GB) and is shared between bus and payload data


CMD schedule resolution = 1 s
Max No of sequences: 300
Max No of commands: 500
Max No of parameters: 1500
Total parameter size: 36 kB

Several commands can be executed in parallel based on priority. Commands are serviced as quickly as the spacecraft processor can process them.
Sequences, commands, and parameters use shared memory and are limited by any of the max. parameters listed

Telemetry bus

Bus SOH rate: 90 kB/min

Maximum telemetry rate based on typical 1 minute snapshots with additional attitude control data saved at 1 Hz

Table 1: Characteristics of the C&DH subsystem

The C2B is in command of the nanosatellite. It handles communication with the ground, controls the solar panels, distributes power to the various components, and provides attitude control with an on-board star tracker/reaction wheel system. It also contains a non-volatile flash file-system consisting of two SD cards, one being the RAID mirror of the other, that we will use to store our images and telemetry data. Proper functioning of the bus, particularly in the attitude control system, is critical to success of the mission (Ref. 21).



Launch: The STARE-A nanosatellite was launched on Sept. 13, 2012 as a secondary payload on an Atlas-5-411 vehicle of ULA (United Launch Alliance) from VAFB, CA. The primary payload on this flight, referred to as NROL-36, National Reconnaissance Organization Launch) were two NRO/MSD (Mission Support Directorate) classified spacecraft, namely NOSS-36A and NOSS-36B. 7) 8) 9)

Mission title


CubeSat size

CINEMA (CubeSat for Ions, Neutrals, Electrons, & MAgnetic fields)

UCB (University of California, Berkeley)


CSSWE (Colorado Student Space Weather Experiment)

University of Colorado at Boulder


CP5 (Cal Poly CubeSat 5)

California Polytechnic State University


CXBN (Cosmic X-ray Background Nanosatellite)

MSU (Morehead State University)


Table 2: Summary of secondary payloads manifested on ELaNa-6 (Ref. 9)

Next to the above list of NASA sponsored secondary payloads, there are additional secondary payloads sponsored by NRO/MSD as shown in Table 3 (containing all secondary payloads).


Launch Sponsor

CubeSat Name



Mass (kg)


1 & 8




3U, Qty 2






Aerospace Corp., El Segundo, CA

1U, Qty 2


Technology Demo




Aerospace Corp



Technology Demo




USC, Marina Del Ray, CA



Cargo Tracking




LLNL, Livermore, CA



Space Debris Mitigation




Univ of Colorado/NSF



Space Weather


NASA/LSP (Launch Services Program)


Morehead State University and Kentucky Space



Space Weather




CalPoly, San Luis Obispo, CA



Debris Mitigation




NSF/UCB Berkeley



Space Environment

Table 3: NROL-36/OUTSat CubeSat manifest with 8 P-PODs 10) 11)

The launch of all CubeSats is being conducted in a new container structure, referred to as NPSCuL (Naval Postgraduate School CubeSat Launcher). This new dispenser platform was designed and developed by students of NPS (Naval Postgraduate School) in Monterey, CA, to integrate/package P-PODs as secondary payloads.

NRO refers to all 11 secondary (or auxiliary) CubeSat payloads on NROL-36 as the OUTSat (Operationally Unique Technologies Satellite) mission using for the first time the NPSCuL platform as a container structure for the 8 P-PODs (Ref. 11).


Figure 6: Photos of the integrated OUTSat P-PODs in the NPSCuL platform (left) along with the proud NPS students (left), image credit: NRO, NPS

Orbit of all secondary payloads: Elliptical orbit of 770 km x 480 km, inclination = 66º.


Launch: The STARE-B 3U CubeSat was launched on November 20, 2013 (01:15 :00 UTC) from the MARS (Mid-Atlantic Regional Spaceport) on Wallops Island, VA on a Minotaur-1 vehicle of OSC (Orbital Sciences Corporation). The launch was part of the ORS-3 (Operationally Responsive Space-3) enabler launch mission. The primary payload on ORS-3 was STPSat-3. 12) 13) 14)

ORS-3 is ushering in launch and range processes of the future. The ORS-3 mission will demonstrate and validate a new launch vehicle flight safety architecture of the future through the AFSS (Autonomous Flight Safety System) payload, which uses launch vehicle orbital targeting and range safety planning processes to protect public safety from an errant launch vehicle during flight. 15) The outcome of this test is of great interest to the military as well as to NASA. The launch also will be part of the Federal Aviation Administration's (FAA) certification process for the Minotaur rocket. The FAA has licensing authority over American commercial rockets.

