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Satellite Missions Catalogue

CeREs (Compact Radiation Belt Explorer)

Sep 3, 2018




Quick facts


Mission typeNon-EO
Launch date16 Dec 2018

CeREs (Compact Radiation Belt Explorer)

Spacecraft    Launch   Sensor Complement   Ground Support   References

CeREs is the first fully NASA-SMD (Science Mission Directorate) funded CubeSat and contributes to NASA's Heliophysics program by: (a) advancing the understanding of the radiation belt electron energization and loss processes by making high-cadence, high-resolution measurements of the energy spectra of electrons over a broad energy range, and (b) will flight-validate an innovative, compact, low-mass, low-power instrument that has many future applications. 1)

Through the extensive involvement of graduate students, CeREs will provide valuable training for the next generation of experimental space physicists. CeREs measurements of radiation belt electrons at LEO will extend and complement NASA's flagship mission, the Van Allen Probes. A photo of the CeREs spacecraft being readied for delivery to the launch provider is shown in Figure 1. The figure also shows the Tyvak rail pod deployer that will deploy the CeREs spacecraft.

The spacecraft bus and the payload were developed, tested, and integrated at NASA/GSFC (Goddard Space Flight Center) with contributions from the Co-Is at SwRI (Southwest Research Institute). The project also involved graduate students from the Catholic University of America in Washington DC and the University of Texas, San Antonio.

Figure 1: CeREs spacecraft (left) with the Tyvak Rail Pod deployer (right), image credit: NASA
Figure 1: CeREs spacecraft (left) with the Tyvak Rail Pod deployer (right), image credit: NASA

As mentioned above, the science objectives addressed by CeREs encompass electron energization and loss in the Earth's radiation belts and solar electron acceleration and transport, with the former being the primary science goal of the mission. In addition, CeREs measurements of energetic protons over the polar-regions provide information about how these hazardous particles access geospace during geomagnetically disturbed times, an important aspect of space weather.

Radiation Belt Electron Dynamics

The electron populations in the Earth's outer Van Allen belts are highly dynamic due to energization and loss processes that occur over time scales ranging from minutes to years. It is currently understood that electron energization results from local in-situ wave particle interactions, radial transport, or a combination of both. It is essential to understand the contribution of both electron loss and energization in order to ascertain the net flux levels in the radiation belts.

While there are several processes that deplete electrons in the outer belt, loss due to microbursts has recently gained prominence since they may potentially empty the belts on time scales of days. Microbursts are short-lived (<1 s) bursts of electrons scattered into the loss cone by interactions with plasma waves in the magnetosphere. Figure 2 shows electron microbursts observed over a radiation belt pass by the HILT (Heavy Ion Large area proportional counter Telescope) sensor onboard the SAMPEX spacecraft. Microbursts are seen as "spikes" over the smooth bell-shaped radiation belt electrons.

Figure 2: HILT observations of microbursts of electrons > 1 MeV (image credit: NASA)
Figure 2: HILT observations of microbursts of electrons > 1 MeV (image credit: NASA)

The MERiT (Miniaturized Electron and pRoton Telescope) instrument will characterize microbursts with much higher time resolution than prior measurements, as well as provide spectral information in multiple differential channels to better understand microburst properties. In order to distinguish between energization due to radial transport vs. in-situ energization, it is necessary to examine the radial profiles of PSD (Phase Space Density). PSD radial profiles differ for internal, i.e., wave-particle energization, as opposed to those due to radial transport. A recent study using PSD radial profiles has shown that it is possible to identify both processes even though they may occur during a single enhancement event. This is illustrated in Figure 3, which shows both monotonically increasing and peaked radial profiles of PSD for electrons of different energies during an energization event driven by simultaneous high-speed solar wind stream (HSS) and CME (Coronal Mass Ejection). Calculation of PSD requires knowledge of both the model-dependent global magnetospheric field and accurate measurements of particle spectra. MERiT measurements with differential energy coverage will reduce uncertainties arising from less precise spectral measurements.

