Minimize PSP

Parker Solar Probe — former SPP (Solar Probe Plus)

Spacecraft     Project Status     Launch    Sensor Complement    References

The Solar Probe Plus mission is part of NASA's LWS (Living With a Star) Program. The program is designed to understand aspects of the sun and Earth's space environment that affect life and society. The program is managed by NASA/GSFC (Goddard Space Flight Center). The Johns Hopkins University Applied Physics Laboratory (JHU/APL) in Laurel, MD., is the prime contractor for the spacecraft. In September 2010, NASA selected the Solar Probe Plus mission for development. A launch of the mission is planned for 2018. 1)

NASA's first mission to go to the sun, the Parker Solar Probe, is named after Eugene Parker who first theorized that the sun constantly sends out a flow of particles and energy called the solar wind.

NASA has renamed the Solar Probe Plus spacecraft humanity's first mission to a star, which will launch in 2018 as the Parker Solar Probe in honor of astrophysicist Eugene Parker. The announcement was made at a ceremony at the University of Chicago, where Parker serves as the S. Chandrasekhar Distinguished Service Professor Emeritus, Department of Astronomy and Astrophysics.

In 1958, Parker then a young professor at the university's Enrico Fermi Institute published an article in the Astrophysical Journal called "Dynamics of the interplanetary gas and magnetic fields." Parker believed there was high speed matter and magnetism constantly escaping the sun, and that it affected the planets and space throughout our solar system.

This phenomenon, now known as the solar wind, has been proven to exist repeatedly through direct observation. Parker's work forms the basis for much of our understanding about how stars interact with the worlds that orbit them.

"This is the first time NASA has named a spacecraft for a living individual," said Thomas Zurbuchen, associate administrator for NASA's Science Mission Directorate in Washington. "It's a testament to the importance of his body of work, founding a new field of science that also inspired my own research and many important science questions NASA continues to study and further understand every day. I'm very excited to be personally involved honoring a great man and his unprecedented legacy."

Table 1: On May 31, 2017, NASA renamed the Solar Probe Mission to honor pioneering physicist Eugene Parker 2)

The SPP science objectives are: 3) 4) 5)

1) Determine the structure and dynamics of the magnetic fields at the sources of the fast and slow solar wind.

2) Trace the flow of energy that heats the corona and accelerates the solar wind.

3) Determine what mechanisms accelerate and transport energetic particles.

4) Explore dusty plasma phenomena in the near-sun environment and their influence on the solar wind and energetic particle formation.

Background: 6) 7) 8) 9) 10)

• The concept for a "solar probe" dates back to "Simpson's CommiIee" of the Space Science Board (National Academy of Sciences, 24 October 1958). The need for extraordinary knowledge of Sun from remote observations, theory, and modeling to answer the questions:

- Why is the solar corona so much hotter than the photosphere?

- How is the solar wind accelerated?

• SPP was a NASA concept study in 2008. The challanging objective of the mission is to explore the near-Sun environment for a better understanding of solar physics. So far, no missions have penetrated closer to the Sun than 0.3 AU (Astronomical Units).

• Helios 1 and 2 were a pair of cooperative US and German deep space probes (launch Dec. 10, 1974 and Jan. 15, 1976, respectively) which set the record for the closest approach to the Sun, at ~45 million km, slightly inside the orbit of Mercury.

• The NASA MESSENGER mission (launch Aug. 3, 2004) was the first spacecraft to orbit planet Mercury. The data of the Sun are unique representing the only in situ measurements of the inner heliosphere as close as 60 solar radii (RS). The unexplored region within this distance is where the corona is accelerated to form the supersonic solar wind, and is critical to our understanding of the Sun's impact on the solar system.

First definitions of Solar Probe missions (studies) at NASA/JPL were started in 1978. The original Solar Probe mission concept of 2005, based on a Jupiter gravity assist trajectory, was no longer feasible under the new guidelines given to the mission. A complete redesign of the mission was required to meet the mission constraints, which called for the development of alternative mission trajectories that excluded a flyby of Jupiter.

In mid-2007, NASA asked JHU/APL to consider another concept for Solar Probe that would perform all science objectives of the 2005 concept, implemented as a non-nuclear powered spacecraft, and executed under a New Frontiers-like cost cap. The resulting mission is called Solar Probe+ in recognition of the potential gains in science of the current concept over predecessors. 11) 12) 13)

In March 2012, the SPP project advanced to Phase-B. 14)

Two key technical challenges make a solar probe much more difficult than other missions: 15)

1) The extremely high temperature and harsh environment in the Sun's proximity, which the spacecraft cannot survive without adequate thermal protection

2) The extreme difficulty of getting close to the Sun, as an enormous amount of velocity must be canceled out from the Earth orbital velocity in order for a probe to get close to the Sun.

SPP will sample the solar corona to reveal how it is heated and the solar wind and solar energetic particles are accelerated. Solving these problems has been a top science goal for over 50 years. 16) During the seven-year mission, seven Venus gravity assist (VGA) maneuvers will gradually lower the perihelia to <10 RS (Radius of sun ~700,000 km), the closest any spacecraft has come to the Sun. Throughout the 7-year nominal mission duration, the spacecraft will spend a total of 937 hours inside 20 RS , 440 hours inside 15 RS , and 14 hours inside 10 RS, sampling the solar wind in all its modalities (slow, fast, and transient) as it evolves with rising solar activity toward an increasingly complex structure. SPP will orbit the Sun in the ecliptic plane, and so will not sample the fast wind directly above the Sun's polar regions (Figure 1). However, the current mission design compensates for the lack of in-situ measurements of the fast wind above the polar regions by the relatively long time SPP spends inside 20 RS.17) - This will allow extended measurement of the equatorial extensions of high-latitude coronal holes and equatorial coronal holes. At a helioradius ~35 RS , there are two periods per orbit (one inbound and one outbound) when SPP will be in quasi-corotation with the Sun and will cross a given longitudinal sector slowly. In these intervals, known as fast radial scans, the spacecraft will sample the solar wind over large radial distances within a given flux tube before moving across the sector. These measurements will yield additional information on the spatial/temporal dependence of structures in the solar wind and on how they merge in the inner heliosphere.

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Figure 1: Solar wind speed as a function of heliographic latitude illustrating the relationship between the structure of the solar wind and coronal structure at solar minimum (a, c) and solar maximum (b). Ulysses SWOOPS solar wind data are superposed on composite solar images obtained with the SOHO EIT and LASCO C2 instruments and with the Mauna Loa K-coronameter. (d) Solar cycle evolution (image credit: D. J. McComas et al.) .18)

Science overview:

The SPP mission targets processes and dynamics that characterize the Sun's expanding corona and solar wind. SPP will explore the inner region of the heliosphere through in-situ and remote sensing observations of the magnetic field, plasma, and energetic particles. The solar magnetic field plays a defining role in forming and structuring the solar corona and the heliosphere. In the corona, closed magnetic field lines confine the hot plasma in loops, while open magnetic field lines guide the solar wind expansion in the inner corona. The energy that heats the corona and drives the wind derives from photospheric motions, and is channeled, stored, and dissipated by the magnetic fields that emerge from the convection zone and expand in the corona where they dominate almost all physical processes therein. Examples of these are waves and instabilities, magnetic reconnection, and turbulence, which operate on a vast range of spatial and temporal scales. Magnetic fields play also a critical role in coronal heating and solar wind acceleration. They are conduits for waves, store energy, and propel plasma into the heliosphere through complex forms of magnetic activity [e.g., CMEs (Coronal Mass Ejections), flares, and small-scale features such as spicules and jets]. How solar convective energy couples to magnetic fields to produce the multifaceted heliosphere is central to SPP science.

SPP will make in-situ and remote measurements from <10 RS to at least 0.25 AU (53.7 RS ). Measurements of the region where the solar wind originates and where the most hazardous solar energetic particles are energized will improve our ability to characterize and forecast the radiation environment of the inner heliosphere. SPP will measure local particle distribution functions, density and velocity field fluctuations, and electromagnetic fields within 0.25 AU of the Sun. These data will help answer the basic questions of how the solar corona is powered, how the energy is channeled into the kinetics of particle distribution functions in the solar corona and wind, and how such processes relate to the turbulence and wave-particle dynamics observed in the heliosphere. Cross-correlation of velocity, density, and electromagnetic fluctuations will allow a partial separation of spatial and temporal effects.

The physical conditions of the region below 20 RS are important in determining largescale properties such as solar wind angular momentum loss and global heliospheric structure. The Alfvénic critical surface, where the solar wind speed overtakes the Alfvén speed, is believed to lie in this region. This surface defines the point beyond which the plasma ceases to corotate with the Sun, i.e., where the magnetic field loses its rigidity to the plasma. In this region solar wind physics changes because of the multi-directionality of wave propagation (waves moving sunward and anti-sunward can affect the local dynamics including the turbulent evolution, heating and acceleration of the plasma). This is also the region where velocity gradients between the fast and slow speed streams develop, forming the initial conditions for the formation, further out, of CIRs (Corotating Interaction Regions).

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Figure 2: SPP, shown along its orbit (dashed curve) near a perihelion pass, will measure solar energetic ions and electrons from a vantage point very near the site where these particles are accelerated. The illustration sketches the occurrence of a solar flare and a CME extending a few RS from the Sun. The shock at the front edge of the CME and the compressed sheath plasma behind the shock form as the CME, with its entrained flux rope (tangled pink lines), pushes outward from the Sun through the ambient solar wind. Swept-up magnetic field lines are refracted and compressed across the shock and draped around the CME. Energetic particles accelerated at both the flare and CME shock are shown spiraling away from the Sun (yellow spirals) along the magnetic field. For simplicity, magnetic field lines around the shock are depicted as smooth. However, it is expected that the field ahead of CME shock and in the sheath will highly structured.Waves ahead of the shock that are produced by high intensities of shock-accelerated ions streaming away from the shock are sketched for the uppermost magnetic field line connected to the CME shock (image credit: Ref.16)

SPP participation:

• 31 institutions participate in SPP science teams

- 23 in the US, 8 foreign

- 17 educational, 5 non-profit, 8 government labs

• 106 science team members

- 69 PIs and Co-Is

- 37 additinal scientists

- Next generation graduate students and post-docs.

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Figure 3: Participating organizations in SPP (image credit: JHU/APL)

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Figure 4: Artist's view of the SPP spacecraft with solar panels folded into the shadows of its protective shield, gathers data on its approach to the Sun (image credit: JHU/APL)

To accomplish the science objectives of addressing the fundamental questions about the Sun by acquiring critical data and measurements to answer questions that cannot be answered by observations from satellites in Earth orbit and from other interplanetary space probes, a solar probe must approach the Sun closely. A solar orbit approach to within the range of 10 solar radii (Rs) from Sun's center must be considered to conduct the necessary in situ measurements and investigations.

Getting directly to the Sun from Earth would require a launch energy C3 as large as 423 km2/s2. This is beyond the capability of launch vehicles currently available (Atlas V, Delta IV Heavy) or to be developed in the near future. The highest launch C3 ever achieved was 164 km2/s2 for the New Horizons mission to Pluto (launch Jan. 2006).

After an extensive analysis by NASA and JHU/APL, the trajectory option 5 was chosen as the baseline trajectory for the new solar probe. The redesigned mission is named SPP (Solar Probe Plus) for its significant advantages in both technical implementation and science accomplishments as compared with the original Solar Probe mission.

The mission design utilizes seven Venus gravity assists to gradually reduce perihelion (Rp) from 35 solar radii (Rs) in the first orbit to < 10 Rs for the final three orbits. The SPP orbit consists of two primary orbit phases, a science phase (0.25AU to perihelion) and a cruise/data downlink phase (0.25AU to aphelion).