Orbit: Near-circular orbit, altitude = 500 km, inclination = 40.5º.

Secondary Payloads: The secondary technology payloads on this flight consist of 26 experiments comprised of free-flying systems and non-separating components (2 experiments). ORS-3 will employ CubeSat wafer adapters, which enable secondary payloads to take advantage of excess lift capacity unavailable to the primary trial. 16) 17)

NASA's LSP (Launch Services Program) ELaNa-4 (Educational Launch of Nanosatellite-4) will launch eight more educational CubeSat missions. The ELaNa-4 CubeSats were originally manifest on the Falcon-9 CRS-2 flight. When NASA received word that the P-PODs on CRS-2 needed to be de-manifested, LSP immediately started looking for other opportunities to launch this complement of CubeSats as soon as possible. 18)


ORS-3 mission sponsor

Spacecraft provider

No of CubeSat Units

ORS-1, ORSES (ORS Enabler Satellite)

ORS (US Army)

Miltec Corporation, Huntsville, AL


ORS-2, ORS Tech 1

ORS Office

JHU/APL, Laurel, MD


ORS-3, ORS Tech 2

ORS Office




SOCOM (Special Operations Command)

LANL (Los Alamos National Laboratory)

1 x 3




1 x 3




1 x 3




1 x 3


STP (Space Test Program)

SMC/XR USAF, Boeing Co.




SMC/XR, USAF, Boeing Co.




NSF (National Science Foundation)



NRO (National Reconnaissance Office)

Lawrence Livermore National Laboratory


Black Knight-1


US Military Academy, West Point, NY




US Naval Academy, Annapolis, MD




Naval Postgraduate School, Monterey, CA




University of Hawaii, Manoa, HI




St Louis University, St. Louis, MO




University of Alabama, Huntsville


SPA¿1 Trailblazer


COSMIAC, University of New Mexico


Vermont Lunar CubeSat


Vermont Technical College, Burlington, VT




University of Florida, Gainsville, FL




University of Louisiana, Lafayette, LA




Drexel University, Philadelphia, PA




Kentucky Space, University of Kentucky




NASA/ARC, Moffett Field, CA


TJ3Sat (CubeSat)


Thomas Jefferson High School, Alexandria, VA


Table 4: ORS-3 manifested CubeSats & Experiments (Ref. 16)



Mission status:

• Dec. 13, 2013: The STARE-B (HORUS) mission is in the commissioning phase at LLNL (Ref. 19).

• April 2013: STARE-A (Re) on orbit. Communication issues prevent operations.- The status of STARE-A remains unchanged in December 2013. 19)

LLNL has conducted a spiral development of the STARE technology through 3 pathfinders. Each pathfinder is a 3U CubeSat and each builds upon technology maturation developed on the previous pathfinder. 20)

1) The first pathfinder, STARE-A, was launched on the NRO-L36 OUTSat (Operationally Unique Technologies Satellite) mission on September 13, 2012 and has experienced communication issues being investigated in the spring/summer of 2013. This first pathfinder has a Cassegrain telescope and an attitude control capability limited to torque coils.

2) The second pathfinder, STARE-B, implementing a full set of reaction wheels for attitude control and a more sensitive imager. STARE-B was launched on November 20, 2013.

3) The third pathfinder, STARE-C, implementing improved reaction wheels and a compact and robust optical telescope, is expected to launch in Q3 of 2014.

The upcoming pathfinders are expected to raise the technology readiness level (TRL) to 7 by exercising refinement in an in-orbit operational environment.



Sensor complement: (Optical Payload, Namuru GPS receiver)

LLNL has had significant involvement in Space Situational Awareness since 2008 when it began implementing a large scale computer simulation called TESSA (Testbed Environment for Space Situational Awareness). TESSA is part of a collaboration between LLNL (Lawrence Livermore National Laboratory), SNL (Sandia National Laboratories), and the AFRL (Air Force Research Laboratory), its primary aim being to improve performance analysis of the SSN (Space Surveillance Network). Although TESSA is not yet a finished product, it has already produced several important results. One of the most important discoveries was that adding an auxiliary set of sensors to the SSN can drastically reduce the number of close conjunction predictions. 21)

It is not immediately clear as to what auxiliary network would provide the most improvement at the smallest cost, but one obvious choice is a set of small SSN dedicated satellites whose sole purpose is to observe other satellites and orbital debris. Placing them in orbit bypasses the problem of image degradation due to atmospheric turbulence and allows for the use of small aperture optics because of the close proximity the satellites will have to their targets. Plus, with the cheap cost of equipping and launching a 3U Cubesat, a large constellation of these sensors can eventually be deployed without significant financial burden.

