Minimize Van Allen Probes

Van Allen Probes / former RBSP Mission

On November 9, 2012, NASA renamed the former RBSP (Radiation Belt Storm Probes) Mission to Van Allen Probes, in honor of the late James Van Allen, the head of the physics department at the University of Iowa, who discovered the radiation belts encircling Earth in 1958. During his career, Van Allen was the principal investigator for scientific investigations on 24 Earth satellites and planetary missions, beginning with the first successful American satellite, Explorer I, and continuing with Pioneer 10 and Pioneer 11. 1) 2)

The Radiation Belt Storm Probes mission is part of NASA’s LWS (Living With a Star) Geospace program to explore fundamental processes that operate throughout the solar system, in particular those that generate hazardous space weather effects near the Earth and phenomena that could affect solar system exploration.

Background: Earth's radiation belts are often referred to as the “Van Allen Belts” due to their discovery by James Van Allen and his team at the University of Iowa in 1958. The radiation belts were discovered during the flight of the very first American satellite. Van Allen and colleagues had installed a Geiger-Müller tube on Explorer 1 to detect cosmic rays, and as the satellite made its eccentric orbit around the Earth, the readings periodically went off the top of the counter’s scale. It happened again during the flight of Explorer 3 several months later. Several followup missions proved that the space around Earth was not empty, but instead enriched with electrons, protons, and energy created by interactions between Earth's magnetic field (or magnetosphere), the solar wind, and (occasionally) cosmic rays arriving from beyond the solar system (Ref. 48).

The prime goal of the RBSP mission is to understand the sun’s influence on the Earth and near-Earth space by studying the planet’s radiation belts on various scales of space and time. The LWS Geospace program will launch two spacecraft, the Radiation Belt Storm Probes, to discover the fundamental physics underlying the source, loss, and transport processes that govern the radiation belts. Observations from the two spacecraft will enable the development of empirical and physics-based models for the radiation belts. The empirical models will be used by engineers to design radiation-hardened spacecraft, while the physics-based models will be used by forecasters to predict geomagnetic storms and alert both astronauts and spacecraft operators to potential hazards. The knowledge gained from the mission will be applicable to particle acceleration processes occurring throughout the plasma universe. 3) 4) 5) 6) 7) 8) 9)

The RBSP mission seeks to resolve decades-old scientific mysteries of how these particles become energized to such high levels, and how the radiation belts vary so dramatically with changing conditions on the sun.

The mission’s science objectives are to:

• Discover which processes, singly or in combination, accelerate and transport radiation belt electrons and ions and under what conditions.

• Understand and quantify the loss of radiation belt electrons and determine the balance between competing acceleration and loss processes.

• Understand how the radiation belts change in the context of geomagnetic storms.


Figure 1: A cutaway model of the radiation belts with the 2 RBSP satellites flying through them (image credit: NASA) 10)

Legend to Figure 1: The radiation belts are two donut-shaped regions encircling Earth, where high-energy particles, mostly electrons and ions, are trapped by Earth’s magnetic field. This radiation is a kind of “weather” in space, analogous to weather on Earth, and can affect the performance and reliability of our technologies, and pose a threat to astronauts and spacecraft.

The inner belt extends from about 1000 to 8000 miles above Earth’s equator. The outer belt extends from about 12,000 to 25,000 miles. This graphic also shows other satellites near the region of trapped radiation.

The instruments on the two RBSP spacecraft will provide the measurements needed to characterize and quantify the processes that produce relativistic ions and electrons.




In July 2000, NASA/GSFC was designated as the NASA Lead Center for the LWS program by the Office of Space Science (OSS), of NASA Headquarters. As Lead Center, GSFC has primary responsibility for managing the implementation of the LWS program and its associated projects.

In 2006, JHU/APL (Johns Hopkins University/Applied Physics Laboratory) of Laurel, MD was awarded a NASA contract to design, built, integrate and operate the twin probes mission. Following a successful confirmation review in late 2008, NASA has given the JHU/APL the go-ahead to continue development of the RBSP mission. APL will build and operate the twin probes that will study the radiation belts surrounding Earth, with a primary mission of two years. 11) 12)

The construction phase of the two RBSP spacecraft started in January 2010 following a three-day CDR (Critical Design Review) in December 2009.

The RBSP mission is composed of two minisatellites with identical sets of instruments to measure charged particle populations, fields, and waves in the inner magnetosphere. The spacecraft are designed to be sun-pointed spinners in near-equatorial elliptical orbits with apogees inside the geosynchronous orbit. Near-equatorial orbits and spin orientation were chosen in order to maximize the coverage of the equatorial pitch angle distributions.


Figure 2: Artist's view of the RBSP spacecraft constellation (image credit: JHU/APL)

Spin axis orientation: Both RBSP spacecraft are stable inertial spinners (nominally at 5 rpm), with the spin axes pointed generally in the direction of the sun. Since the spacecraft spin axes stay inertially fixed, but the Earth moves about the sun nearly 1º/day, it is necessary to maneuver the spin axis of each spacecraft by an average 1º/day to maintain general sun pointing. Attitude maneuvers, however, are scheduled to occur only every 21 days, during which the RA (Right Ascension) of the Sun will increase by approximately 21º. To accommodate this solar motion, the attitude maneuvers will reposition the spin axes approximately 10.5º east (+RA) of the sun. For the next 21 days, the sun will march eastward (in a +RA direction), passing by the spin axes on day 11, and reaching a point 10.5º east of the spin axes on day 21. The process is then repeated throughout the duration of the mission. 13)

Additionally, there is a seasonal north-south bias of the spin axes. When the sun is in the northern hemisphere (March 21 – Sept 23), the spin axes are offset 17º south of the ecliptic plane, while when the sun is in the southern hemisphere (September 23 – March 21), the spin axes are offset 17º north of the ecliptic plane. Figure 3 shows the attitude control concept for both the in-plane and out-of-plane components.


Figure 3: Expected sun angles for spin axis pointing (image credit: JHU/APL)

The mass of the two minisatellites is < 1500 kg. The mission duration is 2 years (expandable to 4 years).

• Operate through challenging radiation environment. The spacecraft and its payload are required to operate continuously while the RBSP spacecraft transits through the heart of the inner-trapped proton Van Allen belt twice every ~9 hour orbit for the nominal 2 year +75 day mission. These energetic (up to hundreds of MeV) protons provide the majority of the penetrating dose and all of the displacement damage. The second major contribution to the total radiation dose is from the outer belt trapped electrons that bombard the spacecraft during the long exposures near apogee.

• Provide attitude control through spin stabilization to provide required instrument fields-of-view, nearly sun pointed with nominal spin rate 5 rpm

• Provide power system to operate through eclipses up to 114 minutes

• Downlink an average daily data volume of at least 6.4 Gbit of recorded plus real-time data per day during the operational phase of the mission

• Accommodate significant payload mass (130 kg) and average power (149 W)

• Provide deployed science booms for magnetometer and search coil instruments

• Provide deployed axial and wire radial booms for radio wave measurements.

Table 1: Overview of basic spacecraft requirements (Ref. 12)


Figure 4: Illustration of the RBSP spacecraft (image credit: JHU/APL)

The spacecraft structure consists of a primary load-bearing central cylinder and aluminum honeycomb decks for mounting instruments and spacecraft components. The two observatories are held together by a RUAG-supplied Lightband low-shock separation system. This same separation system is also used between the stacked observatories and the launch vehicle. 14)

The RBSP spacecraft has a single string fault tolerant architecture. Critical single string spacecraft components use un-switched power and have the ability to be power cycled (or “off-pulsed”) in the event that a radiation induced failure causes a fault that requires removal of power. Critical boxes can be off-pulsed individually or as a group. Both software and hardware command loss timers are part of this “off-pulse” architecture and result in a power cycle of spacecraft electronics if a specific command is not received for a defined duration.

The spacecraft block diagram is shown in Figure 8. The avionics for the system are contained in the IEM (Integrated Electronics Module). The IEM consists of five cards connected across a common backplane. A 32-bit PCI bus, clocked at 16.5 MHz, connects the single board computer (SBC), solid-state recorder (SSR) and Spacecraft Interface Card (SCIF) for flow of commands and telemetry. The SBC is a BAE RAD750 based design clocked at 33 MHz for 50 MIPS (nominal),with 16 MB of SRAM, 4 MB of EEPROM and 64 kB of PROM. The SSR contains 16 Gbit of SDRAM memory with EDAC and hardware scrubbing. The selected SDRAM has a low upset rate (even in the RBSP environment); the few SDRAM errors that are expected will be corrected by SSR EDAC.