Figure 3: Van Allen Probes observations of radial profiles of PSD during an event driven simultaneously by HSS and CME (image credit: NASA) 2)
Figure 3: Van Allen Probes observations of radial profiles of PSD during an event driven simultaneously by HSS and CME (image credit: NASA) 2)

Solar Electron Energization and Transport

As CeREs traverses the polar-regions, it encounters open field lines, i.e., magnetic field lines that connect to the Sun. Charged particles emanating from various solar sources such as flares, CME shock accelerated interplanetary regions, reconnection regions higher in the corona, etc. are all observable from low earth orbit over the polar caps. Open questions remain concerning regions of solar electron acceleration at lower energies (~tens of keV). This is due to lack of high quality measurements at these energies, which will be well covered by MERiT measurements. Figure 4 shows solar electron spectra measured by ACE and SAMPEX, which highlight the lack of measurement below 30 keV, and where MERiT will fill the gaps.

Figure 4: Solar electron spectra showing measurements by ACE and SAMPEX (image credit: NASA)
Figure 4: Solar electron spectra showing measurements by ACE and SAMPEX (image credit: NASA)

Solar Proton Access to Geospace

The geomagnetic field prevents low rigidity (momentum per unit charge) charged particles from reaching low latitudes. The minimum value of rigidity required to reach a point on the Earth is termed geomagnetic cutoff rigidity. This cutoff rigidity, which depends upon the geomagnetic field and direction of arrival, varies during geomagnetic storms when the field is distorted. Proton cutoff variability as measured by sensors onboard SAMPEX during a geomagnetic storm is shown in Figure 5 for 19.0-26.0 (blue), and 22.0-60.0 MeV (red) protons.

The figure also shows the Dst (Disturbance storm time) index, a measure of the distortion of the geomagnetic field. It is evident the cutoff location follows the variation in the geomagnetic field, with protons reaching lower latitude as the strength of the geomagnetic field disturbance increases. This is important from a space weather perspective as at times these energetic particles can reach space station orbits.

Figure 5: Solar proton access to geospace during a geomagnetic storm. The figure shows 19-26 MeV (blue) and 22-60 MeV (red) protons measured by the PET sensor onboard SAMPEX (image credit: NASA)
Figure 5: Solar proton access to geospace during a geomagnetic storm. The figure shows 19-26 MeV (blue) and 22-60 MeV (red) protons measured by the PET sensor onboard SAMPEX (image credit: NASA)




The CeREs CubeSat has 3U form factor, with the payload and the bus occupying almost 1.5U each. Power is supplied by dual deployable solar panels provided by Clyde Space, Glasgow, Scotland and the bus, XB1, provided by Blue Canyon technologies, Boulder, CO. In order to achieve CeREs science, the spacecraft bus was required to:

• Store 854 Mbit of science and 100+ Mbit housekeeping data per day

• Support UHF-band downlink data rate of 2.2 Mbit/s, and uplink date rate of 9.6 kbit/s

• Provide > 9.2 W orbit-average power

• Launch mass < 4.58 kg

• Orbital debris compliant with NPR (NASA Procedural Requirements) 8715.6

• Point MERiT to local zenith to better than 200

• Collect >40 kbit/s of data when >60 latitude

• Provide 148 Full cone clear FOV for MERiT.

Figure 6: CeREs cut-away view illustrating the main components and their functionality (image credit: NASA)
Figure 6: CeREs cut-away view illustrating the main components and their functionality (image credit: NASA)

The 3U structure is made of aluminum with Teflon rails for the TyVak Rail-Pod deployer. The spacecraft has a passive thermal design with a battery heater for safehold. A high level block diagram of CeREs is shown in Figure 7.

Figure 7: CeREs high-level block diagram (image credit: NASA)
Figure 7: CeREs high-level block diagram (image credit: NASA)

XB1 Bus of BCT (Blue Canyon Technologies)

The XB1 is a complete solution, providing bus functionality for GN&C, EPS, Thermal, C&DH, RF Comm, and SSR. A precision star tracker and two Sun pointers (coarse and fine) provide attitude control with pointing accuracy that exceed mission requirements. XB1 also provides for real time commands as well as command macros. The XB1 is powered on and able to receive RF communications during all points of the mission. Should an upset occur, radiation or otherwise, all spacecraft components reset to safe mode.