Parameter

Solar Probe (2005)

Solar Probe Plus (2008)

Minimum perihelion

4 Rs

9.5 Rs

Inclination

90º from ecliptic

3.4º from ecliptic

Number of solar passes

2

24

Total time within 20 Rs

96 hours

961 hours

Time between passes

4.6 years

88 to 150 days

Time from launch to first perihelion

4.1 years

3 months

Mission duration

8.8 years

6.9 years

Aphelion

5.5 AU

1 AU

Table 2: Comparison of Solar Probe and Solar Probe Plus (Ref. 15)

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Figure 5: Reference Mission: Launch and Mission Design Overview (image credit: JHU/APL, NASA)

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Figure 6: Detail of solar encounter timeline for a typical orbit (image credit: NASA, JHU/APL)

Figure 6 shows the orbit within ± 10 days of perihelion, and an expansion of the region ± 20 Rs. This figure also shows the time spent in each part of the solar encounter of scientific interest for one of the final orbits. In total, SPP will spend more than 2100 hrs closer than 30 Rs, nearly 1000 hours below 20 Rs, and 27 hrs in the region below 10 Rs.

SPP is an ambitious mission, requiring significant technology development in several major areas. Table 3 is a summary of the technology readiness assessment for SPP and gives an indication of the basis for technology. For each area, technology development plans have been established, and in each case, significant progress has been made to achieve TRL (Technology Readiness Level) 6 by PDR (Preliminary Design Review).

Item

TRL (Technology Readiness Level)

Comment

TPS (Thermal Protection System)

4

NASA sponsored technology development

Solar array

4

Combines space heritage and concentrator cells

Cooling system

4

Adapted from heritage systems

X-/Ka-band transponder

4

NASA sponsored technology development

LEON3 processor

5

Qualified product in new JHU/APL application

Table 3: Technology development for SPP (shows only items with TRL lower than 6)

 



Spacecraft:

At 9.5 Rs, the solar intensity is 512 times that at 1AU. SPP is packaged behind the carbon-carbon TPS (Thermal Protection System), a 11 cm thick heqat shield, to protect it from this extreme solar environment and allow it to operate at standard space thermal environments while the TPS experiences temperatures of 1400ºC on its sun-facing surface. SPP utilizes actively cooled solar arrays for power generation maintaining the solar cells within required temperature limits (Ref. 3). 19) 20)

Solar Probe Plus is a 3-axis stabilized spacecraft, shown in Figure 7, with functional block diagram in Figure 9.

TPS: The most prominent feature is the 2.3 m diameter TPS, with associated structure used to attach the shield to the spacecraft. The TPS protects the bus and payload within its umbra during solar encounter. The conceptual science instruments are mounted either directly to the bus, on a stand-off bracket near the fairing attachment, or on a science boom extended from the rear of the spacecraft.

In general, the payload is protected from the effects of solar exposure by the TPS. Two notable exceptions are the SPC (Solar Probe Cup), part of the SWEAP investigation, and the electric field antennas carried as part of the FIELDS investigation. Both sensor packages extend beyond the TPS and see the same environment as the TPS sunward-looking face. Both sensors are of high heritage; however the solar environment during solar encounter is significantly more severe than all previous experience. Therefore, technology development programs for each have been implemented to demonstrate the operation of each in the expected SPP environment.

Three deployable conceptual carbon-carbon plasma wave antennas are mounted 120º apart on the side of the bus. These antennas will partially protrude beyond the umbra during encounter. The solar array cooling system dissipates the high solar flux absorbed by solar array wings during closest approach to the sun enabling the solar cells to operate within their temperature constraints while providing the required electrical power. Water in the cooling system is pumped from the outboard-most edge of the solar array substrate, or platen, up through channels in the solar array wings into the four cooling system radiators mounted under the TPS and back through the pump located on the top deck of the spacecraft. The system can dissipate 6000 W of heat at perihelion, and is designed and operated to prevent freezing at aphelion.

The new configuration uses a single pair of arrays to generate power. The bulk of the solar array panel is filled with "primary cells" similar to cells used on the MESSENGER mission to Mercury, while the angled panel on the end of the solar arrays use cells designed to withstand the high illumination during perihelion.

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Figure 7: Illustration of the Solar Probe Plus spacecraft configuration (image credit: JHU/APL)

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Figure 8: Spacecraft overview (image credit: JHU/APL, Ref. 10)

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Figure 9: Block diagram of the Solar Probe Plus spacecraft (image credit: JHU/APL, Ref. 4)

At aphelion, the entire array is exposed to sunlight. As the spacecraft nears the sun, the array is tilted toward the spacecraft body until at perihelion only the end of the array is exposed to sunlight in the penumbra created by the TPS knife edge. The array substrate is a titanium plate with channels running under the cells. Water pumped through this panel carries heat to radiators mounted on the TPS support structure. Figure 10 shows the configuration at perihelion.

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Figure 10: Solar array illumination at perihelion (image credit: JHU/APL)

Technology development work on TPS has resulted in several changes to the design. The TPS remains a carbon-carbon and carbon foam sandwich, with a ceramic coating on the Sun-facing surface to control reflectance and emissivity properties. The shape and size of the TPS has changed to optimize mass while considering manufacturability and the need for longer knife edges for illumination control of the solar arrays. In particular, the TPS has shrunk from nearly 3 m in diameter to reflect more efficient packaging of the spacecraft.

The spacecraft in Figure 7 also reflects the new antenna configuration, including a 0.6 m HGA (High Gain Antenna) mounted on the body of the spacecraft instead of a boom. In addition to mass optimization, this change removes the need to deploy and retract the HGA each orbit to protect it from thermal damage at perihelion, thus increasing the reliability of the system.

The design uses a blowdown monopropellant hydrazine system for propulsion, with thrusters for attitude control and trajectory correction. Star Trackers and an internally redundant IRU (Inertial Measurement Unit) are included for guidance and control. The avionics suite is based on the APL IEM (Integrated Electronics Module) and PDU (Power Distribution Unit) used in most APL missions over the last decade or more. The IEM houses the command and data handling processor, solid-state recorder, interface to the guidance and control instruments, and payload interface. The PDU is an internally redundant box that includes all power switching. RIO (Remote I/O) devices are used to collect spacecraft telemetry, and communicate with the avionics suite through serial data links.

Avionics and SpaceWire Network: 21)

• SpaceWire selected over 1553: SpaceWire offers greater bandwidth and lower emissions

• Redundant processor module: (prime, hot spare, warm spare)

• Redundant electronic modules: SSRs (Solid State Recorders) are cross strapped

• Two cross strapped transponders.

 

Communication Coverage Profile:

Adequate communication links between ground and spacecraft are essential for mission operations. Transmissions of spacecraft operation commands, spacecraft telemetry, science observation sequences, and instrument measurement data between the SPP spacecraft and ground are through the spacecraft Telecomm system and NASA's DSN (Deep Space Network) of tracking stations located at Goldstone in the United States, Canberra in Australia, and Madrid in Spain. Besides the data transmission, navigation of the SPP spacecraft will rely on regular and periodic tracking of the spacecraft through the DSN. The communication coverage of the spacecraft over the mission duration directly affects the spacecraft's operation, science data download, navigation, and the control of the flight trajectory. Due to launch and navigation errors, the flight trajectory needs to be periodically adjusted by applying a TCM (Trajectory Correct Maneuver). Availability of adequate navigation tracking and communication links to the spacecraft dictates the placement of the trajectory correction maneuvers, which has direct impacts on the onboard ΔV budget (Ref. 19).

A comprehensive study of detailed communication coverage over the entire mission was conducted in Phase B across multiple SPP subsystem teams. Because of the unique operation environment of the SPP mission, many factors must be understood in order to maintain adequate communications with the spacecraft. First, the highly elliptical solar orbits across the inner solar system cause frequent solar conjunctions, sometimes with extended periods. And secondly, the spacecraft's TPS obstructs the view of the antenna and causes extra outage of communication times.

The X-band is baselined for spacecraft tracking for navigation and works for both uplink and downlink modes. The Ka-band is mainly for science data downlink and works only for the downlink mode.

Besides the communication outage due to the solar conjunctions attributed to the viewing geometry of Sun, Earth, and the spacecraft, the TPS of the spacecraft sometimes causes additional outage. Because of the extremely high heat radiated from the Sun, the spacecraft bus must be constantly protected from direct solar radiation to prevent overheating. When spacecraft solar distance is less than 0.7 AU, the spacecraft must be oriented with the TPS pointed at the Sun, so that the spacecraft bus and components including the antennas are behind the TPS and are protected inside the TPS umbra. Since the antennas are behind the TPS, some of the radio transmission is obstructed by the TPS. About 14° of the field of view from the center of the TPS is blocked. When the direction of Earth is near the direction of the Sun and within the 14° cone angle about the TPS center, the view from the SPP antenna to Earth is obstructed by the TPS, thereby preventing communication between Earth and the spacecraft.

The survive the extreme solar radiation conditions, the TPS must remain pointed toward the sun at all times. The flight software is required to be capable of controlling attitude within 5 seconds of a processor reset or demotion. The spacecraft has three flight processors (prime, hot spare, and backup spare) to meet this requirement. The tight TPS pointing requirements cause geometric challenges for communications with earth resulting in severely limited communication availability and bandwidth. The SPP spacecraft will use Ka-band downlink transmissions which provide high throughput with CFDP (CCSDS File Delivery Protocol) to return as much data as possible. The SPP spacecraft will use X-band uplink with CFDP to provide efficient guaranteed delivery of commands and save uplink bandwidth when deploying command loads to multiple processors. 22)

Commanding: The SPP flight software reuses heritage code from JHU/APL missions designed to use CCSDS Telecommand packets for commanding. The SPP Ground Software has a database of commands which can create telecommand packets and package them into CLTUs (Command Link Transmission Units). SPP supports the unreliable delivery Expedited Service (BD Service) of the CCSDS Communications Operations Procedure-1 (COP-1) commanding protocol. SPP does not support the reliable Sequence-Controlled Service (AD Service) of COP-1 commanding. The COP-1 AD Service is not well suited for deep space without modification as it provides a limited maximum number of commands without acknowledgment and requires significant retransmission if a single command is dropped.

SPP is a decoupled mission where each SOC (Science Operations Center) can command their instrument as they see fit. Aside from a limited set of calibration activities and earth pointing for communications (when allowed), the spacecraft pointing is fixed at the sun. There is no coordination required between the SOCs and the MOC (Mission Operations Center) to point the spacecraft. The MOC validates that instrument commands are well-formed, targeted to the right destination, and have an APID (Application Identifier) within the assigned instrument APID range prior to allowing transmission to the spacecraft. The MOC does not perform any further validation of instrument commands. The flight software will only send instrument CF contents to the target instrument interface. Instruments can only be commanded via files sent to the MOC by SFTP. These command files are queued and later uplinked to the spacecraft. A separate sequence number will be used for each instrument interface. This guarantees the ordering of files sent to the instrument interface while not impacting sequencing of CDH (Command and Data Handling) or other instrument command files. Due to power constraints, instruments are off during Ka telemetry downlinks, but files can still be uplinked during this period and later streamed to the instrument when it is powered on. Figure 11 highlights the steps involved in sending instrument commands to the target instrument.

The MOC runs a file queue management application that is responsible for initiating the uplink file transfers. The spacecraft CDH and instrument files are all stored in separate queues in this application. Instrument files have an enable time when it is considered acceptable to send them to the spacecraft and a time-out time when an opportunity would have been missed and it no longer makes sense to uplink the file. MOC files are queued in realtime by a contact plan and do not have time out times. Each queue can be enabled for file selection or disabled by the MOC. This application will select the next file by checking for priority files first and then doing a round-robin selection between each enabled command interface with a file that is ready to send.