The main goal of the STARE mission is to demonstrate usefulness of space based sensing for refining orbital parameters of an orbiting object. But exactly what level of refinement do we consider useful? As a concrete example, consider the Iridium constellation of satellites in LEO. With current SPG4 models, the positional error for these objects over one day is typically 1000 m. At this level of uncertainty the Iridium operators receive about 10 warnings per day of close approaches within 1 km. If, instead, the position were known to within 100 m, this number would drop to 1 possible conjunction in 10 days. Furthermore, the certainty with which we can say a collision will not occur based on the uncertainty ellipsoids of the objects will be reduced by four orders of magnitude.

Note: SPG4 (Simplified General Perturbations No. 4) propagator. SPG4 cannot accurately represent most orbits. The positional error is typically 1000 m and can be much worse for orbits with high eccentricity. And this assumes that the TLE's best represent the real orbits in a least squares sense over one day. SGP4 only has good positional accuracy for near circular low earth orbits having near zero inclination. The results of the STARE orbit refinements will instead be accurate force model orbits.


Stretch goal

Useful level

Orbital Accuracy After Refinement

≤ 50 m
≥ 1 day ahead

< 100 m
≥ 1 day ahead

Characteristics of Objects

Tangential velocity

< 1000 km
< 0.1 m2
< 10 km/s

< 100 m
< 1 m2
< 3 km/s

Table 5: A list of the goals for the STARE mission. The second row refers to the minimum limit at which accuracy can be considered valid.

One of the primary constraints on the STARE mission is the size of the satellites themselves. The 3U Cubesat limits the diameter of the primary optic to < 10 cm, which, along with the characteristics of our baseline Cypress IBIS-5B CMOS sensor, limits the maximum distance to the targets the project is trying to image. Based on SNR calculations with these considerations in place, De Vries has simulated an orbital platform to maximize the number of observation opportunities, with the following criteria constituting a valid observation (i.e. one capable of reducing the size of the uncertainty ellipsoid of the target):

• A maximum separation smaller than 100 km (due to sensor choice)

• A relative tangential velocity less than 3 km/s (due to sensor choice)

• A solar separation angle larger than 30º (corresponding to a solar exclusion angle of 30º)

• An Earth exclusion angle of 85º

• A lunar exclusion angle of 1º.

These criteria, along with considerations of downlink opportunities, solar panel orientation and attitude control, drag-limited orbital lifetime, and GPS signal coverage, limit the number of useful orbital regimes for the STARE satellites.

Examining the close conjunctions from the simulations occurring over a one-week period with the cataloged objects in LEO shows that a 700 km polar orbit with an inclination of about 90º is optimal for the purposes of the project. In particular, a sun-synchronous orbit of 98º simplifies satellite attitude control with respect to solar panel power generation. The baseline for the STARE satellites was thus chosen to be a 700 km, sun-synchronous orbit with an inclination of 98º.

Observing strategy: The STARE mission only plans on having one dedicated ground station for communication with the STARE satellites. This station, located at NPS, will allow for about 2 minutes of data transfer per day at 9600 baud. Hence, the project is limited to downloading about 1Mb of data per day, which is close to the size of one 1280 x1024 pixel image.

Fortunately, the vast majority of the 1,310,720 pixels in an image will contain only detector noise and sky background, so they are of no use to the project. The information the project is actually after is the following:

4) Precise position and time of satellite at time of observation: This information is contained in the GPS receiver logs that are recorded simultaneously with the image capture. Each GPS log is approximately 200-300 bytes.

5) Stellar Positions (in detector coordinates): The positions of the stars will give the project a very accurate pointing of the satellite once matched up to cataloged positions. The project will record the location and flux of the 100 brightest stars in the image.

6) Track Endpoint Positions (in detector coordinates): Along with the timing and angular information from the two items above, the track endpoints tell the project exactly where the satellite was at the start and end of the observation (in the transverse plane).

While the project will have the capability to download a full, raw image from the payload to the ground (a typical image averages 600-700 kB in size once compressed) for diagnostic and calibration purposes, this is the information that will be routinely received on the ground. The GPS data will be logged from the on-board receiver and the star and track data will be extracted from the images by using the algorithm which will run in the PXA 270 payload microprocessor.