Two modes exist for the GN&C – sun point mode and fine reference point mode. The sun point mode acts as a safe mode while fine reference pointing being used for all other mission modes and is command able from the ground. For CeREs, the sun point mode has the spacecraft X-axis toward the sun, which is perpendicular to the solar panels, the Z-axis along the length of 3U, and the Y-axis completing the coordinate system. The spacecraft will automatically enter this mode at boot, anytime battery voltage drops below a threshold, or as a result of any fault. Onboard ADCS hardware powered during this mode includes the reaction wheels, star tracker, torque rods, three-axis magnetometer, inertial measurement unit, coarse sun sensors, and the battery heater.

The Fine Reference Point mode is entered when the following requirements are met: (a) valid attitude, obtained by star tracker, propagated by IMU, (b) valid time, obtained by GPS, propagated by onboard oscillator, and, (c) valid orbit / ephemeris – obtained by GPS. The GPS is set up to be activated after initial checkout activities are completed and will run duty cycled to reduce the overall power consumption onboard. The star tracker is always enabled in the XB1 system. Should either the GPS or star tracker fail, the telemetry generated may be uploaded to the spacecraft via a ground command. A high-level block diagram of the XB1 bus is shown in Figure 8.

Figure 8: XB1 high-level block diagram (image credit: BCT)
Figure 8: XB1 high-level block diagram (image credit: BCT)

EPS (Electrical Power Subsystem): The XB1 electrical power system receives power input from the deployable COTS solar arrays purchased from ClydeSpace and stores charge in an internal lithium ion battery pack made of up three series cells with a capacity of 2.6 Ah and a maximum voltage of 12.6 V. Ten switched power rails provide positive 3.3, 5, 7, and 12 V.

RF communications subsystem: CeREs uses the Cadet UHF radio, which is internal to the XB1. The XB1 maintains the 3.3V power to the radio to enable the receipt of RF commands and only enables additional power rails when commanded. The command to transmit enables a beacon, single packet of high FIFO (First-In, First-Out) data, to be sent at a set rate. In order to downlink stored data from the low FIFO, this beacon must be active. As the Cadet radio is directly connected to the XB1 bus, all communications to the payload flows through both components. Similarly, all data generated by MERiT pass through the XB1 to be stored in the four-gigabyte FIFO queue onboard the Cadet.

FSW (Flight Software): The FSW has a modular structure and comprises several core modules. The CFE (Core Flight Executive) is a set of mission independent reusable core flight software services and operating environment that provides standardized APIs (Application Programmer Interfaces). The applications can be added or removed at run-time and contains platform and mission configuration parameters. These core components include:

- ES (Executive Services), which manage the software system;

- SB (Software Bus), which provides publish/ subscribe software bus messaging interface;

- TIME (Time Services), which provides spacecraft time;

- EVS (Event Services), which provides interface for sending, filtering, and logging event messages; and

- TBL (Table Services), which provides interface to manage table images.

FSW also has core flight software services (CFS) and all CFS applications are reused with minor project-specific configuration changes. These services cover all the major aspects such as file uplink downlink, house keeping, commanding, scheduling, etc. Mission specific CFS, i.e., custom applications are provided for instrument component management [APD (Avalanche Photodiodes), SSD (Solid State Detectors), FEE (Front End Electronics)] and science telemetry output.

Figure 9: The CeREs CubeSat is loaded inside its deployment canister at Rocket Lab's facility in Huntington Beach, California (image credit: NASA)
Figure 9: The CeREs CubeSat is loaded inside its deployment canister at Rocket Lab's facility in Huntington Beach, California (image credit: NASA)



The CeREs CubeSat was launched on the Electron vehicle of Rocket Lab on 16 December 2018 at 06:33 UTC (19:33 NZDT) from Rocket Lab Launch Complex 1 on New Zealand's Māhia Peninsula with the ELaNa-19 payloads and others. 3) 4)

NASA is providing the CeREs launch via the ElaNa-CSLI (Educational Launch of Nanosatellites-CubeSat Launch Initiative) program. 5)

Orbit: Circular high-inclination (85º) orbit at an altitude of 500 km.