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Figure 11: Instrument command flow (image credit: JHU/APL)

Telemetry: JHU/APL uses the SLE (Space Link Extension)) Return All Frames (RAF) service to receive CCSDS telemetry frames from the spacecraft. Virtual channels are assigned for real-time telemetry, recorded data on SSR (Solid State Recorder), and realtime fill frames. The process of prioritizing and playing back SSR telemetry via CFDP has been used quite successfully on the MESSENGER and Van Allen Probes missions. SSR Housekeeping telemetry is ingested into the MOC telemetry archive. Instrument SSR telemetry files will be provided directly to SOCs with no processing by the MOC.

SPP will record telemetry immediately before a low data rate contact and use CFDP to guarantee delivery of the most critical data during this contact. Figure 12 shows the high level flow of instrument telemetry from creation to the SOC.

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Figure 12: Instrument telemetry flow (image credit: JHU/APL)

Frontier Radio on SPP mission: 23)

NASA has selected the Frontier Radio DS (Deep Space) version, developed by JHU/APL, for the communication for the SPP (Solar Probe Plus) mission. The VAP (Van Allen Probes) mission successfully transitioned the Frontier Radio technology to TRL-9 in an S-band duplex configuration for Near-Earth applications (Frontier NE). The successful VAP effort and TRL-6 X/X/Ka-band development efforts provided a deep space Frontier Radio (Frontier DS) with high heritage from the TRL-9 near-Earth unit. The low-SWaP (Size, Weight, and Power) and intrinsically high radiation tolerance of the Frontier Radio DS uniquely qualified it for the SPP application and resulted in the mission baselining this radio. As with VAP for the near-Earth radio, the SPP effort supported the maturation of the deep space radio enhancements, including the necessary compatibility testing with the DSN (Deep Space Network). Flight Frontier Radios for the SPP mission (Figure 13) have completed qualification as of August 2016 and will be integrated into the spacecraft during the remainder of 2016. 24)

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Figure 13: A flight Frontier Radio for SPP (image credit: JHU/APL)

 

 

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Figure 14: Configuration of the Avionics and SpaceWire Network (image credit: JHU/APL)

Status of the project:

• July 14, 2016: Following a successful NASA management review on July 7, the Solar Probe Plus mission — which will send a spacecraft on several daring data-collecting runs through the sun's atmosphere — is moving into the system assembly, integration, test and launch stage of the project. NASA terms this period as Phase D, during which the mission team will finish building the spacecraft, install its science instruments, test it under simulated launch and space conditions, and launch it. 25)

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Figure 15: Engineers at JHU/APL in Laurel, Maryland, prepare the developing Solar Probe Plus spacecraft for thermal vacuum tests that simulate conditions in space. Today, the spacecraft includes the primary structure and its propulsion system; still to be installed over the next several months are critical systems such as power, communications and thermal protection, as well as science instruments. The probe is scheduled for launch in summer 2018 (image credit: NASA/JHUAP)

• April 8, 2015: NASA's SPP (Solar Probe Plus) mission reached a major milestone in March when it successfully completed its CDR (Critical Design Review). An independent NASA review board met at the Johns Hopkins University Applied Physics Laboratory (APL) in Laurel, Maryland, from March 16 to 20 to review all aspects of the mission plan; APL has designed and will build and operate the spacecraft for NASA. The CDR certifies that the Solar Probe Plus mission design is at an advanced stage and that fabrication, assembly, integration and testing of the many elements of the mission may proceed. 26)

• In March 2014, Solar Probe Plus will begin advanced design, development and testing — a step NASA designates as Phase C — following a successful design review in which an independent assessment board deemed that the mission team, led by JHU/APL ( Johns Hopkins University/Applied Physics Laboratory) in Laurel, MD, was ready to move ahead with full-scale spacecraft fabrication, assembly, integration and testing. 27)

Launch: A launch of the Solar Probe Plus spacecraft is scheduled for the summer of 2018 from Complex 37 of the Cape Canaveral Air Force Station, Florida. The launch vehicle is a Delta-4 Heavy rocket of ULA (United Launch Alliance), augmented by Orbital ATK's Star-48 solid motor as a third stage, in order to cope with the extremely high energy required for this flagship mission. 28)

Orbit: The trajectory able to send the spacecraft within 10 RS (Solar Radii) of the Sun center is a Venus-Venus-Venus-Venus-Venus-Venus-Venus-Gravity-Assist (V7GA) trajectory, a unique trajectory developed to enable the Solar Probe Plus mission without a Jupiter gravity assist. Even with the most powerful launch vehicle and upper stage, a spacecraft cannot get close to the Sun from Earth directly. Extra energy must be shed off the spacecraft's orbit to further reduce its heliocentric orbital velocity in order to encounter the Sun under 10 RS. The V7GA trajectory allows for the spacecraft to reduce the necessary orbital speed via multiple Venus gravity assists.

The amount of required orbital speed reduction required at aphelion is too large to come from one or two Venus flybys. Attaining the aphelion orbital speed reduction will require seven Venus flybys. Following each Venus flyby, the orbital speed at aphelion will decrease, resulting in a smaller orbit with a shorter perihelion distance. After seven Venus flybys, orbit perihelion distance will gradually decrease to 9.86 RS, the minimum solar distance required for the baseline mission. Throughout the mission there are no additional deep space maneuvers; all the orbit changes as well as the phasing (Venus-to-Venus transfer location and timing) between each Venus flyby are achieved through the control of the Venus flybys by appropriate selection of the Venus flyby target parameters. To minimize the mission duration, both resonant and non-resonant Venus flybys are utilized in this trajectory design (Ref. 19).

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Figure 16: Overview of the V7GA mission trajectory (image credit: JHU/APL)

The SPP mission is comprised of 24 highly elliptical, heliocentric orbits with decreasing orbital periods from 168 days for orbit 1, settling into an 88 day orbit period midway through the mission. Each orbit is broken into two distinct periods, the Solar Encounter period and the Cruise/Downlink period. Figure 17 highlights the primary characteristics of each period (Ref.22) .

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Figure 17: SPP Orbital Operations Concept (image credit: JHU/APL)


 
Minimize Sensor complement: (SWEAP, WISPR, ISIS-EPI, FIELDS)


Sensor complement: (SWEAP, WISPR, ISIS-EPI, FIELDS)

The experiments selected for Solar Probe Plus are specifically designed to solve two key questions of solar physics:

1) Why is the Sun's outer atmosphere so much hotter than the Sun's visible surface?

2) What propels the solar wind that affects Earth and our solar system?

The answers to these questions can be obtained only through in-situ measurements down in the corona. The solar physics community has been struggling with these questions for decades. There is hope that this mission should finally provide those answers.

In 2010 the Heliophysics Division of the Science Mission Directorate at NASA conducted a competitive "Announcement of Opportunity" procurement and in September of that year selected five science investigation teams to provide instruments and scientific study for the Solar Probe Plus Mission. These teams are: 29) 30)

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Figure 18: Instrument accommodations of the reference vehicle, ram facing view (image credit: JHU/APL)

 

SWEAP (Solar Wind Electrons Alphas and Protons)

The SWEAP instrument is being developed at the Smithsonian Astrophysical Observatory (SAO) in Cambridge, MA, USA, PI: Justin Kasper. The instruments that make up the SWEAP investigation, are the SPC (Solar Probe Cup), and the SPAN (Solar Probe Analyzers). SPC is a Faraday Cup that looks around the spacecraft heat shield directly at the Sun and makes rapid measurements of the bulk properties of the solar wind. SPAN is a set of three ESAs (Electrostatic Analyzers) behind the spacecraft heat shield that make detailed measurements of the three dimensional velocity distribution functions of ions and electrons. 31) 32)

SPAN is divided into two modules on either side of the spacecraft bus. SPAN-A is on the ram side of the spacecraft and has an ion and electron ESA. SPAN-B is on the anti-ram side of the spacecraft and has an ESA. SPAN-A, SPAN-B and SPC measurements include:

• Electron pitch angle distributions, moments, ion moments

• Full 3D electron velocity distribution functions

• Full 3D ion velocity distribution functions

• Proton bulk properties

• Ion 2D VDFS

• Ion/Electron total flux and flow angles at high cadence.

The advantages of the SWEAP suite include the following:

• By including SPC, the suite is able to observe the solar wind flow throughout each encounter. This significantly increases the probability of observing significant events within the inner heliosphere such as interplanetary shocks and coronal mass ejections.

• SPC is also able to make burst measurements of the flow angles of the solar wind faster than 100Hz, enabling measurements of ion fluctuations above the cyclotron frequency

• The electron fields of view of SPAN-A and SPAN-B cover most of the sky, allowing SWEAP to detect energetic beams of electrons flowing along the local magnetic field, regardless of its orientation. This allows us to identify times when both ends of the magnetic field line passing through the spacecraft are connected to the solar surface.

• An onboard interface between SWEAP and FIELDS permits generation of 2D VDFs, triggered bursts, and wave-particle correlation functions.

 

SWEAP Overview:

The overall organizational structure of the SWEAP Investigation at the time of the PDR (Preliminary Design Review). The investigation is managed by SAO (Smithsonian Astrophysical Observatory). In addition to the management of the suite, SAO is also responsible for the SWEAP Science Operations Center (SOC) and the Solar Probe Cup (SPC). The UCB/SSL (University of California, Berkeley /Space Sciences Laboratory) is responsible for the SPAN (Solar Probe Analyzers) and SWEM (SWEAP Electronics Module), which controls the instruments, distributes power, formats SWEAP data products, and serves as the single interface to the spacecraft. Additional science team members are from institutions in the United States and France including the University of Michigan, NASA Marshall Space Flight Center, NASA Goddard Space Flight Center, University of Alabama Huntsville, Los Alamos National Laboratory, Massachusetts Institute of Technology, and the University of New Hampshire.

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Figure 19: Placement of the SWEAP instruments (highlighted in yellow) on the Solar Probe Plus spacecraft. This view is from the perspective of an observer North of the ecliptic plane looking down on the spacecraft. In this view the spacecraft is oriented with the Sun-facing thermal protection system (TPS) heat shield on the right, the ram-facing (direction of motion about the Sun) side of the spacecraft looking down, and the anti-ram side looking up. SPC is seen in the upper right side of the drawing, looking around the edge of the TPS. SPAN-B is seen in the top left side of the drawing. SPAN-A is visible in the lower left side. The SWEM is internal to the spacecraft bus and not shown in this visualization, but is located about halfway up the hexagonal spacecraft bus on the same panel as SPAN-B (image credit: SWEAP collaboration)

Figure 19 illustrates the current mechanical design of the Solar Probe Plus spacecraft, along with the placement and orientations of the SPC, SPAN-A, and SPAN-B instruments, which are highlighted. In this image the spacecraft is oriented with the heat shield on the right and the spacecraft bus on the left, the ram-facing side of the spacecraft is pointing down and the anti-ram side of the spacecraft is pointing up. SPC can be seen on the edge of the heat shield with a small suite-provided strut that allows the sensor of the instrument to face the Sun while the high voltage and signal detection electronics sit in the shadow in the small box at the end of the strut. SPAN-A with its two electrostatic analyzers (electrons and ions) is visible on the lower left side of the image on the ram facing side of the spacecraft. SPAN-B, with its electron electrostatic analyzer, is visible on the upper left side of the image. Not seen in this image is the SWEAP Electronics Module (SWEM), which is embedded within the spacecraft bus.

From these locations the instruments are able to detect electrons over almost the entire sky, and ions from within thirty degrees of the Sun and from the ram direction. The SPC strut is designed to place the instrument sufficiently close to the front of the spacecraft heat shield that there are no obstructions to its field of view, which is a 60º full-width cone in the sunward direction. The SPAN electron fields of view and planned angular pixels are shown in a two dimensional projection of the sky in Figure 20. This presents the SPAN-A field of view and pixels as the blue contours, the SPAN-B field of view as the yellow contours, and the outline of the spacecraft, as seen from SPAN-A, as the white trace. See how SPAN-B observes electrons over the region of sky blocked by the spacecraft for SPAN-A. Higher angular resolution in the SPAN pixels is determined by the size of anode detectors within the instruments. Regions of higher angular resolution are oriented to provide the best chance of resolving the electron heat flux when the solar wind magnetic field is in the standard Parker spiral configuration in the ecliptic plane. The fields of view of the SPAN instruments are unobstructed except for minor features including the FIELDS antennas, the FIELDS magnetometer boom, and the solar panels. Note that as the spacecraft approaches the Sun the solar panels move closer to the spacecraft and further from the SPAN fields of view. Adaptive masks will be used to discard pixels too close to the solar panels from processing of electron observations for science products.