The SC IF card contains a custom FPGA design that implements interface logic and thruster control. The board also houses the spacecraft precision oscillator which is used for generation of MET (Mission Elapsed Time). The DC-DC converter card provides regulated secondary voltages derived from the s/c primary power bus and implements the box off-pulse capability. The telemetry card gathers temperature, analog and discrete data and is connected to the SC IF card via an internal I2C bus.

The IEM handles both commands and telemetry data flow to each instrument via 115.2 kbaud UART (Universal Asynchronous Receive Transmit) links. The UARTs are synchronized to the spacecraft timekeeping system via a 1PPS (One Pulse-per-Second) interface. Commands to the instruments and other S/C bus components are sent out via two sequenced transmission buffers with the delay from 1PPS dependent upon the prior command buffer usage.


Figure 5: The IEM is shown installed on spacecraft A (image credit: JHU/APL)


Figure 6: Block diagram of the IEM (image credit: JHU/APL)

An electrically-isolated function within the IEM is the HWCLT (Hardware Command-Loss Timer) utilized as part of the fault management system. This is a discrete, logic-based circuit that maintains a count-down between successive “reset” pulses from the ground. If the HWCLT is not “reset” by a specific command sent from the ground within 3.58 days, a logic pulse is sent to the PDU which initiates a PDU sequence to off-pulse the PDU and then the IEM and XCVR (Transceiver). As with all off-pulse implementations, there are multiple levels of protection on this action including an inhibit feature within the PDU itself and two physical interfaces to each box being off-pulsed. The 3.58 day duration is set based upon other fault management mitigation events such as a software based command loss-timer and specific actions initiated through the ground.

C&DH (Command and Data Handling): The C&DH subsystem is resident in the flight computer. The C&DH software is a set of functional applications and libraries, which were implemented to be used with the core Flight Executive (cFE) software developed by NASA/GSFC and the VxWorks operating system. The cFE software provides standard services with a standard API (Application Programmer’s Interface). The RBSP C&DH software is composed of seventeen (17) unique applications and eight (8) libraries. Each of these applications and libraries has been designed and implemented to perform a specific set of functions. Having a single common design and single common code base had the advantages of improved software review, testing and maintenance, and in the end reduced the total software development effort, producing a quality C&DH flight software system. As RBSP is the first mission developed by APL using this architecture, it is expected that future cFE application software development at APL will benefit from this application framework. 15)

EPS (Electrical Power Subsystem): Each spacecraft utilizes a DET (Direct Energy Transfer) power system topology. The power bus voltage varies with the 8-cell Li-ion battery voltage. The EPS consists of the PSE (Power System Electronics), BME (Battery Management Electronics), SAJB (Solar Array Junction Box), the 50 Ah Li-ion battery, and four deployed solar array panels. A simplified block diagram of the power system is shown in Figure 7.

The PSE consist of a single fault tolerant sixteen stage sequential analog voltage control shunt regulator with maximum battery current limit. The loads are connected to the single 8-cell, 50Ah Li-ion battery via the PDU (Power Distribution Unit). The nominal bus voltage is 30 V and can vary between 24 and 32 V depending on the state of charge and temperature of the battery. Each battery cell can be by-passed by a bypass switch that is activated by ground command to remove a single cell from the battery in case of a pending cell failure. In the case where by-pass switch activation has occurred, the corresponding bus voltage range becomes 21 to 28 V.

The primary battery charge control method is CC/CV (Constant Current followed by a Constant Voltage) taper charge. The battery is charged at a high rate, limited to C/5, where C is the battery capacity, using the available S/A power that is not used by the loads until the battery SOC (State-of-Charge) reaches 60%. The on board coulometer then reduces the battery charge current to C/10. The battery maximum voltage is controlled to preset safe levels via voltage (V) limits that are implemented in the single fault tolerant voltage regulator. Whenever the battery voltage reaches the V limit, the V control loop will force the charge current to taper.


Figure 7: Simplified block diagram of the RBSP spacecraft power system (image credit: JHU/APL)

The BME consists of an interface board and a cell shunt board. Each battery cell has a parallel connected analog shunt used during the mission to balance the end of charge voltage of each Li-Ion battery cell. Each cell shunt is limited to 0.75A maximum current bypassed around the cell in order to limit the amount of power dissipated in the BME. The BME contains eight relays which allow the battery cells to be disconnected from cell shunts to limit leakage current during ground operations or whenever the BME sis not powered. During safe mode operation, the current controller and BME are not powered and the system relies on the single fault tolerant voltage limit regulator. The average spacecraft load power during flight is expected to be 277 W.

Solar array: The RBSP solar array consists of four deployed panels with a total active area of 3.2 m2. Each panel is approximately 0.739 m wide and 1.26 m long. The panel substrates are 25.4mm thick aluminum honeycomb with composite face sheets. The panel front cell side is insulated with Kapton, co-cured with the graphite fiber face sheet. The back face sheet is not painted. Three different solar cell sizes are used to maximize the cell packing density. Each panel contains 12 strings of 24 series connected 28.3 cm2 solar cells. The boomless panels (two panels of the four) contain an additional two strings of 22 series connected 26.62cm2 solar cells and four strings of 22 series connected 11.5 cm2 solar cells. The boom panels (two panels of the four) contain an additional string of 22 series connected 26.62cm2 solar cells and two strings of 22 series connected 11.5cm2 solar cells. The solar cells are triple junction cells with minimum efficiency of 28.5% [BTJ (2nd Generation Triple-Junction)], from EMCORE Photovoltaics. The cover glass on each cell is 0.5 mm thick cerium-doped microsheet, from Qioptiq with Indium tin oxide coating.

The PDU (Power Distribution Unit) provides switched, unswitched, and pulsed power to the spacecraft components. The PDU receives primary power from the PSE and has a serial UART command/telemetry interface with the IEM. The PDU box is a modular slice design. Each slice consists of a printed circuit board housed in a mechanical frame, and the slices stack and bolt together. The slices are electrically connected using internal rigid-flex connectors for signals. A wiring harness external to the box is used for power connections. A solid 350 mil thick aluminum chassis and solid 150 mil aluminum radiation shields (located in thinned areas of the PDU chassis) mitigate the effects of radiation on the electronics parts, allowing the PDU to function nominally in a high radiation environment.

ADCS (Attitude Determination and Control Subsystem): The science investigation requires a post-processed attitude knowledge of ≤ 3º (3σ). Both spacecraft are nominally sun-pointing and spin stabilized. To maintain the observatory’s sun sensors within the operational range and to ensure sufficient power system margins, the total off-pointing angle between each observatory and the Sun-Earth line is required to remain within the range of 15-27º. Furthermore, each observatory is required to operate within the spin range of 4-6 rpm during normal operations, and to maintain the spin rate within 0.25 rpm once the nominal spin rate is established.

Attitude determination is accomplished by two means. The flux-gate magnetometer sensor is the primary sensor for determining definitive observatory attitude. The data from this sensor is not processed on-board, so additional sensing is required to allow for autonomous attitude sense for spacecraft health and safety. This on-board attitude determination is accomplished via two sun sensors. These sensors provide coarse spacecraft attitude information that is sufficient for autonomous spacecraft health and safety. This data is also sent to the ground where it is used in the estimation of definitive spacecraft attitude that is used for science processing. Each Observatory also includes two passive nutation dampers that help maintain a stable spacecraft attitude and damp out any “wobble” after propulsive maneuvers.

The propulsion system is a simple monopropellant blowdown system. Three Inconel tanks store the 56 kg of hydrazine propellant onboard, and feed the eight 0.9 N thrusters. The position and orientation of these eight thrusters allow for spin up and spin down about the primary spin axis, positive and negative precession about the spinplane axes, and velocity change toward and away from the sun. These thrusters are the only active attitude control mechanisms on board the spacecraft, and provide the full set of capability required to maintain the spacecraft attitude and spin rate and perform orbit corrections, collision avoidance maneuvers as required, and the final de-orbit burn at the end of the mission.

The spacecraft also includes diagnostic instrumentation – termed the ERM (Engineering Radiation Monitor) – to monitor the in-situ radiation environment. The long-term health and operability of the spacecraft electronics and materials are directly affected by the total incident radiation dose. The ERM will measure this incident dose, and will provide a means for correlating upsets in spacecraft electronics with the environment present at that time. This monitor will also allow refinement of the standard total dose curves that are traditionally used for the design of spacecraft that operate in the Earth’s radiation belts.