Payload Complement

This mission includes 10 ELaNa-19 (Educational Launch of Nanosatellites-19) payloads, selected by NASA's CubeSat Launch Initiative. The initiative is designed to enhance technology development and student involvement. These payloads will provide information and demonstrations in the following areas:

• CeREs (Compact Radiation belt Explorer), a 3U CubeSat of NASA. High energy particle measurement in Earth's radiation belt.

• STF-1 (Simulation-to-Flight-1), a 3U CubeSat (4 kg) of WVU (West Virginia University). The objective is to demonstrate how established simulation technologies may be adapted for flexible and effective use on missions using the CubeSat Platform.

• AlBus (Advanced Electrical Bus), a 3U CubeSat of NASA/GRC to demonstrate power technology for high density CubeSats.

• CHOMPTT (CubeSat Handling Of Multisystem Precision Time Transfer), a 3U CubeSat of UFL (University of Florida). CHOMPTT is equipped with atomic clocks to be synchronized with a ground clock via laser pulses.

• CubeSail, a mission of the University of Illinois at Urbana-Champaign. A low-cost demonstration of the UltraSail solar sailing concept, using two near-identical 1.5U CubeSat satellites to deploy a 260 m-long, 20 m2 reflecting film.

• NMTSat (New Mexico Tech Satellite), a 3U CubeSat developed by the New Mexico Institute of Mining and Technology with the goal to monitor space weather in low Earth orbit and correlate this data with results from structural and electrical health monitoring systems.

• RSat-P (Repair Satellite-Prototype), a 3U CubeSat of the USNA (US Naval Academy ) in Annapolis Maryland to demonstrate capabilities for in-orbit repair systems (manipulation of robotic arms).

• ISX (Ionospheric Scintillation Explorer), a 3U CubeSat of NASA and CalPoly to investigate the physics of naturally occurring Equatorial Spread F ionospheric irregularities by deploying a passive ultra-high frequency radio scintillation receiver.

• Shields-1, a 3U CubeSat of NASA/LaRC, a technology demonstration of environmentally durable space hardware to increase the technology readiness level of new commercial hardware through performance validation in the relevant space environment.

• Da Vinci, a 3U CubeSat of the North Idaho STEM Charter Academy to teach students about radio waves, aeronautical engineering, space propulsion, and geography by sending a communication signal to schools around the world.

In addition to the 10 CubeSats to be launched through NASA's ELaNa program, there are three more nanosatellites set for liftoff on top of the Electron rocket in New Zealand. NASA also provided a launch opportunity for:

• AeroCube 11 consists of two nearly identical 3U CubeSats developed by the Aerospace Corp. in El Segundo, California. The AeroCube 11 mission's two CubeSats, named TOMSat EagleScout and TOMSat R3, will test miniaturized imagers. One of the CubeSats carries a pushbroom imager to collect vegetation data for comparison to the much larger OLI (Operational Land Imager) aboard the Landsat-8 satellite, and the other TOMSat CubeSat has a focal plane array on-board to take pictures of Earth, the moon and stars. Both satellites feature a laser communication downlink.

• SHFT (Space-based High Frequency Testbed), a 3U CubeSat (5 kg) mission of DARPA, developed by NASA/JPL. The objective is to study variations in the plasma density of the ionosphere by collecting high-frequency radio signals, including those from natural galactic background emissions, from Jupiter, and from transmitters on Earth.

Rocket Lab has christened the mission "This One's for Pickering" in honor of the New Zealand-born scientist William Pickering, who was director of the Jet Propulsion Laboratory in Pasadena, California, for 22 years until his retirement in 1976.



Sensor Complement

MERiT (Miniaturized Electron Proton Telescope)

The MERiT instrument combines a solid state particle telescope and APDs (Avalanche Photodiodes) to extend the energy range of measured charged particles to much lower values (~ few keV) than is usually possible using SSDs (Solid State Detectors) alone. There are 8 electron and 10 proton differential channels for the SSD stack and 16 APD channels. The data from the APDs are of two types: differential channel rates for electrons, protons, and mixed channels ranging in variable cadence; integral proton rate at variable cadence (Ref. 1). 6) 7)

The data from the SSD stack are of three types: singles rates from each of the eight SSDs, differential channel rates for electrons and protons at variable cadence, and PHAs from all eight SSDs, including differential channel IDs, for a variable number of events. The instrument operates two modes: high time resolution or microburst mode (MB) and low time resolution or SEP (Solar Energetic Particle) mode and is fully configurable and can be modified on flight.