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Figure 20: Electron FOV of the instruments, showing combined anode/deflection angle maps for electron energies up to 4.5 keV (blue = SPAN-A, orange = SPAN-B). The (0º, 0º) direction is sunward, and (90º, 0º) points to the ram. Spacecraft obscurations are outlined in blue for SPAN-A and red for SPAN-B. The only portions of the sky not covered by one of the two sensors are those directly sunward (blocked by heat shield) and the small surrounding region (110º, -50º). The portion blocked by the heat shield is visible to SPC up to electron energies of 2 keV (the 29º half-width FOV of SPC is not shown on this figure), image credit: SWEAP collaboration)

Suite Electrical Interface and SWEM: The SWEM (SWEAP Electronics Module) is the data and power interface between the SPC and SPAN instruments and the spacecraft. It provides intermediate data processing (moment calculations, mode switching, data compression and data storage) for the SWEAP data as well as centralized power conditioning. In general, the SWEM acts as the SWEAP suite central processing unit, housing command scripts for sensor initialization and configuration, high voltage control, and attenuator actuation and accumulating data into various packets for inclusion into the data stream to be transferred to the SPP SSR (Solid State Recorder).

The SWEM houses most of the spacecraft interface electronics for SWEAP in a single box to reduce mass, to simplify harnessing and interfaces, and to reduce duplication of interface logic, power converters and other common services. SWEM allows for a single interface to the spacecraft DPU, easing integration and test activities later in the project. The module is composed of 2 separate approximately 6U-VME sized boards (160 x 200 mm): the DCB (Data Controller Board), and the LVPS (Low Voltage Power Supply board).

The SWEM derives its design from THEMIS, HESSI, STEREO, and MAVEN, and makes extensive use of the technical capabilities demonstrated in the IDPUs (Instrument Data Processing Units) developed for these missions. The suite data interface and control design is derived from the MAVEN DCB , modified slightly to incorporate the different SPP interfaces. The Coldfire processor implementation was qualified for use by MAVEN, and the design will be copied for SWEAP. The remaining circuitry needed on the DCB and LVPS are standard designs used extensively in previous missions.

DCB (Data Controller Board): This board receives commands and timing signals from the SPP DPU, and generates initialization and configuration messages for the SPANs and SPC. It acts as a router during data collection and provides intermediate processing such as on-board moment, pitch angle and averaging computations. Upon low voltage turn-on of the SPC and SPAN, the DCB will configure the sensors, enable the low voltage line that powers the high voltage supplies and execute the HV ramp-up commands that bring the supplies to nominal voltage. The DCB includes hardware and software safety latches to prevent accidental high voltage turn on. Attenuator position is controlled by the DCB, which also compresses the data into averages and calculates onboard moments, implemented in a single FPGA/embedded processor. The DCB FPGA also provides the SWEAP to SPP DPU interface, controlling the flow of raw data through the system. It receives commands from the SPP DPU processor and manages the SWEAP suite. Following the RBSP and MAVEN architecture, the DCB has memory on-board (rad-hard and SEE tolerant 10 x 8 GB Flash Memory that is the same as is currently flying on MAVEN) and takes formatted packetized data from the instruments directly making them available when requested for transmission to the ground without further processing. Support circuitry on this board includes address/data demultiplexing and an ADC for housekeeping data.

FSW (Flight Software): The core of the SWEM DCB is a Coldfire processor implemented in an FPGA, as an IP-core. The Coldfire processor was used on MAVEN and builds on heritage code that has flown on many missions. This processor and the board electrical design was developed for UCB's Instrument package on RBSP, allowing SWEAP to take advantage of both hardware and software heritage.

The SWEAP FSW architecture uses dedicated FPGAs for routine data and memory management, instrument interfaces, and other repetitive tasks in order to leave the processor free for specialized duties and less frequent, higher level functions. Instrument housekeeping and commanding is implemented via low speed serial data interfaces in a point-to-point architecture. Instrument science data is relayed over separate, higher speed serial lines to the DCB where data is directly fed into on-board memory, formatted as packets, and sent to the S/C mass memory for eventual downlink transmission. There are no time intensive tasks and CPU usage as a fraction of total available resources is low. Computation intensive tasks are performed in the instrument FPGAs using dedicated HW, the CPU only supports their operation.

LVPS (Low Voltage Power Supply board ): The SWEM LVPS generates the various voltages required by the SWEAP instrument suite. Regulated voltages are PWM (Pulse Width Modulated) and regulated to better than ±5 %. This board also provides current monitors for secondary voltages. Each power supply is current limited on its primary side and is galvanically isolated primary to secondary. The input from the 28 V spacecraft power is soft started and filtered to meet EMI (Electromagnetic Interference) requirements.

Science Data and Operations: The Suite is operated through the SWEM. The science data is taken along with calibration sequence run according to a command plan uploaded and then executed by the SWEM.

The SWEM has a number of independent operating configurations, which mainly affect instrument science data accumulation rates. Only two basic modes (Safe and Science Mode) affect power consumption and dissipation. In Safe Mode, only the core systems are powered on. Safe Mode is the power on condition, with only the core power and processing system powered on (the instrument sensors are not powered on). Science Mode is the normal operating state. In this mode, SWEAP is ready for full science data collection and is autonomously controlled using table driven mode configurations. In Science Mode, sensor data is sent to the SWEM at a constant rate. While the specific contents of SPC and SPAN data are different, all communication between the SWEM and instruments is accomplished through the standardized CCSDS protocol. Upon receipt of instrument data the SWEM makes use of FPGA-based logic to steer incoming data to the assigned memory locations in real-time based on programmable tables. Once in memory, the FPGA logic moves data through the data processing and compression pipeline, completes the data packetization, inserting ApIDs, and queues the data for transfer to the spacecraft.

Operation of SPC by the SWEM is a simple matter of uploading operating parameters to the SPC FPGA. The FPGA will then control the high-voltage modulator board and record telemetry that is then sent back to the SWEM. The SWEM contains a number of pre-programmed configurations to place SPC into the various combinations of ion, electron, peak-tracking, full-sweep, flux-angle and calibration modes. SPC sends 1 science packet/s to the SWEM.

The SPAN sensors are powered on via command from the SPP DPU to the SWEM. The SWEM contains command scripts for SPAN sensors to initialize the LV electronics and place the sensors in a default test mode, enable HV and bring HV to programmable levels, run various test sequences that confirm proper operations, and command the sensor into an instrument mode. The SPANs always make measurements at the same rate, but SWEM can load new high voltage lookup tables and/or trigger a switch between the operative high voltage table in the SPANs, which affects the energies and angles that a SPAN sensor scans in each measurement cycle. SWEM can also load and/or trigger a switch between the lookup tables that determine how counts are accumulated and binned in the SPAN sensor FPGAs. The combination of high voltage and product lookup tables defines the instrument operation, with the resulting sensor operational modes.

SPC (Solar Probe Cup):

SPC is a Faraday Cup (FC) with a 60º FOV (full width) that views the Sun from the edge of the SPP heat shield and measures the RDF (Reduced Distribution Function) and flow angles of ions and electrons as a function of energy/charge at high frequency. SPC measurements are essential for SPP and SWEAP because otherwise the solar wind would often be blocked by the heat shield. FCs measure the current produced on a metal plate by charged particles with sufficient energy/charge to pass through a grid placed at a variable HV (High Voltage). The underlying technology is straightforward and similar to the operating principles of vacuum tubes. SPC filters charged particles based on the component (E||/q) of their energy/charge parallel to the instrument line of sight. For a given E||/q, particles with any E/q can enter SPC as long as their flow angle is within the FOV. In a single measurement the HV oscillates between two voltages. The population of plasma with E||/q between these two voltages produces an AC current on the plate, and electronics isolate the AC signal and record it. SPC measures the currents produced on a circular plate divided into four quadrants and placed behind a smaller circular aperture. A cross section of the SPC sensor and an illustration of the measurement process is provided in Figure 21.

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Figure 21: Upper-left: a distribution function showing three voltage windows highlighted in different colors. Lower-left: Cross-section of FSU (Faraday cup Sensor Unit). The upper-section with a larger diameter is the ‘modulator assembly' with the high-voltage modulator grid highlighted in orange. The electric field set up between the modulator grid and the adjacent ground grids is shown as yellow arrows. As the modulator grid in the FSU oscillates between two voltages (upper-right panel), only particles with energy/charge between the voltages produce an AC current on the collector plates (lower-right panel). These currents are amplified and fed to a synchronous detection circuit (image credit: SWEAP collaboration)

The combined currents are directly related to the RDF, and ratios of the currents recorded by the different collector plates can be solved for the precise flow angle of the plasma as a function of energy/charge. Detailed properties of the solar wind such as velocity, density, and anisotropic temperatures are then determined by convolving a model velocity distribution function with a detailed instrument response function and deriving the best fit solar wind parameters. The AC detection process makes SPC insensitive to lower-frequency noise sources, such as thermionic emission from hot surfaces, photoelectron emission from surfaces exposed to sunlight, and penetrating radiation from intense energetic particle events.

The FC has been a workhorse for space plasma measurements from the beginning of space exploration. FCs have flown on Explorers 10, 18, 21, 28, 33, 35, 47 and 50, SOLRAD A and B, Pioneer-6 and -7, Mariners-3, -4, -5 and -10, OGO-1 and OGO-3, Voyager-1 and -2, and Wind. FCs have played a major discovery role in exploring space, from the first observations of ion fluxes in space through the first observations of the termination shock. The FCs on Voyager-2 and Wind are still operating and providing invaluable solar wind measurements. The FC components and measurement process both function in a high temperature environment. Since the sensor is simply a metal plate, the response of the instrument is stable with time. The FCs on Wind have a drift in response of about 0.01%/year.

A significant effort has been expended in an assessment of materials for SPC, solar furnace testing of materials in an optical, thermal, and radiation environment well in excess of the closest SPP solar encounters, design and fabrication of a prototype, and successful operation a prototype instrument in a flight-like environment. Thermal and optical simulations predict that SPC temperatures at closest approach will reach over 1600ºC. Standard aerospace materials at these high temperatures can evaporate or outgas, especially when sputtering of solar wind ions and electron beam heating are factored in, leading to changes in thermal/optical properties and degrading performance. Due to the Sun-exposed materials at high voltage, the project must also factor in thermionic emission due to both the hot cathode and field emission mechanisms, along with photoelectron induced emission hundreds of times more intense than generally seen near Earth. For the grids, the project found the best materials to withstand the harsh conditions are high purity tungsten (W). Tungsten has a high work function and the highest melting temperature of the refractory metals, and is also simple to etch into grid patterns. Other materials employed in the sensor include pure Niobium, and an alloy of molybdenum, titanium and zirconium called TZM. The niobium is employed in the heat shield and in wires due to its flexibility and the ease with which it can be laser welded. The TZM was selected mainly because it is able to retain its strength at high temperatures and due to the relative ease with which it can be machined. Sapphire has emerged as the preferred material for insulating surfaces due to its high electrical resistivity at high temperatures combined with its relative mechanical strength compared to other materials such as pyrolytic Boron Nitride.