Fault protection: Each spacecraft also includes robust APA (Fault Protection and Autonomy) systems that work together to maintain the overall health and safety of the flight segment. Because the spacecraft includes limited hardware redundancy, the FPA systems are of particular importance on RBSP. The architecture uses a layered response approach to maximize science data collection in the event of a fault. The system protects against the extended loss of communications by way of both software and hardware command loss timers. It also monitors the sun angle of a given observatory, and can safe the system and notify the ground in case of an exceedance of the minimum or maximum sun angle. Similarly, the FPA system monitors the bus voltage and battery state of charge, and can safe the system in case of a LVS (Low Voltage Sense) or a LBSOC (Low Battery State of Charge) where the bus voltage or battery state of charge drop below a minimum preset level.

The system also monitors the current condition and the health of spacecraft components, and it has the ability to individually off-pulse the primary unswitched loads (the IEM, the PDU and the transceiver) to restore those systems to a known startup configuration and presumably to clear any faults that my be present. In addition to monitoring and managing spacecraft bus health and safety, the system can also monitor instrument currents and heartbeats and can power off instruments in the case of a fault, and they can also individually power off the instruments based on a turn off request generated by that instrument.

All parts used in the RBSP observatory were specified to survive a total ionizing dose of 34 krad (Si) [23 krad (Si) for the IEM] without parametric or functional failure. This value is based on a 2-year (plus 75 day) life, with a RDM (Radiation Design Margin) factor of 2, and a nominal shield depth of 350 mils (6.3 mm) [500 mils (12.7 mm) for the IEM] of aluminum.

Lastly, the fault protection system manages the separation sequence after launch, deploying the solar arrays and powering on the RF downlink (uplink is enabled by default at launch).


Figure 8: Block diagram of the RBSP spacecraft (image credit: JHU/APL, revision Oct. 28, 2011, Ref. 12)


Figure 9: RBSP spacecraft layout with side panels in a non-flight “open” orientation (image credit: JHU/APL)


Figure 10: Photo of the RBSP-A spacecraft in Nov. 2011 during testing on the solar arrays (image credit: JHU/APL) 16)

Spacecraft parameter

Current best estimate

Not to exceed capability


Spacecraft dry mass

609.4 kg

743 kg



56 kg

56 kg


- Normal: 15-27º
- Safe: 27-33º

277 W
233 W

350 W
332 W


Thermal bus environment

0 to +30ºC

-20 to +45ºC




151.4 m/s


G&C attitude knowledge of S/C
-Spin axis control
- Spin rate control

3.1º (3σ)
±0.25 rpm

3.1º (3σ)
±0.25 rpm


Average instrument data rate

72 kbit/s

78 kbit/s


Onboard data storage

16 Gbit

16 Gbit


Table 2: Overview of RBSP spacecraft resources (Ref. 12)


Figure 11: Technicians at the Astrotech payload processing facility prepare the RBSP spacecraft for encapsulation in the payload fairing (image credit: NASA)


Launch: The two identical RBSP spacecraft were launched on August 30, 2012 from the Cape Canaveral Air Force Station (launch complex 41) in Florida. The launch provider was ULA (United Launch Alliance), using an Atlas-V 401 launch vehicle. 17) 18) 19)

The probes were released from the rocket's Centaur upper stage one at a time and sent off into different orbits, kicking off the two-year mission to study Earth's radiation belts. The RBSP-A spacecraft separated from the Atlas rocket's Centaur booster 1 hour, 18 minutes, 52 seconds after launch. The second spacecraft, RBSP-B, separated 12 minutes, 14 seconds later.

After deployment of RBSP-A, the Atlas-V will raise its own apogee to approximately 30540 km and then deploy RBSP-B. This slight difference in apogee altitude values will cause RBSP-A to lap RBSP-B approximately once every 75 days (201 & 200 orbits respectively).

Orbit: Near-equatorial HEO (Highly Elliptical Orbit, almost like GTO), perigee ~ 620 km, apogee ~ 5.8 RE (~ 30,500 km), inclination = 10º, period ~9 hours.


Figure 12: Illustration of the two RBSP spacecraft in their orbit (image credit: JHU/APL)


Figure 13: RBSP mission design (image credit: JHU/APL)

RF communications: The RF communications system contains a single transceiver, an 8 W SSPA (Solid-State Power Amplifier), a diplexer, and two broadbeam, near-hemispherical antennas. The system provides S-band uplink, downlink and radiometric tracking capability. It supports both ½ turbo and convolutional encoding, and it uses coherent downlink to allow for Doppler navigation.

The mission uses the S-band only version of the Frontier Radio, built on JHU/APL's SDR (Software Defined Radio) architecture. The Frontier Radio is based on coherent transceiver (XCVR) technology and is compatible with NASA’s STRS (Space Telecommunications Radio System) architecture. The downlink data rate is up to 2 Mbit/s (QPSK modulation). In addition, highly accurate coherent Doppler data is needed for spacecraft navigation (the orbit determination process uses the Doppler data). 20) 21)

Due to its (nominally) 5 rpm spin, the spacecraft is designed with the two antennas’ boresights parallel to the spin axis to ensure uninterrupted telecommunications while spinning. Each antenna provides sufficient gain (–4 dBic minimum) from its boresight to 70º from boresight.

• Two antennas: S-band conical bifilar helix, circular polarization, broadbeam

• RF routing: procured power divider & diplexer (L-3 Communications Corporation)

• SSPA (Solid-State Power Amplifier): 8 W, S-band power amplifier

• Frontier Radio: DC/DC, DSP (Digital Signal Processor), receiver, exciter slices.


Figure 14: The RBSP mission requires broadbeam antenna coverage from boresight to 70º for each antenna. The antenna is shown on the right, with and without a radome (image credit: JHU/APL)


Figure 15: Orbital configuration of the RBSP observatory (image credit: JHU/APL)



Frontier Radio (Demonstration Payload):

The Frontier Radio is a low-power, low-mass, modular SDR (Software Defined Radio) platform designed for communications, navigation, radio science, and sensor applications- the first radio with full software implementation. The objective is to demonstrate a TRL-6 (Technology Readiness Level) of 6 and to be qualified for a spaceflight mission. 22) 23) 24)


Figure 16: Block diagram of the single-string RF communications subsystem (image credit: JHU/APL)

Link Function



Receive (rx)


≥ 1


< 5 MHz


B/QPSK, PM, Subcarrier





DSP Power [HW (Hardware), FW (Firmware), SW (Software)]

< 2 W (primary 28 V bus)

Transmit (Tx)

Coherent w/Rx



≥ 2


> 100 MHz


I/Q, ≥ 4 bit/symbol




Convolutional, Turbo

DSP power

Rx Power + < 3 W



Mission critical



≥ 50, ≥ 100 krad (goal)


Temperature range

-30ºC to +60ºC


Mass, volume

≤ 400 g, ≤ 400 cm3

Table 3: High level design constraints for the DSP (Digital Signal Processor) system, including hardware, firmware, and software 25)

Hardware: The radio level constraints focused the DSP system design effort on specific paths. The first set of trades focused on the hardware necessary to process the communication formats listed in Table 3. The demodulation scheme for even these simple phase modulation formats costs 350 to 450 MOPS, which is derived from the original New Horizons FPGA design. These operations do not include state machines or overhead control logic, they are mostly simple 32 bit operations on the main data stream. 10 to 20 MOPS (Million Operations Per Second) are necessary for controlling the radio hardware and communicating with the spacecraft, and (if a lookup table based modulation and encoding architecture is used) the transmission path costs 3-5 OPS per transmitted symbol (300-500 MOPS for the radio goal data rates). There was no space grade standard processor that could come close to meeting these MOPS requirements while fitting within the receive only power consumption specification. The TI C6701 DSP processor came close to these capabilities, but the radiation reliability, SEE (Single Event Effect) rate, on this device did not meet the requirements. There was no single commercial solution available.

Processor combinations with a FPGA were investigated as another option, but were again ruled out due to total power consumption or the board area required to create multiple high efficiency core voltages. During the investigation process the math operations themselves were looked at more closely. Nearly all of the high rate operations were additions or subtractions for the CIC (Cascaded Integrator-Comb) filter based receive channel, and base 2 finite field arithmetic for the transmit channel. Most of the difficult operations, multiplication and division, were carried out at a much lower rate, <5 MOPS, to create control loop filters or gain elements. The remaining operations were low rate basic state machines and mathematics for hardware control and communication with the spacecraft interface.

Very quickly it became clear that the ideal solution was a front end FPGA for filtering and demodulation, a low power DSP processor running in real time to close control loops, and a low rate RISC processor to handle abstraction of the radio hardware and firmware up to the spacecraft interface. At that time FPGA gate count had risen to the point where it became possible to imbed all of these technologies into one FPGA device. Designing and implementing a system of this level of complexity for a new radio had significant cost and schedule risk. The benefits of limiting the DSP design to one device, thereby reducing board area and interconnect I/O power, prevailed in this trade.