Of course, this all relies on the assumption that the images contain a track and a suitable number of stars to yield an astrometric solution. To ensure this is the case, the satellite will be commanded to point toward a given target (when it is passing through a field with an ample number of bright stars) and begin acquiring images at the calculated time of conjunction. In a typical observing sequence, 10 consecutive one-second exposures will be taken along with their corresponding time-stamps. The 10 image allotment should guarantee that one or two images contain the track even with the 1000 m uncertainty of its TLE (Two Line Elements).


Figure 7: The 3U CubeSat contains the LLNL developed STARE optical payload (image credit: LLNL, Ref. 3)


Optical payload:

The optical payload was designed and developed at LLNL. The instrument is filling up approximately 1.5U of the C2B volume. The optical tracking payload is designed to acquire images of small orbiting objects, pre-process the data relevant to the target’s orbit and pass the processed data to the ground station via the communication system (Ref. 21).

The optical payload is comprised of a modified reflective Cassegrain telescope and a CMOS imager at its focus. The telescope consists of two reflective conics with one corrector lens near the detector intended to reduce the aberrations at the field edges. With a 225 mm focal length and an aperture of 85 mm, the system will yield a resolution of about 29 µrad/pixel, which corresponds to 6.100/pixel. At a range of 100 km, this is about 2.9 m/pixel. The entire field size is 2.08º x 1.67º. A baffling system, not shown in Figure 8, will reduce the stray light. 22)


Figure 8: Illustration of the Cassegrain telescope (image credit: LLNL)

A detailed view of the payload shows the two-mirror telescope design, the location of the imager board, the GPS (Global Positioning System) receiver board, GPS antenna and the interface board. The interface board was developed at NPS to connect the STARE payload to the Colony 2 Bus. The 1.5U payload was designed to fit into the space allotted shown in Figure 9.

A wide field of view is obviously beneficial for the intended application since it: 1) increases the chance that to capture the entire streak in one exposure and 2) increases the maximum velocity the target can have relative to the spacecraft. To obtain a wide field with minimal aberrations in the small 10 cm3 space offered in the Cubesat payload is a challenge, though, especially since a refractive design is prone to severe weathering in LEO. This makes the STARE telescope design rather unique.

Telescope: At the center of the primary mirror, carved from the same piece of glass, is a lens that corrects for the aberrations at the edge of the field. The telescope delivers an approximately f/2.5 beam, and with our 8.6 mm x 6.9 mm imager, this equates to a field of view of about 2.08º x 1.67º.

A further challenge of the optical system is that there is no focusing mechanism. Thermal expansion and contraction in the space environment are thus of great concern. The telescope is designed to have a depth of focus of 10 µm and an Invar support structure will be used to provide stiffness under changing temperatures. Preliminary thermal calculations show that the focus will be maintained over the -20 to +60ºC range expected in the STARE orbit.

Imager: After collecting the star and track light, the telescope will focus it onto a Cypress IBIS5-B-1300 CMOS imager. The detector, which has a 1280 x1024 format with 6.7 µm pixels, is mainly intended for video rate imaging. The project selected the IBIS5-B-1300 detector , because Boeing is able to provide it in a fully integrated system (which includes the PXA 270 microprocessor) that will facilitate communications with the C2B (Colony-2 Bus) and save the team a great deal of development time and expenses, allowing to finish these pathfinder satellites in time for launch.


Figure 9: The Colony 2 bus (left) communicates with the optical payload (right) via an RS-442 connection (image credit: LLNL)

The LLNL imaging system has a processing board, camera and optics comprised of a two-mirror telescope with corrector lens. The payload is used to image targets less than or equal to 300 km in distance traveling at a speed of ≤ 3 km/s.

Number of pixels

1280 x 1024

Focal length, f/number

225 mm, 2.65

FOV (Field of View)

2.08º x 1.67º

Pixel size


Readout resolution

8 bit

Exposure time

1 s

Aperture diameter

85 mm

Instrument size

< 9.75 cm x 9.75 cm x 15 cm

Instrument mass

< 1.83 kg

Output data rate

< 50 kbit/s

Table 6: Specification of the optical payload

Technical goals: The initial goal of the project is to create two pathfinder CubeSats to test the feasibility of using a 3U CubeSat with an optical payload to refine conjunction analysis in support of SSA. In the event the CubeSats are successful in their mission, more work will be done in the area to further the use of CubeSats for space based space surveillance. This system is not intended to be heavily requirement driven, but to demonstrate the usefulness of space based sensing for refining orbital parameters.