All the key constants that determine thresholds and differential channels can be uploaded via ground command, as are the instrument mode boundary settings. The nominal boundaries correspond approximately to the radiation belts and polar caps for the MB and SEP modes, respectively. The entire stack is enclosed in tungsten-aluminum (W-Al) shielding to prevent contamination from side penetrating particles, and a requirement of contiguous hits in the detector stack within 250 ns (changeable via ground command) ensures minimal background due to chance coincidences. Onboard processing for MERiT is done by CSP (CHREC Space Processor), a radiation tolerant Zynq FPGA-ARM processor, developed by the CHREC (Center for High-Performance Reconfigurable Computing) consortium. 8)

Figure 10 shows the fully-assembled sensor stack surrounded by the shielding and the front end electronics being tested in the laboratory. The resource parameters and the energy ranges of electrons and protons measured by MERiT are listed in Table 1.



Energy range

Measurement and Functional Parameters


~5 keV to ~ 8 MeV


~200 keV to ~ 100 MeV

Energy Resolution

ΔE/E < 30%

Time resolution

4 ms to 1 s

Geometry factor

~ 20 cm2 sr (> 800 keV)
0.05 cm2 sr (~5 to ~450 keV)

Resource Parameters


1.13 kg


0.3 W orbit average

Data rate

~ 42 kbit/s

Table 1: MERiT resource and parameters
Figure 10: Photos showing the MERiT sensor stack (left) and FEE (Front End Electronics) board (right), image credit: NASA)
Figure 10: Photos showing the MERiT sensor stack (left) and FEE (Front End Electronics) board (right), image credit: NASA)



Spacecraft Integration and Testing

CeREs was integrated at GSFC's Heliophysics Energetic Particles Laboratory after component-level testing. All components, including the solar panels, science payload, and spacecraft bus were tested for functionality and performance verified. Some components, e.g., SSDs and the APDs, were flight qualified to NASA specs by the vendor. After the spacecraft was fully integrated, random vibration, thermal vacuum, and balance tests were performed, with the spacecraft meeting or exceeding the launch provider's requirements. Acoustic testing was found to be unnecessary based on the results of the vibration test results. A Rail-Pod shock analysis showed compliance with launch requirements. Venting analysis confirmed that both the payload and XB1 were compliant and capable of surviving launch ascent pressure profile. NASA provides the CeREs launch via the ElaNa-CSLI program on the Electron launch vehicle built by Rocket Labs.

Random vibration test: The random vibration test and low-level signature sweeps tests were conducted for each spacecraft axis, at GEVS -acceptance levels, to demonstrate compliance with the Launch Vehicle requirements. The test sequence for each axis started and ended with a low-level signature sweep, interleaved by random vibration. The results of the test were that all requirements were met with no significant structural changes or failures, and full functionality of the spacecraft was maintained after completion of the test.

Thermal vacuum and Thermal Balance test: Although the launch provider only required a bake out, CeREs underwent a four-cycle thermal vacuum and thermal balance test, including a bake out on the first cycle. Bake out was held at 60ºC for 6 hours, per requirement. The extended duration profile was selected based on survival temperatures of key components. The test setup consisted of thermocouples placed on the outside of the spacecraft, heater panel facing the -X face of CeREs, and the shroud set to space temperature. Thermal balance testing proved that the design is valid. No component violated its temperature limits through the testing. Overall the balance temperatures were on average within 3°C of thermal model predictions.



Ground Segment

The Wallops Flight Facility (WFF) handles the communications, commanding, and data downlink for CeREs. WFF maintains UHF Radar that supports a number of small satellites all using the same Cadet radio and GSE. The natural gain provided by the 18.3 m dish enables high data rates (3.0 Mbit/s) for these missions. WFF is staffed for pass times that occur between 07:30 and 23:30 EDT during the week, with commissioning exceptions available, as necessary, with coordination.