The next hardware trade was the selection of the target FPGA device, and the main contenders were the Actel RTAX and Xilinx Virtex-4 device series. SRAM based FPGA devices had exceptional processing capabilities due to their larger gate count and faster master clock rate, plus they are completely reconfigurable which makes them attractive for an SDR (Software Defined Radio). The main problems with this choice were reliability in radiation and power consumption. The former issue fed the latter because of triple voting and the need for an external device to continually refresh the configuration memory. The SRAM device also required an array of non-volatile memories for the configuration file which would use up valuable board space resources. The Actel RTAX device family was finally selected because it allowed the entire DSP system to meet all of the radio high level design constraints listed in Table 3.

Firmware platform: The first work done on the Frontier Radio firmware architecture was to create a platform to which all of the other firmware and hardware elements would be connected. This effort was the selection of the low rate RISC processor and its interfaces. At that time, there were only a few processor IP cores available from Actel, but a wide selection in the greater FPGA community. Most of these processors fell into the three main categories of too small, too large, and too unique. The small processors were 8 to 16 bit, and most were based upon the 8051 architecture.

Manipulating the 32 and 64 bit fixed point numbers of the existing New Horizons firmware design would have been difficult for these processors, as well as controlling the >24 bit address space. Processors like the LEON3 offered extensive capability, but performed a significant number of unnecessary functions like IEEE-754 floating point mathematics. These designs would have forced the firmware to grow into, and possibly beyond, the RTAX40009 device size. Most of the other processor IP designs made significant use of specific Xilinx or Altera devices resources that were not available in the RTAX design. Many of these custom processor designs were also unique, like the SCIP (Scalable Configurable Instrument Processor) developed at JHU/APL for Actel FPGA based space instrument control. This is a stack based processor that only had an existing compiler for the Forth language. These unique processors would have taken too long to modify for the RTAX family, space applications, or to connect to the rest of the radio. They would have also taken too long to develop a programming flow.

The final selection choice was an instantiation of the MIPS3000 architecture. This particular design did not include a floating point unit or any other unwanted peripherals, which was desirable for the current radio requirements and allowed for upgrades to add these capabilities in the future. The most favorable aspect of the MIPS3000 was the wide variety of software tools available to program the device.

Once the general purpose processor was chosen the firmware architecture began to take shape. Each processing element, the MIPS processor being one of them, was named a "core", and all of the cores were connected to a standard microprocessor bus. The cores that control access to the external memory banks and the spacecraft interface were placed directly on the main bus of the MIPS processor. All of the other cores were placed on a separate APB (Advanced Peripheral Bus), which was connected to the MIPS bus via a custom controller core. Separation of the MIPS bus from the APB was done very early, before specifications were complete on each of the cores. The unknown requirements (bus clock rates, latency, and activity) for the new DSP cores were the main reason for this separation. All high rate communication between cores would be carried out with custom core to core buses.

The separation of the core bus connections helped the design and development process in several ways. Foremost, the design allowed the design team to develop the general purpose processing platform (MIPS bus) separate from the DSP section. A simple APB to UART controller was built to support DSP core development and testing through a Matlab interface. The Matlab low level control tool is still in use today for debugging or detailed testing. In parallel, a firmware and software team was preparing the basic software development tools, planning out memory maps, and outlining the first radio software framework. Placing these two development efforts in series would have broken the project schedule. The other main benefit of the APB is its ability to handle variable response latency gracefully, which is good for real-time processors (the DSP cores) that cannot respond immediately to read/write requests. The APB controller isolates the MIPS from these variable delays, and prevents individual polling routines, and handshaking or interrupt circuitry. It does this by off loading two main burst read/write operations from the MIPS processor and bus.

Firmware DSP Cores: The goal for each of the DSP Core designs was to be as programmable, software defined, as possible while not exceeding the capacity of a RTAX2000 FPGA. The MIPS processor platform itself was allocated ~40% of the FPGA, including the MIPS processor, memory controller, spacecraft interface, APB controller, clock management, and the hardware interfaces. The allocation was derived fairly easily due to reuse of existing designs, as well as the straight forward nature of these circuits. This left the DSP Cores with about 6,000 flip-flops and 11,000 4-input logic cells. At this early stage reusable designs already existed for Convolutional & Turbo encoding (including framing), critical command decoding, demodulation, and individual control loop filters. Given the size of these designs and the FPGA gate area remaining, several key decisions were made about each DSP core. The functionalities, boundaries, and names for each of the cores were mapped out first. Then trades were made on each core concerning software based flexibility, realistic requirements, and FPGA resources. The resulting firmware architecture has remained the same since its creation in 2007, and is shown in Figure 17.

The final step in the firmware process was to break each core down one step further into distinct "blocks" that could each be individually specified, designed, implemented, and tested within approximately one month of time. This was done for three main reasons; modularity, team work, and incremental testing. Modularity is important for STRS compliance, and its use has great benefits in reducing new development cost. This allows one functional element within each core (e.g. CIC filter, ALU, or encoder core circuit) to be removed, inserted, modified, or completely redesigned while having minimal effects on the surrounding circuitry. Modularity also allows for automated customizations at the time the firmware is built for a particular implementation.


Figure 17: Block diagram of the Frontier Radio firmware architecture, including external memory (image credit: JHU/APL)

The Frontier Radio is the focal point of the RF telecommunication system. It receives an S-band uplink for spacecraft commanding, generates a coherent S-band downlink for telemetry and science data (either convolutional or turbo coded) and navigation, and interfaces with the C&DH (Command and Data Handling) subsystem of the spacecraft.

The radio is a low-power, low-mass, modular SDR consisting of four slices: power converter, receiver, exciter, and DSP (Digital Signal Processing) slice. The DSP slice processes both the uplink and downlink signals and handles mode and status communications with the C&DH subsystem. The exciter slice receives baseband data from the DSP slice and modulates it onto a high-quality S-band carrier to generate the downlink. The receiver slice acquires the uplink signal from the ground station and downconverts it to an IF (Intermediate Frequency) that is passed to the DSP slice. The receiver slice also contains an OCXO (Oven Controlled Crystal Oscillator) which is used to generate various clock signals required for the radio. The DSP digitizes the IF signal and further filters and processes it in firmware within an Actel RTAX-2000 FPGA (Field Programmable Gate Array). The power converter slice generates secondary voltages from the main 30 V spacecraft bus (Figure 17).

The Frontier Radio platform was designed for flexibility. Uplink and downlink frequencies, line coding, error correction coding, control loops, and many other features are all programmable. For the RBSP implementation, these features have been determined by the mission requirements and have been hardcoded to maximize reliability in the harsh radiation environment that is anticipated. However, the underlying flexibility of the architecture allowed quick iterations of the design during prototype development.

The two RBSP spacecraft will operate in different RF channels. Both channels adhere to coherent turnaround ratios of 240/221. The information required for these frequency assignments, along with acquisition control loops, calibrations, and other software needed to define the characteristics of the radio are referred to as the Radio Personality and are stored in memory within the radio.

The mission communications system has been designed to operate with the JHU/APL 18 m ground station (APL-18) in Laurel, Maryland, USA, as the primary ground station. Periodic additional contacts will be made over two USN (United Space Network) 13 m ground stations in South Point, Hawaii, USA, and Dongara, Australia. These ground stations provide additional coverage for launch and early operations, emergencies, and science downlink bandwidth when needed.

In addition, communications with NASA's TDRSS (Tracking and Data Relay Satellite System) are foreseen. Low rate downlink and uplink modes are required for TDRSS contacts. The radio microcode had to be optimized for both high and low bit rate performance. The downlink bit rate ranges from 1 kbit/s to 2 Mbit/s for this mission. The downlink bit rate is controlled completely within the radio and is determined by programmable prescaler and divider circuits within the FPGA. Although only six downlink rates are planned for use in the RBSP mission, these circuits allow the radio to be programmed to a variety of other bit rates below 2 Mbit/s. Above that rate, hardware specific to the RBSP design becomes a limiting factor. For this RBSP implementation, the uplink bit rates are 2000 bit/s and 125 bit/s. Other rates are available in the generic radio design.

SDR technology has the most potential to benefit near Earth applications in the immediate future because of the inherent low cost adaptability from build to build, and the ability to change communication formats in flight.

As an example, the RBSP mission communicates primarily with a ground station at JHU/APL, but is also required to communicate through the NASA TDRSS and broadcast space weather information to other ground stations throughout the world. The Frontier Radio allows RBSP to instantly switch between these communication formats using one mode command.