SMILE (STARE Mission Hardware in the Loop Environment) software will be used for operations on the ground with STARE. The payload data from STARE will feed into SMILE for improved ephemeris data for conjunction analysis.

For the STARE optical imager on the payload to take an image, the nanosatellite’s view must not be obstructed by the Earth or moon. It should be in the Earth’s shadow or umbra, while the target must be in the full sunlight. The sun constraint shown in Figure 10 was one of several constraints modeled in the STK simulation of the STARE orbit.


Figure 10: Depiction of umbra and full sunlight (image credit: NPS, Ref. 5)


Namuru V3.2 GPS receiver:

The Namuru V3-2 is an L1 FPGA based GNSS receiver designed for integration into a Colony 2 CubeSat. The new Namuru V3.2 GPS receiver hardware design has been in development by Kevin Parkinson of General Dynamics Corporation Ltd., Albany, New Zealand. This development has followed the design philosophy of using commercially available components rather than fully space- qualified components in order to constrain the cost of the device. 23)


Figure 11: The PCB (Printed Circuit Board) of the NamuruV3.2 GPS receiver (image credit: University of New South Wales)

Some of the features of the Namuru V3.2 are: 24)

• The mechanical dimensions of 82 × 82 mm are tailored to the Colony 2 bus (C2B).

• The electrical interface includes the RS422 connections required by the C2B, as well as the particular I/O signals required for the mission.

• The design employs a capable processor in the form of an Actel SmartFusion A2F500 system on a chip that includes a hard wired ARM Cortex M3 processor, 64 kB of internal SRAM, 512 kB of internal Flash, a Flash based FPGA fabric with 11,520 FPGA tiles, and other peripherals.

• The design includes an Actel ProASIC3E Flash based FPGA with 38,400 tiles for the inclusion of large digital hardware designs.

• A serial Flash memory has been included for non-volatile storage and the storage of FPGA and firmware images required for over-the-air reprogramming.

• A Zarlink GP2015 L1 RF front end has been used because it is a well-understood front end, even though the device is no longer in production. The RF portion of the receiver includes an RF shield to reduce electromagnetic susceptibility.

• A 10 MHz voltage controlled temperature compensated crystal oscillator (VC-TCXO) has been used in order to discipline the 10 MHz local clock and eliminate saw-tooth timing errors.

• 1 MB of fast external SRAM has been included in order to ensure that sufficient SRAM is available for firmware and data variables.

• The design includes latch up detection on the power supply that causes the power to be cycled if the current drain exceeds a set threshold.

• Mechanical stresses caused by thermal cycling have been considered in the design of the printed circuit board through the use of solid (filled) vias and the use of transitional tin-lead solder alloys.

• A super-cap allows the RTC included in the SmartFusion A2F500 to be run for at least 24 hours.

• An SMA antenna connection has been employed to allow the antenna to be robustly connected to the receiver.



Ground Segment:

The MC3 (Mobile CubeSat Command and Control) system provides ground support for the Colony -2 program. The NRO commitment to the Colony -2 program is manifested in the purchase of twenty C2Bs (Colony-2 satellite Buses) and six MC3 ground stations. The NRO contract with Boeing is written to permit procurement of an additional 30 C2Bs. 25) 26) 27)

The original MC3 is the ground station that supported Colony-1 missions and when updated will support Colony-2 missions. The MC3 ground station uses COTS (Commercial Off-the-Shelf) equipment tied together with GOTS (Government Off-the-Shelf) software to provide a low-cost, easy to operate ground station specifically designed to support the C2B nanosatellites.

The general MC3 architecture consists of: 28)

• The MC3 ground station network to connect multiple government and university nodes together via a Wide Area Network.

• MC3 utilizes CGA (Common Ground Architecture) software to command, control, schedule, and monitor spacecraft operations.

- CGA has been in operation since 1982 on a wide variety of NRL satellite programs

- CGA runs on CentOS (Linux)

- CGA is a Government-owned open architecture software which can be coded for any aspect of spacecraft operations

- Can be run autonomously; user inputs schedules - CGA will automatically assign resources on the network to track and pass data/commands when the satellites are over a node.


The MC3 CGA allows the satellite communications systems to be tested on the same equipment it will use to operate, using the same commands, radios, and antennas. Once deployed, the world-wide network of MC3s will provide coverage at geographically advantageous areas to universities, service academies, and research facilities who operate Colony-2 satellites. The MC3 comes in one computer rack and uses two portable antennas for easy installation at any participating facility.