CeREs uses the VMMOC (Virtualized Multi-Mission Operations Center) at GSFC. The VMMOC will connect directly to the GSE at WFF and will be the primary point of contact with the spacecraft. ITOS (Integrated Test and Operations System) will have a specific database for the CeREs mission, leveraging those generated by other supported CubeSat missions. The VMMOC will handle all commanding and downlinking of telemetry from the spacecraft. All telemetry will be captured by the VMMOC, forwarded to the MEDS (Mission Engineering Data System), and submitted to the TaaS (Telemetry as a Service) web portal for use by the SOC (Science Operations Center). TaaS will archive telemetry as received from the VMMOC and provide a web interface for plotting or export selected data. MEDS will process the raw telemetry from the spacecraft utilizing the COSMOS command and telemetry system. MEDS is a virtual machine that exists on multiple NASA computers at GSFC and is managed by the CeREs PI, PM, and Lead Systems Engineer. Since MEDS is a virtual machine, it is fully portable and can be distributed as necessary.

The communications system connecting WFF and CeREs is shown in Figure 11. The downlinked data from the VMMOC is obtained and will be sent to the SOC (Science Operations Center) located at SwRI to process raw telemetry into scientifically usable data.

Figure 11: Communication system connecting the MOC to CeREs using the Wallops Flight Facility (image credit: NASA)
Figure 11: Communication system connecting the MOC to CeREs using the Wallops Flight Facility (image credit: NASA)

Commanding: There are two types of stored XB1 commands, absolute and relative time sequences or macros. The absolute time stored commands use TAI time (in seconds since J2000) at which the command should be executed. Four hundred of these commands may be stored onboard the spacecraft and up to eight may be executed in one 5Hz cycle. When uploaded from the ground, these commands are stored in RAM and will not persist through a reset. Stored commands may be overwritten and a ground command exists to clear all stored commands.

CeREs will typically use the absolute time sequences to trigger macros for nominal operations. Macros can be triggered to immediately begin or to start at an absolute time, utilizing a macro execute command encapsulated in an absolute time command uploaded from the ground. Note that macros also have the ability to trigger themselves upon completion, if desired. A macro is a list of commands whose execution is scheduled for a specified time after initiation. Absolute commands may be used to initiate a macro a specific time. A macro may call itself, creating a repeating process. Macros are numbered and their ordering is important due to the way in which they are disabled. For example Macro #51 covers nominal orbit configurations.

Mission Operations

Deployment and Commissioning: As the spacecraft exits the deployer, separation switches will be released, allowing the spacecraft to power up and boot the FSW. The XB1 will begin the transition to sun point mode and a deployment macro will activate and run through the deployment sequence for both the antennas and the solar panels, twice. This is timed to ensure at least one of the attempts will be made during insolation (sun exposure). The deployment macro disables the low voltage checks, as a successful deployment is required to run in order to charge the spacecraft in sun point mode and to communicate with the ground. Upon completion of the deployment sequence, the beacon will begin at a rate of once per minute until the rate is modified by the ground station via macro command from 1 minute to 10 seconds for approximately 15 minutes before returning to 1 minute.

Nominal Orbit: A nominal science orbit, defined and controlled via an orbit macro, is broken into nine stages. Each stage is targeted to run over a range of latitudes. The actual run time for each stage is based solely on wait commands that may need to be adjusted on orbit, using predicted TLEs and orbit determination, to account for the actual orbit's timing. Six of the eight stages are a form of science mode, either MB or SEP, and are based around the polar-regions, while the other two are utilized to publish telemetry to the radio. The different regions of a CeREs orbit and the modes of operations in those regions are shown in Figure 12.

While in a science mode, the spacecraft will be in the fine reference-pointing mode with the primary axis (–Z) pointed to zenith and the secondary axis (solar array, –X) to the sun. While not in science mode, the pointing of primary and secondary axes will be exchanged to prioritize charging. Once all stages are complete, the macro simply runs again immediately for the next orbit. The orbit macro is designed to start at the equator and will be re-triggered occasionally via absolute time sequence to keep the stages in sync with the orbit. When the CeREs team detects drift, a new absolute time sequence will be loaded via ground command to restart the orbit macro. The nominal values for the transitions into different modes are command able from the ground, with the nominal values designed to capture mostly microbursts while traversing the radiation belt horns, and, mostly solar particles over the polar caps.