Analysis and testing performed for the RBSP mission has shown very low fault rates under averse radiation conditions, allowing the single string radio to become a key part of the spacecraft autonomous fault recovery system.

Table 4: Frontier radio support for near-Earth applications (Ref. 23)


Figure 18: Configuration of the ground support equipment (image credit: JHU/APL)

Instrument mass

< 1.92 kg

Power consumption

7 W full duplex, 5 W Rx only (30 V nominal)



Coherent with uplink with a turnaround ratio of 240/221 within a 1 σ standard deviation of 3.3 parts in 10-12 at 10 s

Error correction coding

Convolutional: rate ½
Turbo : rate ½, frame length 8920


QPSK, PM (1.2 radian nominal, 0.9 radian for space weather broadcast)

RF power out

> 4 dBm, ± 0.75 dB over temperature

Bit rates

2 Mbit/s, 1 Mbit/s, 500 kbit/s, 250 kbit/s, 125 kbit/s, 1 kbit/s (nominal tested bit rates)

Line coding

Bi-phase-L and NRZ-L

BER performance (convolutional)

Within 1 dB of theory

FER Performance (turbo)

Within 1.5 dB of theory


Noise figure

< 3.5 dB

Acquisition and tracking threshold

-150 dBm

Dynamic range

-70 to threshold


BPSK on a 16 kHz subcarrier

Bit rates

125 and 2000 bit/s

BER performance (unencoded)

Within 1 dB of theory

Table 5: Key performance parameters of the RBSP Frontier Radio

Mission approach:

- Simultaneous two­point measurements by identical spacecraft in common orbits with a slow separation in phase, lapping one another 4-5 times/year.

- Covering the full range of local times in 2 years

- Apogee of ~ 5.8 RE to sample the outer belt and ring current

- Perigee of ~ 630 km to sample the inner belt.

The two RBSP spacecraft will operate entirely within the radiation belts throughout their mission. When intense space weather occurs and the density and energy of particles within the belts increases, the probes will not have the luxury of going into a safe mode, as many other spacecraft must do during storms. The spacecraft engineers must therefore design probes and instruments that are “hardened” to continue working even in the harshest conditions.


Figure 19: Photo of the Frontier Radio (image credit: JHU/APL) 26)


Operation of the Frontier Radio on the Van Allen Probes and future outlook:

The first two Frontier Radios provide a robust, high capability SDR with very low SWaP (Size, Weight, and Power). The targeted applications typically require highly sensitive communications and radio navigation modes, resulting in a baseline architecture that excels at meeting the performance requirements of a large variety of other spaceborne applications. The Van Allen Probes make use of this SDR in an S-band duplex configuration to support mission-critical spacecraft commanding, the return of spacecraft housekeeping telemetry and science data. A transponder function provides for two-way Doppler navigation of the spacecraft. 27)

An on-board CCD (Critical Command Decoder) provides a method for command reception without an external spacecraft processor when needed; this CCD could be used for spacecraft fault detection, radio reconfiguration, or in-flight software uploads should a particular mission require any of those functions. Other advanced features include support for a variety of coding formats (typical convolutional and Turbo codes) and an internal ovenized oscillator for improved timekeeping accuracy and RF performance such as low phase noise and implementation loss.

The highly modular architecture of the Frontier Radio facilitates low-cost, low-risk infusion of new technology or alternate configurations to meet new, unique, or future mission requirements. This is particularly useful for ORS (Operational Responsive Space) applications, where pre- and post-launch reconfiguration is an essential capability. Modularity is a theme carried throughout the hardware, firmware, and software architectures.

Alternate Frontier Radio configurations and modules have been developed that provide Ka-band (26, 32, or 38 GHz) transmit and receive capability, with symbol rates as high as 150 Msample/s and support for high order vector modulation schemes (e.g., 300 Mbit/s QPSK, 450 Mbit/s 8-PSK, and so forth). At the other extreme, an X-band duplex (with an optional dual-band X/Ka-band exciter) configuration supports data rates to below 10 bit/s and two-way ranging for deep space applications.

The SPP (Solar Probe Plus) mission of NASA, with the launch planned for 2018, is highly constrained in mass and power, thus is leveraging this deep space configuration of the Frontier Radio. The wide bandwidth capability of this SDR will enable the SPP mission to take advantage of substantially increased precision during ΔDOR (delta Differential One-way Ranging) operations; this is critical in simplifying the navigation solution for a spacecraft that will fly as fast as 200 km/s at closest approach to the Sun. SPP also requires the Ka-band downlink capability to close highly constrained communications links as the spacecraft optimizes pointing to keep its sun shade properly oriented (in contrast to optimizing for the communications link).

The advanced signal processing capabilities of this SDR can be used to enable new mission operations scenarios that could make or break the viability of a mission. The ability to implement very narrow bandwidth tracking algorithms and more exotic phaselocked-loop implementations (e.g. adaptive or high order loop filters) provide receive sensitivity that mirrors that of the ground receivers used for deep space missions. This further enables the use of advanced decoding algorithms with minimal implementation loss at deep space data rates (e.g. 7.8 bit/s). These features enable advanced capabilities such as future deep space relay or highly sensitive remote sensing (deep space, near earth, or airborne) with a very low SWaP impact. Software management within the SDR further facilitates more advanced operations that may require performance optimization for different operations scenarios or more autonomous signal acquisition, operation, or fault recovery. Software modifications within the current function set can be implemented and tested in as little as a few days to weeks, an essential capability for ORS applications.

The versatility and low SWaP of the existing Frontier Radio platform has proven useful in applying this SDR to remote sensing applications. More specifically, remote characterization of the ionosphere in concert with the DORIS system has many civilian and defense-related applications, and realizing a global system of these remote sensors becomes more economically feasible where hosted payload space and rides of opportunity are leveraged (Figure 20).


Figure 20: GPS and DORIS-based trans-ionospheric remote sensing applications relevant to the Frontier Radio (image credit: JHU/APL)

These situations are much more viable when the SWaP of the payload is minimized. Interest in this area has spawned recent development of the waveforms required to operate the Frontier Radio for ionosphericcharacterization applications. The very low SWaP of the existing platform sets this SDR apart from other high-reliability space-grade SDRs, making it a highly viable choice for hosted payload applications with long mission life. Fortunately, most of the signal processing algorithms required for this application are synonymous with the existing algorithms that are used for the radio communications and navigation functions. Therefore, a minimum development effort is required to support the new waveform. The wide variety of hosted payload opportunities provides for a moving target in terms of available accommodations and resources. The volume of the current Frontier Radio lends itself well to small satellites, in as small as 500 to 1000 cm3 for remote sensing configurations. The required volume depends on the accommodations provided by the host spacecraft. For example, a power converter and customized data interfaces would not be required for Cubesat rides where power regulation and conditioning are routinely provided and relatively standard data interfaces are utilized. In this configuration, power consumption is ~2.5 W for a 5 Vregulated input, which includes an internal ovenized precision oscillator.

Wide application of the Frontier Radio to a variety of small satellite applications is a natural fit, given the very low SWaP yet high capability of this SDR. The built-in processing capabilities (currently an embedded MIPS in FPGA) are more than sufficient for simple autonomous decision making required in some small satellites and probes. By leveraging ever-increasing available FPGA resources, higher capability processors, like the LEON3, can readily be embedded into the SDR’s existing FPGA, alongside the DSP cores. This would provide the capability required for full-featured small satellites, for example, where guidance and control systems might provide 3-axis active pointing. Substantial SWaP savings can be realized in small satellites that leverage these options, essentially embedding multiple systems like communications/RF, command and data handling, and guidance and control within one FPGA. The RF functions within the SDR’s FPGA could equally be ported to another FPGA on the spacecraft to minimize overall SWaP or improve redundancy and reliability. Highly integrated spacecraft systems like these have the power to substantially reduce the spacecraft SWaP (and associated hardware cost) and/or dramatically increase the available science/mission data for each launch. Further, this opens up the window to more advanced cooperative systems such as a mesh of small spacecraft or nodes that perform distributed sensing or form distributed RF apertures.

The Frontier SDR has been mission enabling for its low mass and power and high functionality and multi-band capability. These same advantages lead to its use on the SPP (Solar Probe Plus) mission and on a LEO mission needing high bandwidth, low power, and radiation tolerance. Cubesat and nanosat implementations are in the early stages of design, and are likely to bring the incorporation of multiple spacecraft functions into one central processor. High integration, or the ability to port firmware and software IP to different host processors within a spacecraft, is a game changing capability that will shape future missions and open up new opportunities. The existing SWaP and capability of the Frontier SDR provide an exceptional baseline for new small satellite missions, with substantially untapped processing capability for adding advanced new features (Ref. 27).