NPS involvement with the Colony program started with the introduction of the Colony-2 satellite bus and MC3 ground station. NPS was approached by the NRO CubeSat Program Office to build, test, and host an MC3 ground station and act as the master, coordinating node for all MC3s installed at universities.


Figure 12: Overview of the initial MC3 architecture (image credit: NPS)

• NPS (Naval Postgraduate School), Monterey, CA

• USU (Utah State University), Logan, UT

• UH (University of Hawaii) at Manoa, HI

• AFIT (Air Force Institute of Technology), located on Wright-Patterson AFB, Ohio.

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14) Roz Brown, “Ball Aerospace's STPSat-3 to Fly Solar TIM Instrument for NOAA,” BATC, July 19, 2012, URL:

15) Michael P. Kleiman, “ORS Office organizing three new programs,” AFMC (Air Force Materiel Command), Aug. 30, 2012, URL:

16) Peter Wegner, “ORS Program Status,” Reinventing Space Conference, El Segundo, CA, USA, May 7-10, 2012, URL:

17) Joe Maly, “ESPA CubeSat Accommodations and Qualification of 6U Mount (SUM),” 10th Annual CubeSat Developer’s Workshop, Cal Poly State University, San Luis Obispo, CA, USA, April 24-25, 2013, URL:

18) Garrett Lee Skrobot, Roland Coelho, “ELaNa – Educational Launch of Nanosatellite Providing Routine RideShare Opportunities,” Proceedings of the 26th Annual AIAA/USU Conference on Small Satellites, Logan, Utah, USA, August 13-16, 2012, paper: SSC12-V-5

19) Information provided by Steve Wampler of LLNL (Lawrence Livermore National Laboratory) in Livermore, CA, USA.

20) Vincent Riot, Willem de Vries, Lance Simms, Brian Bauman, Darrell Carter, Don Phillion, Scot Olivier, “The Space-based Telescopes for Actionable Refinement of Ephemeris (STARE) mission,” Proceedings of the 27th AIAA/USU Conference, Small Satellite Constellations, Logan, Utah, USA, Aug. 10-15, 2013, paper: SSC13-XI-11, URL:, URL of presentation:

21) Lance M. Simms, Vincent Riot, Wilhem De Vries, Scot S. Olivier, Alex Pertica, Brian J. Bauman, Don Phillion, Sergei Nikolaev, “Optical payload for the STARE mission,” LLNL-CONF-474234, Lawrence Livermore National Laboratory, SPIE Defense and Security, Orlando, FL, USA,March 25-29, 2011, URL:

22) Lance M. Simms, Wilhem De Vries, Vincent Riot, Scot S. Olivier, Alex Pertica, Brian J. Bauman, Don Phillion, Sergei Nikolaev, “Space-based telescopes for actionable refinement of ephemeris pathfinder mission,” Optical Engineering, Volume 51, Issue 1, Jan. 19, 2012, URL:

23) Éamonn Glennon, Kevin Parkinson, Peter Mumford, Nagaraj Shivaramaiah Yong Li, Rui Li, Yuanyuan Jiao, “A GPS Receiver Designed for Cubesat Operations,” 2011, URL:

24) K. J. Parkinson, P. J. Mumford, E. P. Glennon, N. C. Shivaramaiah, A. G. Dempster, and C. Rizos, "A Low cost Namuru V3 recever for Spacecraft operations," International Global Navigation Satellite Systems Society IGNSS Symposium, Nov. 2011, University of New South Wales, Sydney, Australia

25) Gregory C. Morrison, “Mobile CubeSat Command and Control: Assembly and Lessons Learned,” NPS Thesis, September 2011, URL:

26) Robert C. Griffith, “Mobile CubeSat Command and Control,” NPS Thesis, Sept. 2011, URL:

27) Giovanni Minelli, Philip Ibbitson, Aaron Felt, Ernesto Yzquierdo, James Horning, David Rigmaiden, James Newman, “Mobile CubeSat Command & Control (MC3) Ground Stations,” 2012 Summer CubeSat Developers’ Workshop, Logan, Utah, USA, Aug. 11-12, 2012, URL:

28) Jim Newman, “Mobile CubeSat Command and Control,” 9th Annual Spring CubeSat Developer's Workshop, Cal Poly State University, San Luis Obispo, CA, USA, April 18-20, 2012, URL:

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