Figure 12: CeREs nominal orbits showing the different modes and their dwell times (image credit: NASA)
Figure 12: CeREs nominal orbits showing the different modes and their dwell times (image credit: NASA)

Contingencies and special events: As the CeREs orbit begins to decay or change due to drag, the macros used to drive science operations will also become invalid. This is due to the fact that they are based upon relative time passing from the start and not actual orbit parameters. To account for this, the science macros will be stopped, updated as necessary, and restarted again. Should an event check be triggered on board the spacecraft, the immediate response will be to downlink data that was generated after the prior pass for further examination. Once mission engineers have verified the health of the spacecraft steps may be taken to resume nominal operations. If the event check is no longer enabled, individual commands will used to set the event checks back to the nominal state. A majority of event checks simply place the spacecraft back into sun-point mode. The nominal science setup will also need to be performed in order to resume operations.

Special events, such as a CME impact, or a conjunction with other spacecraft studying allied phenomena could benefit from measurements from the CeREs spacecraft. Additional macros will be created and placed onboard that include a wait and a duration of time the high data rate collection should be run. The nominal orbit macro will be halted in order for this to occur and resumed once special operations are concluded.



1) Shrikanth Kanekal, Lauren Blum, Eric Christian, Gary Crum, Jeff Dumonthier, Allison Evans, Thomas Flately, Ashley Greeley, Sergio Guerro, Agbontaen Imasuen, John Lucas, James Mackinnon, Nikolaos Paschalidis, Deepak Patel, Khary Parker, Quintin Schiller, Errol Summerlin, George Suarez, Mihir Desai, Stefano Livi, Kristi Llera, Joey Mukherjee, Keiichi Ogasawara, "CeREs: The Compact Radiation belt Explorer," Proceedings of the 32nd Annual AIAA/USU Conference on Small Satellites, Logan UT, USA, Aug. 4-9, 2018, paper: 2018-02-06, URL:

2) S. G. Kanekal, D. N. Baker, M. G. Henderson, W. Li, J. F. Fennell, Y. Zheng, I. G. Richardson, A. Jones, A. F. Ali, S. R. Elkington, A. Jaynes, X. Li, J. B. Blake, G. D. Reeves, H. E. Spence, C. A. Kletzing, "Relativistic electron response to the combined magnetospheric impact of a coronal mass ejection overlapping with a high‐speed stream: Van Allen Probes observations," Journal of Geophysical Research, Space Physics, Vol. 120, pp: 7629–7641, September 2015,

3) "Rocket Lab successfully launches NASA CubeSats to orbit on first ever Venture Class Launch Services mission," Rocket Lab, 16 December 2018, URL:

4) Stephen Clark, "NASA, Rocket Lab partner on successful satellite launch from New Zealand," Spaceflight Now, 17 December 2018, URL:

5) "Upcoming ELaNa CubeSat Launches-ELaNa XIX," NASA, August 2018, URL:

6) S. G. Kanekal, et al., "CeREs, A Compact Radiation Belt Explorer to study charged particle dynamics in geospace," Journal of Geophysical Research, to be submitted

7) Shri Kanekal, Eric Christian, Georgia de Nolfo, Tom Flatley, Nick Paschalidis, Mihir Desai, Stefano Livi, Keiichi Ogasawara, "CeREs: A Compact Radiation Belt Explorer," 2014, URL:

8) Dylan Rudolph, Christopher Wilson, Jacob Stewart, Patrick Gauvin, Alan George, Herman Lam, Gary Crum, Mike Wirthlin, Alex Wilson, Aaron Stoddard, "CSP: A Multifaceted Hybrid Architecture for Space Computing," Proceedings of the 28th Annual AIAA/USU Conference on Small Satellites, Logan, Utah, USA, August 2-7, 2014, paper: SSC14-III-3, 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 (

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