Mission status:

• March 7, 2014: Using data from NASA's Van Allen Probes, researchers have tested and improved a model to help forecast what's happening in the radiation environment of near-Earth space — a place seething with fast-moving particles and a space weather system that varies in response to incoming energy and particles from the sun. 28)
When events in the two giant doughnuts of radiation around Earth – called the Van Allen radiation belts — cause the belts to swell and electrons to accelerate to 99 percent the speed of light, nearby satellites can feel the effects. Scientists ultimately want to be able to predict these changes, which requires understanding of what causes them.

Now, two sets of related research published in the Geophysical Research Letters improve on these goals. By combining new data from the Van Allen Probes with a high-powered computer model, the new research provides a robust way to simulate events in the Van Allen belts. The recent work centers around using Van Allen Probes data to improve a three-dimensional model created by scientists at LANL (Los Alamos National Laboratory), called DREAM3D (Dynamic Radiation Environment Assimilation Model in 3 Dimensions). Until now the model relied heavily on the averaged data from the CRRES (Combined Release and Radiation Effects Satellite) mission. 29)

The new technique provides for gathering real-time global measurements of chorus waves, which are crucial in providing energy to electrons in the radiation belts. The team compared Van Allen Probes data of chorus wave behavior in the belts to data from the NOAA/POES weather satellite series in LEO. Using this data and some other historical examples, they correlated the low-energy electrons falling out of the belts to what was happening directly in the belts. Once the relationship between the chorus waves and the precipitating electrons could be established, the research team could use the POES satellite constellation - which has quite a few satellites orbiting Earth and get really good coverage of the electrons coming out of the belts.

The second paper describes a process of augmenting the DREAM3D model with data from the chorus wave technique, from the Van Allen Probes, and from NASA's Advanced Composition Explorer, or ACE, which measures particles from the solar wind. Los Alamos researchers compared simulations from their model – which now was able to incorporate real-time information for the first time – to a solar storm from October 2012. 30)

• In research published Dec. 19, 2013 in Nature, lead author Richard Thorne and colleagues report on high-resolution measurements, made by the Van Allen Probes, of high-energy electrons during a geomagnetic storm of Oct. 9, 2012, which they have analyzed together with a data-driven global wave model. Their analysis reveals that linear, stochastic scattering by intense, natural very low-frequency radio waves known as “chorus” in the Earth's upper atmosphere can account for the observed relativistic electron build-up. 31) 32)

Such electrons in the Earth’s outer radiation belt can exhibit pronounced increases in intensity, in response to activity on the sun, and changes in the solar wind — but the dominant physical mechanisms responsible for radiation belt electron acceleration has remained unresolved for decades.

Two primary candidates for electron acceleration exist, one external and one internal. From outside the belts, a theoretical process known as “inward radial diffusive transport” has been developed; from within the belts, scientists hypothesize that the electrons are undergoing strong local acceleration from very low frequency plasma waves. And controversies exist as to the very nature of the wave acceleration: Is it “stochastic” – that is, a linear and diffusive process – or is it non-linear and coherent?


Figure 21: Schematic illustration of local electron acceleration by chorus (image credit: Jacob Bortnik,UCLA)

Legend to Figure 21: The top panel shows electron fluxes before (left) and after (right) a geomagnetic storm. The injection of low-energy plasma sheet electrons into the inner magnetosphere (1) causes chorus wave excitation in the low-density region outside the cold plasmasphere (2). Local energy diffusion associated with wave scattering leads to the development of strongly enhanced phase space density just outside the plasmapause (3). Subsequently, radial diffusion can redistribute the accelerated electrons inwards or outwards from the developing peak (4).

The successful point-by-point inter comparison of radiation belt features observed by the Van Allen Probes with the predictions of the state of the art model developed by Richard Thorne and his group at UCLA dramatically demonstrates the significance of in situ particle acceleration within the Earth's radiation belts.

• December 2013: Although the Earth’s Van Allen radiation belts were discovered over 50 years ago, the dominant processes responsible for relativistic electron acceleration, transport and loss remain poorly understood. Evidence is presented for the action of coherent acceleration due to resonance with ultra-low frequency waves on a planetary scale. Data from the , launch July 25, 1990, and the Van Allen Probes mission (launch Aug. 12, 2012), with supporting modelling, collectively show coherent ultra-low frequency interactions which high energy resolution data reveals are far more common than either previously thought or observed. The observed modulations and energy-dependent spatial structure indicate a mode of action analogous to a geophysical synchrotron; this new mode of response represents a significant shift in known Van Allen radiation belt dynamics and structure. These periodic collisionless betatron acceleration processes also have applications in understanding the dynamics of, and periodic electromagnetic emissions from, distant plasma-astrophysical systems. 33)

The latest discovery uses measurements, taken by a UNH-led ECT (Energetic Particle, Composition, and Thermal Plasma) instrument on board the Van Allen Probes twin spacecraft, reveal that the high-energy particles populating the radiation belts can be accelerated to nearly the speed of light. This mode of action is analogous to that of a particle accelerator like the Large Hadron Collider. However, in this case, the Earth’s vast magnetic field, or magnetosphere, which contains the Van Allen belts, revs up drifting electrons to ever-higher speeds as they circle the planet from west to east. 34)

The recent finding comes on the heels of a related discovery — also made by the UNH-led EPIC (Energetic Particles and Ion Composition Experiment) instrument suite on CRRES in GTO — showing similar particle acceleration, but on a microscopic rather than a planetary scale. - Now, having the twin spacecraft in orbit with the Van Allen Probes and making simultaneous measurements in different regions of nearby space is a key part of the mission, allows the scientists to look at data separated in both space and time.


Figure 22: Artist’s conception of NASA’s Van Allen Probes twin spacecraft (image credit: Andy Kale, University of Alberta)

• July 2013: A team of scientists, led by the Los Alamos National Laboratory in New Mexico and involving the University of Colorado Boulder, have discovered a massive particle accelerator in the heart of one of the harshest regions of near-Earth space, a region of super-energetic, charged particles surrounding the globe, called the Van Allen radiation belts. Scientists knew that something in space accelerated particles in the radiation belts to more than 99 % of the speed of light, but they didn't know what that something was. New results from NASA's Van Allen Probes now show, that the acceleration energy comes from within the belts themselves. Particles inside the belts are sped up by local kicks of energy, buffeting the particles to ever faster speeds, much like a perfectly timed push on a moving swing. 35) 36)

The discovery that the particles are accelerated by a local energy source is akin to the discovery that hurricanes grow from a local energy source, such as a region of warm ocean water. In the case of the radiation belts, the source is a region of intense electromagnetic waves, tapping energy from other particles located in the same region. Knowing the location of the acceleration will help scientists improve space weather predictions, because changes in the radiation belts can be risky for satellites near Earth.


Figure 23: Artist's rendition of the Van Allen belts surrounding Earth - and an energy-acceleration graph that local particles experience in the belts (image credit: NASA, G. Reeves/M. Henderson)

• June 2013: In new NASA-funded research, the radiation belt group in the UCLA Department of Atmospheric and Oceanic Sciences explained the development of this third belt and its decay over a period of slightly more than four weeks. - By performing a "quantitative treatment of the scattering of relativistic electrons by electromagnetic whistler-mode waves inside the dense plasmasphere," the investigators were able to account for the "distinctively slow decay of the injected relativistic electron flux" and demonstrate why this unusual third radiation belt is observed only at energies above 2 MeV. 37)

• Feb. 2013: NASA's Van Allen Probes mission has discovered a previously unknown third radiation belt around Earth, revealing the existence of unexpected structures and processes within these hazardous regions of space. Previous observations of Earth's Van Allen belts have long documented two distinct regions of trapped radiation surrounding our planet. - The REPT (Relativistic Electron Proton Telescope) instruments aboard the twin Van Allen Probes quickly revealed to scientists the existence of this new, transient, third radiation belt. This discovery shows the dynamic and variable nature of the radiation belts and improves our understanding of how they respond to solar activity. 38) 39) 40)


Figure 24: Artist's rendition of the new model of the three radiation belt regions (image credit: NASA)

Legend to Figure 24: Two giant swaths of radiation, known as the Van Allen Belts, surrounding Earth were discovered in 1958. In 2012, observations from the Van Allen Probes showed that a third belt can sometimes appear. The radiation is shown here in yellow, with green representing the spaces between the belts.