SPC is divided into several smaller units, which are described and shown in Figure 22. The FSU (FC Sensor Unit) is the actual FC sensor that is exposed to sunlight. Two annular niobium plates at the front of the FSU permit plasma to flow into the instrument while shielding the edges of the SPC from sunlight. The FSU is mounted on the end of a titanium strut that interfaces SPC to the SPP transition structure. The FEU (Faraday cup Electronics Unit) is located on a platform directly next to the interface point on the SPP transition structure. The FEU contains preamplifiers that amplify the signals from each of the four collector plates. An FPGA and ADC (Analog-to-Digital Converter) in the FEU digitize the amplified waveforms and perform the synchronous detection algorithm. The digitized currents are sent from the FEU to the APB (Analog Processing Board) in the SWEM, where further calculations are performed and the resulting packets are sent to the spacecraft solid-state recorder. The FEU also contains the HMB (High-voltage ModulatorBoard), which is driven by the FPGA and generates the High Voltage (HV) waveforms that drive the HV modulator grid within the FSU. The FEU is within the shadow of the SPP heat shield, allowing for a relatively benign operating environment, with maximum temperatures expected in the electronics box to be approximately 70ºC.

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Figure 22: Layout of SPC alongside the spacecraft with key components labeled (image credit: SWEAP collaboration)

A cross section of the FSU is shown in Figure 23. The modulator assembly filters particles by energy/charge and the smaller collector assembly records the flux and flow angles of the particles. Plasma enters the FSU through the large circular entrance aperture at the top of the instrument and enters the modulator assembly. The entrance aperture is designed to be sufficiently large so that the smaller limiting aperture between the modulator and collector assemblies is always fully illuminated by plasma over the entire range of angles of incidence. The limiting aperture is sized to detect the minimum flux with sufficient signal to noise, including the drop in signal strength due to the transparency of the grids. Fine conducting grids with high transparency throughout both assemblies act to filter particles and shield stray electric fields. After the entrance aperture a grounded grid prevents stray fields from leaving the instrument to ensure compatibility with electromagnetic instruments on the spacecraft.

Within the modulator are three more grids: a HV modulator grid with a time-varying retarding potential sandwiched between two grounded grids. The particles that pass through the limiting aperture then produce a current on one of four 90º wedge-shaped metal collector plates. The collector plates are larger than the limiting aperture, so an incident beam falls entirely on the collector plates. Three plates are sufficient to solve for flow angles, but a fourth plate provides an additional level of redundancy. A suppressor grid is maintained at -55 V immediately above the collector plates. The suppressor reflects secondary electrons kicked off the plates from ion impacts back onto the plates and prevents low energy electrons from thermionic and photoelectric emission at the front of the FSU from reaching the plates, although this is not necessary because they do not produce an AC signal. The currents from the four plates are fed through coaxial cable to the FEU, where the measurement electronics amplify and digitize the waveform.

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Figure 23: Cross-section of SPC with key components labeled (image credit: SWEAP collaboration)

SPC Estimated Performance: The three key SPC performance parameters are energy resolution, signal to noise, and accuracy of angle measurements. Initial performance testing has been done with the SPC prototype and indicates that SPC will meet all of its high-level performance requirements. A twopronged testing approach was taken, with high-fidelity ion beam testing taking place at the SWF (Solar Wind Facility ) at MSFC (Marshall Space Flight Center) and solar illumination testing taking place in a custom-built chamber at the SAO (Smithsonian Astrophysical Observatory). In order to provide some simultaneous testing, the SWF was fitted with a contact heater used to raise the temperature of the back of the collector housing and collector plates and the chamber at SAO was fitted with a lower precision ion source. Performance testing of the prototype in phase B demonstrated that SPC is able to meet all of its level-1 performance requirements.

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Figure 24: Response of different energy windows to an ion beam of increasing energy (image credit: SWEAP collaboration)

To measure the energy resolution of SPC, a prototype was tested using the ion source at SAO. The instrument was run in a limited ‘full-scan' mode, in which the current is measured in successive voltage windows by scanning in an upward direction and then again in a downward direction. In this case, the voltage windows are 7 % wide and overlap by about 2 %. The energy of the beam is then incrementally increased in a slow manner, so that multiple measurements are made in each voltage window at each beam energy. The results are shown in Figure 24. Voltage windows are shown as different colors (the colors repeat every 6th window). The peak of each voltage window is separated from the adjacent windows, showing that the instrument is able to resolve the energy of an ion beam to better than the separation of these energies windows (in this case 5 %). Further tests in phase C will further quantify the energy resolution of the instrument.

The angular response of the SPC prototype was measured using the SWF ion beam facility. A mono-energetic ion beam was produced that covered the full aperture of the SPC prototype. The instrument was then rotated through a range of angles while being exposed to this beam. Figure 25 shows the normalized current (in color) as a function of the angle of incidence of the ion beam and the energy window.

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Figure 25: SPC prototype response to a fixed energy beam at different angles of incidence (image credit: SWEAP collaboration)

As the instrument is rotated, the ion beam appears in different energy windows because the modulator portion of the instrument affects only the parallel energy of the incoming ions. As such we expect a cos2(θ) dependence of energy on incidence angle. The expected response is shown as a black line. The vertical green bar indicates SPC's FOV requirement. The beam is easily measured up to 10º beyond the 30º requirement. Testing was also performed to demonstrate the angular resolution of the instrument. Observations showed that the resolution is much better than 1º. This is confirmed by an analysis of the geometry of the cup and the measured noise performance of the measurement circuitry. Further testing in phase C will continue characterizing the angular response of the instrument.

SPC Electrical: The SPC Electronics Module is located at the base of the SPC strut and consists of a HVPS (High Voltage Power Supply), a LVPS (Low Voltage Power Supply), and a signal processing and control board referred to as the FEU (Faraday cup Electronics Unit). The HVPS provides the modulator and suppressor grid voltages. The modulation voltage is adjustable from -2 kV to +8 kVDC with an adjustable amplitude 50–800 V, 1280 Hz, AC component. The high voltage DC, AC component peak to peak voltage, and frequency are programmable via digital interface. The suppressor grid voltage is nominally -55 VDC. For each collector plate, the signal processing electronics first convert the input current waveform to a voltage waveform. This voltage waveform is amplified by four parallel gain stages, with the resultant waveforms are digitized by an ADC. The digitized waveforms are then routed to an FPGA that implements a discrete Fourier transform at the modulation frequency. The output of this algorithm is the peak-to-peak amplitude of the collector plate current at the modulation frequency. The gain-stage with the best SNR without saturating the amplifier is then selected and stored. The noise level of the system is about 5 x10-13 A, and currents (including both alphas and protons) are expected to range between 5 x 10-13 to 10-7 A. Thus, the SNR in a single measurement will range between 1 and 105. Protons, being the dominant species, will always have an SNR of more than 10.

SPC Mechanical - FSU materials and Fabrication: The dimensions of the FSU are directly derived from the desired sensitivity (the limiting aperture sized for the minimum expected current at 0.25 AU), range of observable flow angles (56º FOV) and energy range (modulator sized to withstand 8 kV). The FSU was designed following standard practices for solar wind FCs but with certain mechanical parts replaced with components made from materials more appropriate for higher temperatures. A stack of two thin annular niobium plates mounted to the front of the FSU limit direct exposure of the FSU and strut to sunlight while permitting access to the limiting aperture. The modulator and collector assemblies are machined from a high-temperature alloy called Molybdenum TZM (moly-TZM). The Delran insulators that have historically been used in Faraday cup to hold the HV grid in place and isolated from the modulator walls has been replaced with machined sapphire rods. The woven tungsten wire meshes used previously have been replaced with monolithic grids etched from single wafers of high purity W. We have found that these new grids are significantly stronger and more suitable for FCs than the old wire meshes, which were prone to failure through breaking wires and took weeks to manufacture.

Components are stacked within each subassembly, which is then sealed from the back with a TZM plate. The stack consists of niobium spacer rings, individual grids, and sapphire spacer rods and rings. High-voltage is transmitted to the modulator grid using a niobium wire fed from the side of the modulator assembly and bonded to a niobium annulus via crimping and laser-welding. The HV grid is then sandwiched between the niobium annulus and a tungsten annulus. The HV wire exits the side of the modulator housing and enters a custom hard coaxial cable. In that cable, sapphire tubes insulate the HV wire from a niobium outer tube, which is at ground. In the modulator assembly, ground grids make a firm mechanical and electrical connection to each other, to the ends of the assembly, and to the niobium spacer rings. This stacking technique makes assembly straightforward and permits mechanical flexibility in the presence of temperature gradients.

The structure of the collector assembly is similar, but ends with the four Niobium collector plates seated within a shallow sapphire cup. Signal wires are laser welded to the protrusions on the back of each collector plate that extend through the collector housing cap. Each signal wire then runs through hard coaxial cable (of similar construction as the HV coaxial cable) to the FEU. Figure 26 shows an image of the collector plate assembly which is the final component to be stacked into the collector housing. The four niobium collector plates sit on a bed of sapphire and are then clamped in place by a sapphire annulus that is held down by four bolts. In final stack-up, the suppressor grid would be placed directly against the top surface of the upper sapphire annulus.

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Figure 26: The four collector plates within their sapphire housing (image credit: SWEAP collaboration)

SPC Thermal: Four factors motivate keeping temperatures low within SPC: electron emission, mechanical integrity, signal amplification, and heat conduction to SPP. Hot conducting surfaces will generate thermionic emission that will be enhanced in the presence of strong electric fields by field-effect emission. These low energy electrons (several eV) can draw power from the HV supply by flowing to the modulator grid, and potentially interfere with the SPC solar wind signals. Thermionic emission from W materials peaks at about 1200 °C and is predicted to produce no more than 0.1 mW extra power draw on the SPC HV supply. Note that the accuracy of SPC flow angle and flux measurements are decoupled from any thermal or mechanical changes within the modulator assembly. Any change in the spacing between grids does not change the energy cutoff because they are maintained at a fixed voltage. Since the limiting aperture is always illuminated it is insensitive to any shifting of the entrance aperture.

Thermal simulations of SPC have been performed and results are shown in Figure 27. During closest encounters the first grid in the FSU will reach above 1600ºC and the modulator housing walls will reach 1000ºC. Though still hot, the components at the anti-sunward side of the cup are relatively cooler; the collector plates will be about 700 ºC and the back of the collector housing will be approximately 600 ºC. This is well below the temperature range the materials and components have been tested to.

The thermal model was compared to measurements taken under realistic illumination conditions with the phase B prototype in the SES ( Solar Environment Simulator). The SES uses four short-arc Xenon lamps with broad-spectrum reflectors, water cooled mirrors and focusing optics, and a water-cooled window into a vacuum chamber to deliver precisely calibrated, stable light that mimics the intensity, spectrum, and angular spread of sunlight at different phases in the SPP mission.

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Figure 27: Thermal model of SPC, showing predicted temperatures at close approach. Temperatures are in degrees Celsius (image credit: SWEAP collaboration)

A scanning optical thermopile with NIST-traceable calibration is used to calibrate and control the photon flux of the SES onto SPC, allowing us to precisely validate the optical and thermal performance of the instrument against a known photon input. A comparison of the thermal model and measurements from thermocouples distributed throughout the instrument prototype is shown in Figure 28. In this particular exposure three different illumination cases were presented to the instrument, with the temperatures allowed to come to equilibrium before moving to the next case. One can clearly see that not only is the model a good predictor of the equilibrium temperatures, it also performs well during the transition periods, providing further confidence that the assumptions in the thermal model are accurately reflecting the instrument. In addition to validating the instrument thermal design and performance, the SES has also been used to demonstrate that the instrument can survive the extreme thermal gradients and mechanical strains included by the photon-induced heating, to confirm predictions of thermionic and photoelectric emission, to verify that the instrument high voltage performance as expected over temperature, and to confirm the energy resolution and measurement noise are uneffected by changes in light level and temperature.