• On Oct. 28, 2012, the Van Allen Probes/RBSP mission completed their 60-day commissioning phase, and began their two-year primary science mission. 41)

• In early September of 2012, scientists with the newly launched Van Allen Probes got permission to turn on one of their instruments after only three days in space instead of waiting for weeks (until the completion of the commissioning phase), as planned. They wanted to turn on the REPT (Relativistic Electron Proton Telescope) so that its observations would overlap with another mission, SAMPEX (Solar, Anomalous, and Magnetospheric Particle Explorer), that was soon going to de-orbit and reenter Earth’s atmosphere. 42)

Due to a stroke of luck, the dynamic nature of space weather provided the formation of the third radiation belt. This phenomenon was previously unknown. The scientist watched – in disbelief – while their data showed the extra belt forming, then suddenly disappear, like it had been cut away with a knife. They have not yet seen a recurrence of a third belt.

What happened is that shortly before the REPT instrument was turned on, solar activity on the sun had sent energy toward Earth that caused the radiation belts to swell. The energetic particles then settled into a new configuration, showing an extra, third belt extending out into space. — By the fifth day REPT was on, the project was able to plot out the observations and watch the formation of a third radiation belt (Figure 25). The third belt persisted beautifully, day after day, for four weeks - when a CME from the sun destroyed the third belt. The observations were made from institutions including; LASP (Laboratory for Atmospheric and Space Physics) at the University of Colorado, Boulder, CO; NASA/GSFC (Goddard Space Flight Center), Greenbelt, MD; LANL (Los Alamos National Laboratory) in Los Alamos, N.M.; and the Institute for the Study of Earth, Oceans, and Space at the University of New Hampshire in Durham (Ref. 38).


Figure 25: Energetic electron data gathered by the REPT instruments from Sept. 1, 2012 to Oct. 4, 2012 (horizontal axis), image credit: LASP, NASA (Ref. 10)

Legend to Figure 25: This graph shows three discrete energy channels (measured in MeV). The third belt region (in yellow) and second slot (in green) are highlighted, and exist up until a CME (Coronal Mass Ejection) destroys them on Oct. 1, 2012. The vertical axis in each is L*, effectively the distance in Earth radii at which a magnetic field line crosses the magnetic equatorial plane.

• Immediately after launch, RBSP entered a 60 day commissioning phase of operations, where all of the spacecrafts' systems and instruments are activated, monitored, and made ready for the two-year primary science mission. 43) 44)



Sensor complement: (ECT, EMFISIS, EFW, RBSPICE, RPS)

In the summer of 2006, NASA selected four university teams to provide experiments and supporting hardware for the RBSP spacecraft to study the near-Earth space radiation environment of the inner magnetosphere. The instruments are being developed by Boston University, University of Iowa, University of Minnesota, New Jersey Institute of Technology. In addition, NASA signed a memorandum of agreement with NRO (National Reconnaissance Office) to provide a fifth science investigation. 45) 46) 47)

Each RBSP spacecraft carries five instruments or instrument suites to perform in situ measurements of the ions and electrons, electric and magnetic fields, and electric and magnetic waves in the radiation belts.

Science instrument


Power allocation

Average data rate

ECT (Energetic Particle Composition and Thermal Plasma Suite)

66 kg

89.7 W

20.4 kbit/s

EMFISIS (Electric and Magnetic Field Instrument Suite)

27.4 kg

15.5 W

12 kbit/s

EFW (Electric Field and Waves)

9.2 kg

14.4 W

2 kbit/s

RBSPICE (RBSP Ion Composition Experiment)

6.6 kg

7.1 W

5.4 kbit/s

RPS (Relativistic Proton Spectrometer)

20.9 kg

22.5 W

32.2 kbit/s

Totals for science instruments

130.1 kg

149.2 W

72 kbit/s

Table 6: Science instrument accommodation parameters


Figure 26: The particle experiments of the RBSP mission (image credit: NASA, Ref. 3)


Figure 27: Artist's conception shows the Van Allen radiation belts (green), which are two doughnut-shaped (torus) regions full of high-energy particles that fill the near-space around Earth (image credit: NASA)

Legend to Figure 27: The blue and red lines between and around the belts depict the north and south polarity of the planet’s magnetic field. The inner belt, a blend of protons and electrons, can reach down as low as 1,000 km in altitude. The outer belt, comprised mainly of energetic electrons, can swell to as much as 60,000 km above Earth’s surface. Both rings extend to roughly ±65º in north and south latitude. 48)


ECT (Energetic Particle, Composition, and Thermal Plasma Suite):

The ECT instrumentation is being developed at Boston University, Boston MA, PI: Harlan E. Spence. The overall objective of ECT is to determine the spatial, temporal, and pitch angle distributions of electrons and ions over a broad and continuous range of energies. The coordinated ECT particle measurements, analyzed in combination with fields and waves observations and state-of-the-art theory and modeling, are necessary yet sufficient for understanding the acceleration, global distribution, and variability of radiation belt electrons and ions, key science objectives of the LWS program and the RBSP mission. 49)

The science objectives are: 50) 51)

- Determine the physical processes that produce radiation belt enhancements

- Determine the dominant mechanisms for relativistic electron loss

- Determine how the inner magnetospheric plasma environment controls radiation belt acceleration and loss

- Develop empirical and physical models for understanding and predicting radiation belt space weather effects.

The ECT suite consists of three highly-coordinated instruments (MagEIS, HOPE, and REPT) that cover comprehensively the full electron and ion spectra from one eV to 10’s of MeV with sufficient energy resolution, pitch angle coverage and resolution, and with composition measurements in the critical energy range up to 50 keV, and also from a few to 50 MeV/nucleon. All three instruments are based on measurement techniques proven in the radiation belts, optimized to provide unambiguous separation of ions and electrons and clean energy responses even in the presence of extreme penetrating background environments.

MagEIS (Magnetic Electron Ion Spectrometer): MagEIS uses magnetic focusing and pulse height analysis to provide the cleanest possible energetic electron measurements over the critical energy range of 30 keV to 4 MeV for electrons and 20 keV to 1 MeV for ions. Magnetic focusing is accomplished by magnets being used to deflect particles towards the senors to distinguish them from background radiation. MagEIS will provide the cleanest measurements of radiation belt electrons to date. A total of four MagEIS Instruments are carried aboard each spacecraft, each covering a particular part of the energy spectrum and a large range of pitch angles.

HOPE (Helium Oxygen Proton Electron): HOPE uses an electrostatic top-hat analyzer and time-gated coincidence detectors to measure electrons, protons, and helium and oxygen ions with energies from less than or equal to 20 eV or spacecraft potential (whichever is greater) to greater than or equal to 45 keV while rejecting penetrating backgrounds. Particles that are being measured by HOPE exist throughout the solar system and play a significant role in Van Allen Belt Dynamics because these slower particles generate electromagnetic waves which can affect higher energy particles trapped in the belts. The HOPE device is being provided by LANL (Los Alamos National Laboratory). 52)

REPT (Relativistic Electron Proton Telescope): REPT covers the challenging electron range of 4-10 MeV and proton energy range of 20-75 MeV to capture most intense events.


Figure 28: Illustration of the MagEIS (left) and HOPE (right) instruments (image credit: JHU/APL)


Figure 29: Illustration of the REPT instrument (image credit: JHU/APL)

The ECT Science Team is distributed across eight funded US institutions (The Aerospace Corporation, Boston University, Dartmouth College, Los Alamos National Laboratory (LANL), MIT, Southwest Research Institute, University of California Los Angeles, and University of Colorado), one unfunded US Government partner agency (NOAA/SEC) and three unfunded international collaborators (University of Alberta, British Antarctic Survey, and CERT-ONERA). The Science Team comprises radiation belt community leaders in areas of: instrument design and operation; science data analysis and management; theory and modeling; and space weather/radiation belt applications.


Figure 30: Overview of the ECT science team (image credit: BU)


EMFISIS (Electric and Magnetic Field Instrument Suite and Integrated Science)

The EMFISIS instrumentation is being developed at UI (University of Iowa), Iowa City, PI: Craig A. Kletzing. The objective of the investigation focuses on the important role played by magnetic fields and plasma waves in the processes of radiation belt particle acceleration and loss. EMFISIS offers the opportunity to understand the origin of important magnetospheric plasma waves as well as the evolution of the magnetic field that defines the basic coordinate system controlling the structure of the radiation belts and the storm-time ring current. 53)

The science goals are: 54)

- Differentiate among competing processes affecting the acceleration and transport of radiation particles

- Differentiate among competing processes affecting the precipitation and loss of radiation belt particles

- Quantify the relative contribution of adiabatic and non-adiabatic processes on energetic particles

- Understand the effects of the ring current and other storm phenomena on radiation electrons and ions

- Understand how and why the ring current and associated phenomena vary during storms.

The EMFISIS instrument suite consists of the following components:

• MAG (Tri-axial Magnetometer): MAG is a tri-axial fluxgate magnetometer: Vector B, DC-15 Hz, 0.1 nT accuracy, three sensors on rigid boom.

• Waves components:

- Magnetic field - vector B: 10 Hz - 12 kHz, sensitivity: 3 x 10-11 nT 2Hz-1 @ 1 kHz, 3 sensors on rigid boom

- Electric field - spin-plane E: 10 Hz- 12 kHz (vector), 10 kHz-400 kHz (single channel), sensitivity: 3 x 10-17 V 2m-2Hz-1 @ 1 kHz, shares booms with EFW (Electric Fields and Waves) instrument.


Figure 31: Boom-mounted components of EMFISIS (image credit: JHU/APL)


EFW (Electric Field and Waves Suite)

The EFW instrumentation is being developed at UMN (University of Minnesota), Minneapolis, PI: John R. Wygant. The EFW objective is to study the electric fields in near-Earth space that energize radiation particles and modify the structure of the inner magnetosphere.

This investigation consists of a set of four spin-plane electric field (E-field) antennae and a set of two spin-axis stacer (tubular, extendable) booms. The investigation will provide understanding of the electric fields associated with particle energization, scattering and transport, and the role of the large-scale convection electric field in modifying the structure of the inner magnetosphere.

The science objectives are: 55) 56)

- Energization by the large-scale convection E-field

- Energization by substorm injection fronts propagating in from the magnetotail

- Radial diffusion of energetic particles mediated by ULF (Ultra-Low Frequency) magnetohydrodynamic (MHD) waves

- Transport and energization by intense magnetosonic waves generated by interplanetary shock impacts upon the magnetosphere

- Coherent and stochastic acceleration and scattering of particles by small-scale, large-amplitude plasma structures, turbulence and waves (electromagnetic and electrostatic ion cyclotron waves, kinetic Alfven waves, solitary waves, electron phase space holes, zero frequency turbulence).


Figure 32: Components of the EFW instrument suite (image credit: JHU/APL)

Key measurements of EFW:

- Spin plane component of E at DC - 12 Hz (0.05 mV/m accuracy)

- Spin axis component of E at DC - 12 Hz (~3 mV/m accuracy)

- E- and B-field spectra for nearly-parallel and nearly-perpendicular to B components between 1 Hz and 12 kHz at 6-second cadence

- Spacecraft potential estimate covering cold plasma densities of 0.1 to ~100 cm-3 at 1 second cadence

- Burst recordings of high-frequency E- and B-field waveforms, as well as individual sensor potentials for interferometric analyses.

The EFW instrument measures the three dimensional electric field and cold plasma density estimates from the spacecraft potential all over a frequency range from DC to ~500 kHz. Measurements from the spatially separated spacecraft will provide information on azimuthal and radial spatial scales and propagation velocities of large scale structures. The spin plane electric field vector is obtained from spherical sensors at the ends of two pair of orthogonal booms with tip-to-tip separations of 80 and 100 m. The spin axis measurement is obtained from a pair of stace booms with a tip-to-tip separation of ~12 m. The electric field below 12 Hz is telemetered continuously while higher time resolution is obtained from a programmable burst memory with a maximum sampling rate for six quantities of ~ 30,000 samples/s. 57)

The high time resolution data includes interferometric timing measurements between individual probes at the ends of the booms which provide information on small scale structures and phase velocities. DC magnetic fields from the fluxgate magnetometer and wave magnetic fields from the search coil, both associated with the University of Iowa instrument are input into the EFW instrument for processing in the burst memory and in the Digital Signal Processing Board (DSP). The DSP provides wave spectra and cross spectra of electric and magnetic field and cold plasma density fluctuations. The EFW instrument also provides wave electric field signals to the University of Iowa EMFISIS instrument.


RBSPICE (Radiation Belt Storm Probes Ion Composition Experiment)

The RBSPICE instrumentation is being developed at NJIT (New Jersey Institute of Technology), Newark, NJ, PI: Louis J. Lanzerotti. The objective of RBSPICE is to determine how space weather creates what is called the “storm-time ring current” around Earth and determine how that ring current supplies and supports the creation of radiation populations.

The geomagnetic field drives relativistic electron motion via time-dependent gradient-curvature drift. Thus, structural variations of the inner magnetospheric field due to storm-time ring current growth control transport and losses in the outer belt.

This investigation will accurately measure the ring current pressure distribution, which is needed to understand how the inner magnetosphere changes during geomagnetic storms and how that storm environment supplies and supports the acceleration and loss processes involved in creating and sustaining hazardous radiation particle populations.

The science objectives are: 58)

- Understand the effects of the ring current and other storm phenomena on radiation electrons and ions

- Understand how and why the ring current and associated phenomena vary during storms

- Support development and validation of specification models of the radiation belts for solar cycle time scales.

Measurement requirements: Hot plasma pressure must be derived to calculate the ring current contribution to storm-time magnetic fields. Thus, it is necessary to resolve the full energy spectrum of the ring current as well as its composition (H, He, O).


Figure 33: Illustration of the RBSPICE instrument (image credit: JHU/APL)

The measurement quality is independent of the angle between the B-field and the spin axis.

- Ion composition energy range is low enough to determine the complete Ring Current energy density.

- High angle and energy resolution provide detailed pitch-angle and energy spectra: Δθ = 22.5º, ΔE/E = 0.1.


RPS (Relativistic Proton Spectrometer)

NASA entered into a partnership agreement with the NRO (National Reconnaissance Office) to provide the RPS instrument, PI: Clark Groves. - Note: the instrument is also referred to as PSBR (Proton Spectrometer Belt Research).

The objective of RPS is to measure the inner Van Allen belt protons with energies from 50 MeV to 2 GeV. Presently, the intensity of trapped protons with energies beyond about 150 MeV is not well known and thought to be underestimated in existing specification models. Such protons are known to pose a number of hazards to astronauts and spacecraft, including total ionizing dose, displacement damage, single event effects, and nuclear activation. This instrument will address a priority highly ranked by the scientific and technical community and will extend the measurement capability of this mission to a range beyond that originally planned. The project’s goal is the development of a new standard radiation model for spacecraft design.

The science objectives are: 59) 60)

- Support development of a new AP9/AE9 standard radiation model for spacecraft design

- AFRL to develop and test model for RBSP data in general and RPS specifically

- AP9 (protons) and AE9 (electrons) will provide standardized worst-case specifications: dose rate; internal charging/deep dielectric charging; surface charging (most intense fluxes in keV electrons).


Figure 34: Illustration of the RPS instrument (image credit: JHU/APL)

RPS measures energy spectra and angular distributions of protons from 50 MeV to 2 GeV (expect full inner-zone spatial distributions with better-than-weekly cadence):

- Energetic protons responsible for total dose in MEO for shielding thickness over 200 mils aluminum

- Protons responsible for displacement damage

Telescope consists of 8 silicon detectors and a Cherenkov detector:

- Stacked Si detectors used for 50 MeV to ~400 MeV, incident angle constrained by 8-fold coincidence

- Cherenkov detector used for > 400 MeV

- Absolute flux accuracy: dJ/J ~10%

- Energy resolution: dE/E ~30% @ 50 MeV, to 100% @ 2 GeV

- Angular resolution: 30º instantaneous, 5º de-convolved.



Ground segment:

Mission operations: The mission operations concept is designed for mostly unattended spacecraft operations, with distributed science operations. Because the spacecraft are spin-stabilized and nominally sun-pointing, they are inherently in a safe state and the need for constant monitoring is minimized. All critical activities – including commissioning activities and all propulsive maneuvers – are performed in contact with the ground, but nominal science operations are not constrained to occur during “staffed” periods of time. Similarly, the instrument operations are performed “offline” of the MOC (Mission Operations Center), and instrument commands are queued up at the MOC remotely from the SOCs (Science Operations Centers), and uploaded to the spacecraft during the next regular contact. This approach of unattended, decoupled operations greatly reduces the cost of the operational phase of the mission, and it has been successfully demonstrated at APL on the STEREO mission.

The MOC is located at APL, and it serves as the central hub through which all commands and telemetry flow. Figure 5 shows the distributed nature of the operations architecture. APL’s SCF (Satellite Control Facility), with its 18 m antenna, serves as the primary ground station for the mission. Supplemental contacts using the NEN (Near Earth Network) will be used to ensure sufficient daily data download to maintain data volume margin on the spacecraft solid-state recorders. The spacecraft will also utilize the TDRSS system for communications shortly after launch and during early operations, and also rely on TDRSS for emergency communications (Ref. 14).


Figure 35: MOC and SOC architecture of the RBSP distributed ground system (image credit: JHU/APL)

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