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Figure 28: Thermal model predictions (dashed lines) compared to measurements in the SES (solid lines) with the phase B prototype as a function of time. In this 1.1 hour experiment SPC was exposed to three different light levels, with time allowed for the instrument to come into thermal equilibrium at the new level (hundreds of seconds) and for instrument performance characterization measurements to be taken (thousands of seconds), image credit: SWEAP collaboration

SPAN (Solar Probe ANalyzers):

SWEAP has three ESAs (Electrostatic Analyzers) to measure the three dimensional velocity distribution functions of ions and electrons with high angular, temporal, and energy resolution. The SPAN-A module has two ESAs to measure ions and electrons from the ram direction and nadir, and SPAN-B consists of a single ESA to measure electrons from the anti-ram direction. SPAN-A is located on the ram-direction side of SPP and SPAN-B is on the anti-ram side. Significant savings in mass are realized by combining the electron and ion ESAs (a lesson learned from FAST and THEMIS). Electrostatic deflectors extend the narrow planar intrinsic angular FOV of each ESA to 240° x 120°. Together the SPAN electron sensors provide a nearly 4π sr FOV for electrons (excluding the region of the sky blocked by the heat shield). Meanwhile, SPAN-A and SPC provide a continuous view of the solar wind ions, with SPAN-A providing the primary measurement at closest approach, when the velocity aberration from the lateral motion of the spacecraft brings the solar wind into the field of view. SPAN-A includes a pre-acceleration stage and carbon foils, closely based on the MAVEN STATIC design, allowing the separation of solar wind protons, alpha particles, and heavier species. All three sensors include both mechanical and electrostatic attenuators that provide a broad dynamic range, allowing optimal sensitivity over the entire SPP orbit.

SPAN Estimated Performance: In the plane of the instrument aperture, the intrinsic azimuthal resolution of the SPAN sensors is very high because of the focusing properties of the top-hat electrostatic analyzer, but the actual azimuthal angular resolution is determined by the size of the discrete anodes. In order to minimize power required from the discrete preamplifiers, angular resolution over the 240° FOV of both the SPAN-A and SPAN-B electron sensors is divided into eight high resolution 6° anodes, and eight lower resolution 24° anodes. The high-resolution anodes are chosen to image near the sun in order to capture the electron strahl. The SPAN-A and SPAN-B electron sensor FOVs are oriented to optimize their combined FOV. This orientation, with the SPAN-B FOV rotated 90° relative to the SPAN-A FOV, provides optimal coverage of electrons up to 4.5 keV over the whole sky. The finely segmented anodes are placed optimally to measure the strahl, and the alignment of the electron FOVs ensures full energy coverage towards the Sun.

All three SPAN sensors typically operate in "Alternating Sweep" mode, interleaving low resolution measurements of the entire available phase space (sparsely sampled) with full resolution targeted measurements of selected portions of the distribution (fully sampled). In Figure 29, an example measurement of the solar wind protons is shown to demonstrate SPAN-A's ability to cover all of phase space, while still resolving the peak, at high cadence (both distributions can be obtained within less than half a second). The alpha particles, also present in this energy range, are separated using the mass-resolving capability of SPAN-A, and not shown in this Figure. With the expected mass/charge resolving capability of SPAN-A, the project will be able to generate fully distinct VDFs for protons and alphas.

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Figure 29: Proton energy spectra and angular distributions demonstrating the expected measurement resolution for coarse (left) and targeted (right) sweeps, for a two-component solar wind proton distribution at closest approach (image credit: SWEAP collaboration)

SPAN Electrical: Primary power for each SPAN sensor is provided by the SWEM. Each sensor has a low voltage power converter that receives regulated 22 V from the SWEM and converts it to provide analog and digital secondary voltages isolated from the primary service. These voltages are used to operate a digital board, high voltage power supplies, and front end electronics. A backplane also provides interconnects between the individual electronics boards, some of which are also connected directly by stacking connectors.

The high voltage power supplies provide fixed (adjustable) bias voltages for the MCP (Microchannel Plate) detectors and sweeping (lookup table-driven) voltages for the inner hemisphere, electrostatic deflectors and spoiler electrostatic attenuator. For the ions, an additional 15 kV supply provides a pre-acceleration voltage for the time of flight section. The sweeping supplies utilize optocoupler circuitry for the outputs. The high voltage supplies utilize the primary 22 V from the SWEM, switched through the digital board, and controlled by DAC outputs from the digital board. For the electrons, all high voltage power supplies share one board, while for the ions the pre-acceleration and MCP supply are housed on a separate board from the sweep supplies.

The front end for the electron sensors consists of a single board which provides a mounting point for the MCPs, collects charge pulses from the back face of the MCPs on a segmented anode, and utilizes a 16-channel preamplifier ASIC to produce digital pulses suitable for accumulation in the digital board.

The front end for the ion sensors consists of an anode board that provides a mounting point for the Z-stack microchannel plates and collects output pulses on segmented start and stop anodes. Each anode is capacitively coupled to an ASIC-based constant fraction discriminator providing a digital output pulse. These pulses are fed into a time-to-digital conversion utilizing ASIC parts, located on the digital board.

SPAN Mechanical: The mechanical design of SPAN-A and SPAN-B as of Mission PDR is summarized in Figure 30 (SPAN-B) and Figure 31 (SPAN-A). Since the electron ESA in SPAN-A is essentially identical to the electron ESA in SPAN-B, we begin with a review of SPAN-B.

Starting from the left side of Figure 30, the top of SPAN-B consists of a deployable onetime cover (re-closeable on the ground) that protects the ESA from contamination before launch. Immediately below the cover is a mechanical attenuator that is used to regulate the geometric factor of the sensor. The main optics section begins with a broad entrance aperture surrounding a pair of electrostatic deflectors that steer the instantaneous field up and down. After passing through the deflectors, particles travel between a pair of hemispheres, which have a voltage between them that only allows particles in a narrow band of energy to reach the exit. By placing a small voltage on the bottom half of the outer hemisphere ("the spoiler") we can reduce the sensitivity of the analyzer by an order of magnitude, essentially be de-tuning the response of the analyzer. Particles that successfully traverse the electrostatic optics impact a chevron pair of microchannel plates with a high voltage bias across them, which produce a secondary electron cascade with a multiplication factor of a few million, producing a pulse with measurable amplitude that is collected on a segmented anode.

A central tower penetrates the anode and digital boards, bringing high voltage to the inner hemisphere from the high voltage power supply. The anode board is directly mounted in a transition plate which provides a smooth join between the analyzer and the electronics box, while the succeeding electronics cards (digital, high voltage power supply, and low voltage power converter) are mounted in frames which stack together, with a back plane mounted perpendicularly and providing signal and power routing between electronics boards.

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Figure 30: Cross section of the SPAN-B instrument (image credit: SWEAP collaboration)

SPAN-A is shown in cross section in Fig. 31. The SPAN-A electron sensor has an identical mechanical design to SPAN-B, but is mounted together with the SPAN-A ion sensor. The ion sensor also has a nearly identical design, including the same analyzer optics, but with a few critical differences. Instead of an anode board directly at the exit of the analyzer, the SPAN-A ion sensor has a time of flight sensor. A pre-acceleration voltage accelerates ions to a high enough energy to reliably penetrate a carbon foil, producing start electrons that are then guided to the inner half of the microchannel plates. The ion then continues to a second carbon foil, generating a stop pulse. The stop electrons have sufficient energy to penetrate a thick foil and generate a stop pulse, while the primary ions stop in the thick foil, eliminating a primary background source for ion mass composition sensors. The remainder of the electronics box has a similar design, albeit with more boards than the electron sensors in order to accommodate the additional high voltage power supply for the pre-acceleration.

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Figure 31: Cross section of the SPAN-A+ instrument. In this orientation the electron instrument is on the left and the ion instrument is on the right (image credit: SWEAP collaboration)

SPAN Thermal: Since the SPAN instruments are in shadow during scientific operations close to the Sun, the thermal design of these sensors is much more straightforward than for SPC. However, this does not mean that it is completely straightforward. Indeed, the SPAN sensors must contend with a somewhat non-intuitive thermal problem. Far from the Sun, the spacecraft must be able to rotate in order to orient the high gain antenna on the anti-ram side of the spacecraft toward Earth for communications. This results in prolonged illumination of the SPAN-B sensor. The level of heat absorbed by SPAN-B must be limited in order to prevent damage to the deployment and attenuator mechanisms in the electron ESA sensor. Unfortunately, treatment of SPAN-B surfaces to reduce absorption of heat far from the Sun when the spacecraft is rotated and the anti-ram side is illuminated results in very low operating temperatures for SPAN-B when it is close to the Sun. Ironically, at closest approach, while the Sun-facing instruments and spacecraft components struggle with extreme high temperatures, SPAN-B risks falling below its minimum operating temperature. As a result, an operational heater was added to SPAN-B.

Science Operations, Data Processing & Data Products:

The SWEAP instrument suite operates autonomously during each solar encounter, with the SWEM controlling each of the SWEAP sensors, formatting data products, archiving high resolution data to internal storage in the SWEM, and transmitting summary data to the spacecraft for quick transmission to Earth following each encounter. As of PDR, the SWEAP Suite was allocated 20 Gbit of data volume per orbit to meet the Level 1 science goals.

Science Operations: SPP runs on a bent-pipe operations scheme, in which the SPP MOC (Mission Operations Center) and the SPP spacecraft will act only as a pass through of the command plans and telemetry to and from the SWEAP Investigation SOC (Science Operations Center). The SWEAP SOC will plan orbital command plans, select full resolution data, data processing, distribution and archiving. The SOC will also be utilized during integration and test activities, mission simulations, and to commission the SWEAP Suite.

In summary, the SWEAP Investigation will obtain the thermal coronal and solar wind plasma observations needed by the scientific community to address these compelling questions. The four SWEAP sensors provide complete measurements from several eV to tens of keV of the velocity distribution functions of electrons and ionized helium and hydrogen (alpha particles and protons) that constitute the bulk of the solar wind and coronal plasma, along with properties of other ions sorted by their mass/charge. Entering a new and harsh environment for the first time, the instruments face technical challenges unlike any other mission to date.

 


 

WISPR (Wide-field Imager for Solar Probe)

WISPR is an NRL (Naval Research Laboratory) instrument. The imager is a telescope, which looks off to the side of the heat shield, and will make 2D images of the sun's corona as the spacecraft flies through. But like a medical CAT scan, the orbit of the spacecraft through the corona will enable 3D images and a determination of the 3D structure of the corona. The experiment actually will see the solar wind and provide 3D images of clouds and shocks as they approach and pass the spacecraft. 33) 34) 35)

- NRL is WISPR PI institution for NASA; consortium of US and international science partners

- NRL will develop the WISPR instrument and will operate it after launch

- Builds on successful NRL SECCHI heliospheric imagers on the NASA STEREO mission.

The objectives of WISPR are:

• Understand the morphology, velocity, acceleration, and density of evolving solar wind structures when they are close to the Sun.

• Derive the 3D structure of the solar corona through which in-situ measurements are made to determine the sources of the solar wind.

• Determine the roles of turbulence, waves, and pressure-balanced structures in the solar wind.

• Measure the physical properties of SEP (Solar Energetic Particle)-producing shocks and their CME drivers as they evolve in the corona and inner heliosphere.

WISPR, with a 95° radial by 58° transverse field of view, will image the fine-scale coronal structure of the corona, derive the 3D structure of the large-scale corona, and determine whether a dust-free zone exists near the Sun. Given the tight mass constrains of the mission, WISPR incorporates an efficient design of two wide-field telescopes and their associated focal plane arrays based on novel large-format (2 k x 2 k) APS CMOS detectors into the smallest heliospheric imaging package to date. The flexible control electronics allow WISPR to collect individual images at cadences up to 1 second at perihelion or sum several of them to increase the signal-to-noise during the outbound part of the orbit. The use of two telescopes minimizes the risk of dust damage which may be considerable close to the Sun. The dependency of the Thomson scattering emission of the corona on the imaging geometry dictates that WISPR will be very sensitive to the emission from plasma close to the spacecraft in contrast to the situation for imaging from Earth orbit. WISPR will be the first ‘local' imager providing a crucial link between the large scale corona and the in-situ measurements.

 

WISPR instrument design:

The WISPR instrument concept is in effect a miniaturization of the SECCHI/HI concept with adaptations from the SoloHI design. It is a two-telescope system, similar SECCHI/HI, with an inner telescope extending from 13.5 º to 53º and an outer telescope extending from 50º to 108º (Figure 2, right). The two-telescope implementation is driven by the need to baffle the telescope against the intrusion of two of the FIELDS antennas (Figure 2, left) into the WISPR unobstructed FOV. The instrument uses the heat shield as the first occulter and hence the alignment between the heat shield and the first occulter baffle, F1 is a critical element for the successful control of the stray light. The inner FOV cutoff is set at an elongation of 13.5º from sun center, corresponding to a heliocentric distance of 2.3 Rs at 9.86 Rs perihelion. The cutoff is dictated by two requirements: (1) to remain within the heat shield umbra (8º, including a 2º maximum spacecraft offpoint), and (2) to accommodate the instrument on the spacecraft bus at a reasonable height and with reasonable mass. The overall instrument characteristics are shown in Table 4.

Telescope type

Wide angle lenses, aperture stop placed in front of lens.
Inner: f=28 mm, aperture=42 mm2, 490-740 nm (bandpass)
Outer: f=19.8mm, aperture=51mm2, 475-725 nm (bandpass)

Plate scale

1.2 – 1.7 arcmin/pixel (inner-outer)

FOV (Field of View)

95º radial x 58º transverse, inner field limit 13.5º from Sun center

Image quality

Predicted RMS spot including allowable tolerances at 20º from boresight:
Inner: 19.5 µm (2.34 arcmin)
Outer: 19.9 µm (3.38 arcmin)

Detector

APS (Active Pixel Sensor), 10 µm pitch, 2048 x 1920 pixels

Baffle Design / Stray Light Rejection

Front heat shield edge, forward baffle and diffraction light trap designed to reject incoming solar radiation, interior baffles and aperture enclosures designed to reject scattered solar radiation from spacecraft, structures, and thermal radiation from antennas. Average predicted stray light: <2 x 10-9 B/Bs @ 9.86Rs and <2 x 10-12 B/Bs @ 0.25 AU, well below the K+F corona.

Pointing

Instrument axes aligned to S/C to < 0.5º, F1 and heat shield leading edge placement error < 13 mm.
Baffles achieve adequate rejection with 2º excursion from sun center at perihelion.

Calibration

<20% absolute radiometric, platescale <4%, pointing: accuracy 5 arcmin (3σ), jitter 0.8 arcmin (1σ), windowed stability 1.6 arcmin (1σ)

Mass

WIM (WISPR Instrument Module) 9.8 kg; Instrument DPU (spacecraft provided) 1.1 kg

Average power

7 W (including 4W operational heater power)

Envelope

WIM module: 58 cm x 30 cm x 46 cm (door closed)

Average telemetry rate

Allocated data rate 26.6 kbit/s (during 10-day operational periods); 23 Gbit/orbit

Table 4: WISPR instrument characteristics

The instrument concept and s/c accommodation are shown in Figure 32. A set of forward occulters (F1-F3) are located on a ledge to reduce the diffraction from the heat shield. An internal baffle assembly I1-I7 reduces this stray light component further as well as stray light diffracted from the radio antennas and other spacecraft structures. A set of other baffles are located at the apertures of the two telescopes to prevent any further reflections from reaching the detectors. Because of the orbit profile, the WISPR stray light rejection requirements vary as a function of elongation angle and heliocentric distance by about an order of magnitude. The most stringent requirement is 1.8 x 10-12 MSB (Mean Solar Brightness) at the outer edge of the FOV (90º elongation) at the largest distance from the Sun (0.25 AU).

The sophisticated baffle design allows WISPR to meet this requirement and allows for a high SNR (Signal-to-Noise Ratio) imaging, ranging from 20 at the inner FOV at closest perihelion to 5 at the largest distance and FOV angles. The detectors are 2048 x 1920 format APS (Active Pixel Sensor) CMOS devices developed for the SoloHI (Solar and Heliospheric Imager) instrument program. 36) APS devices are much less susceptible to radiation damage than the more common CCD devices and are therefore the best option for this mission. They also come with significant savings in terms of power and mass. These devices are described in more detail in the following reference. 37) The devices are cooled to -60ºC via a passive radiator. A one-shot door is used to protect the baffles and optics from contamination during ground operations, launch, and early flight operations.

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Figure 32: WISPR instrument concept. Insert: WISPR accommodations on the SPP spacecraft (image credit: WISPR collaboration)

Optical Design: The WISPR optics comprises two lens assemblies with the parameters given in Table5 and shown in Figure 33. Both designs are based on the SECCHI/HI optics adjusted to the F# required for the WISPR application. The resolution is optimized for the FOV center, 33.5º and 79º, for the inner and outer telescope, respectively. Glasses such as LAK9G15, SF4, and SF6 were assumed for this design but the adoption of BK7 for the first lenses is being considered, because it may be more resistant to dust impacts . As can be seen from Table 2, the current optical design is excellent. It provides both fast lenses (low F#) and high spatial resolution (~2 pixels) for the inner and outer telescopes, respectively. This means that WISPR is potentially capable of capturing images at spatial resolutions of <2 arcmin (2200 km or ~ 3 arcsec from 1AU) which are comparable to eclipse imaging from Earth. This is truly remarkable for a wide-field coronal telescope and the capability will be exploited as mission and solar condition allow. However, the current observing plan is to obtain images with 2 x 2 binning, as is done for SECCHI/HI, to increase the SNR and reduce the telemetry load. Higher image binning (4 x 4) will be required at large heliocentric distances to maintain an SNR of 5 at the outer edge of the FOV.

 

FOV

Spectral range (nm)

Entrance pupil (nm)

F# (F number)

No of lenses

RMS Spot Size (µm)

Inner Telescope

40º x 40º

490-740

7.31

3.83

5-element

19

Outer Telescope

58º x 58º

475-725

8.08

4.04

6-element

20

Table 5: WISPR optical design

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Figure 33: WISPR lens assemblies for the Inner (left) and Outer (right) telescopes (image credit: WISPR collaboration)

Stray Light Control: The control of stray light due to spacecraft accommodations has been the major focus of the WISPR team during the preliminary design phase of the project. The AO strawman payload outlined a very optimistic imager concept with a single wide-filed lens and an unobstructed 180º FOV because it did not take into account the accommodations of the radio antennas in the forward section of the s/c and into direct sunlight. As a result, two of these antennas impinged either directly into the WISPR FOV or extended into the unobstructed FOV allowing diffracted sunlight to enter the aperture at unacceptable levels. In addition, the tips of the antennas will get so hot (~1800ºC) that they will radiate into the visible creating another (and novel) source of stray light. The only solution for allowing the instrument to operate was to baffle directly those two sources of stray light. In order to achieve this without sacrificing most of its FOV, the WISPR single lens was split into two separate imaging assemblies.

This change allowed the design of peripheral baffles that capture the diffracted and radiated light from the antennas and reduce the stray light to acceptable levels as shown in Figure 34. This is a preliminary result, however. The optimization of the peripheral baffle system is still under way in an effort to reduce the stray light further within the outer telescope FOV. The stray light modeling is performed via Monte-Carlo techniques with the FRED Optical Engineering software using a CAD model of the instrument and FIELDS antennas. This approach allows not only the modeling of the antenna diffracted and radiated light but also the testing of various coatings for the baffle surface and even the modeling of the effects of dust impacts during the mission. These new stray light modeling methods, driven by the need to accommodate occulting-like imagers in crowed s/c environments, far exceed the corresponding modeling efforts in past coronagraphs and imagers where tight controls of structure intrusions in the unobstructed FOVs were possible. They demonstrate that visible light imagers can be accommodated and operate safely even when structures intrude into their direct FOVs.

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Figure 34: Left: The peripheral baffles of the WISPR telescopes. Right: The improvement in stray light levels resulting from the 1- (top panel) to 2-telescope (bottom panel) design change and the optimization of the peripheral baffles (image credit: WISPR collaboration)

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Figure 35: Illustration of the WISPR instrument (image credit: NRL) 38)

Detector and Electronics: The WISPR electronics are divided into two components: the WIM ( WISPR Instrument Module) which controls the two cameras, the door deployment and the operational heaters, and the IDPU (Instrument Data Processing Unit), which passes the commands to the WIM and receives the camera data from the WIM. The WIM controls the two cameras, receives the analog data, digitizes it to 14 bits, removes cosmic rays, and adds individual images together to increase SNR. The functional block diagram is shown in Figure 36. The IDPU is being developed by APL based on specifications provided by the NRL team. The WISPR CIE (Camera Interface Electronics) is an adaption of the SoloHI electronics. The data is transferred from WIM to IDPU via a Camera-Link interface, is compressed and packetized and is then transferred to the onboard SSR (Solid State Recorder) via SpaceWire.

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Figure 36: WISPR electronics functional block diagram (image credit: WISPR collaboration)

Testing and Modeling of High Speed Dust Impacts on the WISPR Optics: The orbital velocities at the SPP perihelia are very high (~170 km/s) and therefore the distribution of the kinetic energy of the dust particles and their fluence have to be accounted for in the WISPR design. The mass and size distributions are not known but models suggest that the majority of the particles will have diameters < 10 µm. For example, the APL/UTEP model predicts about 100 impacts from 10 µm particles and 1000 impacts from 0.1 µm particles at the heat shield during the seven years of the mission. Dust impacts pose two risks for WISPR: (1) they can damage the edges of the forward baffles and thus increase the stray light levels and, (2) they can damage the front lens surfaces and lead to increased stray light by pitting and/or cratering the glass surface. Larger particles can of course break the lens but the probability of a catastrophic hit is exceedingly small (<10-5 for >1mm particle).

There is very little information on the effects of high velocity impacts on optical systems since such environments have not been encountered before. To understand the effects of dust on instrument performance, the WISPR team has established a glass testing and modeling program during the design phase as part of the contribution of the WISPR Co-I team in Germany (V. Bothmer, PI). Although there is no dust accelerator that can accelerate particles to 100s km/s velocities, the Dust Accelerator at the MPIK (Max-Planck Institute für Kernphysik) in Heidelberg is potentially capable of up to 60 km/s velocities. The facility was used in October 2012 to test three different candidate glass materials for the WISPR optics; BK7, BK7 with a diamond coating, and Sapphire. The tests were performed with a variety of iron particle distributions (0.5 – 3 µm) and speeds (0.5 - 8 km/s) against three different impact angles (0º, 45º, 70º) (shaded area in Figure 37).

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Figure 37: Left: The Size-Velocity distribution of the iron particles used in the WISPR Dust Impact experiment. Right: Examples of crater damage in the 3 glass types used in the experiment (image credit: WISPR collaboration)

The examination of the impacted glasses showed that sapphire was the most impact-resistant material with very small (2 µm) and symmetric crates. The diamond-coated BK7, on the hand, did not perform well. The impact caused a halo around the impact crate. The project believes that the halo is the result of the detachment of the coating locally due to the heat produced by the impact. The regular BK7 has relatively small craters (~5 µm diameter). The spall diameters are very consistent with the APL/UTEP model and give confidence in the overall SPP project dust analysis and rick mitigation procedures.

An automated software program was developed to measure the sized and numbers of craters in the images to assess the extent of the damaged area. The results (Figure 38) show a relatively linear damage increase with dust speed. Based on these results, we decided to adopt the regular BK7 as the baseline for the WISPR optics.

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Figure 38: Left: The effect of dust fluence on percent damage area for the three glass types used in the WISPR Dust Impact experiment. The lines are simple linear fits to the data. Right: Fits to the BTDFs (Bidirectional Transmittance Distribution Functions) of pristine (blue curve) and damaged BK7 (red/green curves correspond to different samples). The data points are marked by crosses (image credit: WISPR collaboration)

To study the effect on the imaging performance we then had to measure the change in the BSDF (Bidirectional Scattering Distribution Function) or Harvey-Shack function compared to the BSDF of the undamaged glass. So, the project measured the BTDF (Bidirectional Transmittance Distribution Function) of the damaged glasses and used the percent damaged area to make a linear fit to the BTDF (Figure 38, right). The measurements revealed that the model BK7 used in FRED was too conservative. We adjust those curves and run our stray light calculations for pristine and damaged WISPR lenses to calculate the beginning- and end-of-life performance of the optics. The results are shown in Figure 39.

To summarize, the dust testing and subsequent analysis has allowed the team to (1) validate the APL/UTEP model at those velocities and accepted for evaluating the risk for higher speeds, (2) reject exotic materials and coatings as an alternative to regular BK7, (3) develop a realistic BSDF model to evaluate the stray light effects of dust impacts on the imaging performance, and (4) get an estimate on the approximate damaged area of the WISPR optics. The team has kept the option open to return to MPIK for one more round of testing of either the flight glass or possibly a recently-identified ceramic glass material.

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Figure 39: Left: Comparison of the estimated stray light levels at Beginning- (BOL) and End-Of-Life (EOL) for the WISPR instrument. The modeling is based on damage to 0.6% of the lens area as predicted by the dust model and uses BSDF measurements from the dusts tests at MPIK. The higher levels in the inner telescope are a result of the much brighter scene at those elongations (image credit: WISPR collaboration)

WISPR Science Overview (Ref. 38):

• Wide-Field, Visible Light Imager : 13.5º - 105º from the Sun

• WISPR will map the morphology, velocity, acceleration and density of evolving solar wind structures when they are close to the Sun, as they approach, and then as they pass over the spacecraft

• Small density "blobs" accelerate to 30 Rs – is this occurring all the time?

• Shocks are seen ahead of CMEs – When do they form and are they the source SEPs (Solar Energetic Particles) ?

• Streamer belt has longitudinal density fluctuations – What is this due to?

Summary:

• WISPR will image:

- Slow and fast solar wind structures and fluctuations directly

- CMEs /Coronal Mass Ejections) and shocks and follow their propagation, evolution, and connection to the site of production of SEPs

• WISPR will measure electron density turbulence:

- Fast cadence readout mode to generate power spectral density to compare to in-situ observations of density and magnetic field spectral density.

• WISPR provides the links between the:

- Solar wind structure and SPP in-situ instruments

- Solar Orbiter and Solar Probe Plus missions.

 


 

ISIS-EPI (Integrated Science Investigation of the Sun - Energetic Particle Instruments)

The ISIS instrument is being developed at SwRI (Southwest Research Institute), San Antonio, TX, PI: David McComas. The objective of ISIS-EPI is to find the sources and acceleration mechanisms of solar energetic particles that are dangerous for human space explorers and can adversely affect our highly technology-based lives here on Earth. 39) 40)

ISIS-EPI will measure key properties of the accelerated particles ejected from the Sun. The ISIS-EPI low energy instrument measures the composition and intensities of protons and heavy elements as well as energetic electrons in multiple directions at the lower energies where the acceleration processes begin, while the ISIS-EPI high energy instrument measures the energy spectra, composition, and angular distributions of protons, heavy elements and electrons at the higher, more hazardous energies.

• Determine in both gradual & impulsive Solar EP events:

- Energy spectra

- Composition (electrons, protons, major heavy elements)

- Timing

- Pitch angle distributions

• Measure 3He as a key indicator of impulsive events

• Measurements of other populations (CIRs, ACRs, and GCRs) provide important new information on the radial dependences of these particles

ISIS Science Requirements and Performance:

ISIS provides comprehensive measurements of the energy spectra, anisotropy, and composition of suprathermal and solar energetic ions from ~0.02–200 MeV/nuc, as well as the energy spectra and arrival direction of ~0.025–6 MeV electrons (energy ranges indicate expected performance (Figure 40). The ISIS driving requirements were identified in the original ISIS proposal to NASA and have been vetted and agreed to through the development of various SPP Requirements documents. By combining measurements with other SPP instruments and from instruments on other spacecraft, ISIS addresses key questions concerning the origin, acceleration, and transport of different types of inner heliospheric particle populations, including those associated with CIRs (Corotating Interaction Regions), ACRs (Anomalous Cosmic Rays) and GCRs /Galactic Cosmic Rays). During the course of the SPP mission, ISIS will measure a sufficient number of impulsive and gradual SEP events (Figure 41) to meet all of its scientific objectives. Driving requirements for key instrument functional parameters are provided in Table 6; we note that while meeting these requirements is adequate to achieve all of our scientific objectives, the expected performance is even better.

Functional parameter

Measurement requirements

Energy range

e-: <0.05 to >3 MeV; p+/ions: <0.05 to >50 MeV

Energy binning (resolution)

>6 bins/decade

Cadence

e-: <1 s for selected electron rates; p+/ions: 5 s for selected ion rates

FOV (Field of View)

>π/2 ster coverage in both sunward and anti-sunward hemispheres, including coverage within 10 degrees of the nominal Parker spiral field direction at perihelion

Angular sectoring

e-: <45º sectors; p+/ions: <30º sectors

Composition

At least H, He, 3He, C, O, Ne,Mg, Si, Fe

Max intensity <1 MeV

>106 particles cm-2 sr-1 s-1

Max intensity >1 MeV

>5×105 particles cm-2 sr-1 s-1

Table 6: Measurement requirements for ISIS-EP suite

Figure 41 shows that the numbers of gradual and impulsive SEP (Solar Energetic Particle) events expected to be observable with EPI-Hi during the SPP prime mission, as a function of particle fluence and heliocentric radius, are more than sufficient to address all the science questions discussed above and to enable great discovery science on the SPP mission.

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Figure 40: Ion energy spectra of different inner heliospheric particle populations that SPP will encounter and the required energy range coverage for EPI-Lα (green) and EPI-Hi (blue) as well as the broader expected performance (lighter shades) for ISIS' overall energy coverage. Also shown is the energy range coverage for electrons (image credit: ISIS-EPI collaboration)

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Figure 41: Number of gradual (left) and impulsive (right) SEP events during the SPP prime mission inside a given heliocentric radius as a function of event size. These plots include the relative amount of time spent at each radial distance. The estimates for gradual events are based on NOAA/GOES data from 1976–2008, an assumed 11-year solar cycle, and a radial gradient of R-2.4 based on Lario et al. (2006). Estimates for impulsive 3He-rich events are based on 1998–2006 data from ACE/ULEIS, measurements of impulsive electron events by Wind/3DP (Wang 2010), and on an assumed fluence radial gradient of R-2 (image credit: ISIS-EPI collaboration)

 

ISIS Suite Overview:

The ISIS Energetic Particle Suite measures energetic electrons, protons, and heavy ions across a broad range of energies. To provide this wide coverage, ISIS includes two instruments that contain multiple sensors with detectors optimized for various parts of the energy measurement range. The suite combines EPI-Lo and EPI-Hi mounted together on the ISIS Bracket (Figure 42). ISIS is mounted at the aft end of the ram side of the SPP spacecraft providing an open FOV toward the direction of the nominal Parker spiral over much of the solar encounter phase, while staying within the umbra of the SPP TPS (Thermal Protection System).

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Figure 42: Illustration of the ISIS suite (image credit: ISIS-EPI collaboration)

EPI-Lo (Figure 43) measures energetic ions from 0.02 MeV/nuc to ~15 MeV total energy and energetic electrons from 25–1000 keV. To provide a large FOV with hemispherical coverage, EPI-Lo has eight wedges. Each wedge has 10 apertures that collimate energetic particles into pathways through the electrooptics, which allows their speed and energy to be measured while also registering which aperture they passed through. The eight sensor wedges are serviced by an electronics box that contains the Event Board and the Power Board. The Event Board contains the analog and digital processing circuits to record the events and communicates to the spacecraft through command and telemetry channels. The Power Board contains both the low voltage power converters as well as the high voltage power required for the sensors.

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Figure 43: EPI-Lo mechanical design (image credit: ISIS-EPI collaboration)

EPI-Hi (Figure 44) measures energetic protons and He nuclei from ~1 to ~100 MeV/nuc (and higher energies for heavier elements) and energetic electrons from ~0.5 to ~6 MeV. To cover this energy range, and to provide wide FOV coverage, EPI-Hi has three telescopes, a double-ended high energy telescope (HET), a double-ended LET1 (Low Energy Telescope), and a single-ended LET2 (Low Energy Telescope). These telescopes are mounted on the EPI-Hi Electronics Box, which contains an analog and digital processing electronics board for each telescope; a detector bias power supply; a digital processing unit for the instrument to coordinate its operations and communicate by command and telemetry channels with the spacecraft; and a low voltage power supply.

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Figure 44: EPI-Hi mechanical design (image credit: ISIS-EPI collaboration)

ISIS is located as far aft on the spacecraft body as possible, on the ram side, just inside the umbra line (Figure 45). This provides protection from direct solar heating, but still allows ISIS to view within 10º of the Sun-Probe line, thereby providing access to the nominal direction of the Parker Spiral magnetic field over much of the solar encounter phase.

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Figure 45: Location of ISIS on SPP spacecraft (image credit: ISIS-EPI collaboration)

FOV Maps: Sionce ISIS must remain in the umbra, its FOV has several blockages from the spacecraft itself. The TPS is the predominant FOV blockage (Figure 46 and 47). At the extremes of ISIS' FOV, there are small blockages due to the deployed solar arrays and the solar limb sensors. The EPI-Hi FOVs are a series of five overlapping 45º half-angle cones, three of which are provided by the low energy telescopes (the double-ended LET1, and the single-ended LET2). The other two 45º half-angle cones are provided by the double-ended, HET (Higher Energy Telescope). The EPI-Hi telescope FOVs are overlapped to provide full energy coverage in the sunward and anti-sunward direction in two, nearly-complete 45º half-angle cones. Further coverage is provided in the directions that are not blocked by the spacecraft by the low energy telescopes alone. The EPI-Lo FOV is an array of 80 apertures, ten on each of eight wedges. These apertures are arranged to sample a hemispherical FOV that includes viewing in the sunward and anti-sunward hemispheres as well as coverage near the direction of the nominal Parker spiral.

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Figure 46: EPI-Hi FOV Map. Three blue diamonds indicate the locations of the average Parker spiral magnetic field for a solar wind velocity of 400 km/s at heliocentric distances of 0.05, 0.25, and 0.7 AU (left to right), image credit: ISIS-EPI collaboration

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Figure 47: EPI-Lo FOV Map. Three blue diamonds indicate the locations of the average Parker spiral magnetic field for a solar wind velocity of 400 km/s at heliocentric distances of 0.05, 0.25, and 0.7 AU (left to right), image credit: ISIS-EPI collaboration

ISIS Bracket: The ISIS suite is integrated onto a single, combined ISIS bracket. The bracket is designed to allow flexibility and position the ISIS instruments as close to the allowable umbra-line as possible since the final position of the SPP TPS may be shifted slightly during spacecraft I&T (Integration and Testing). The nominal bracket design holds ISIS just within the allowed limit with ~2° safety margin to the actual umbra line, which keeps the ISIS FOVs as close as possible to the Sun-Probe line. However, the SPP spacecraft center-of-mass must be very close to the center-of-pressure on the TPS in order to minimize pointing perturbations. To do this, the TPS, which is initially oversized, will be trimmed to its final configuration during spacecraft I&T. By design, the TPS will not be trimmed below the limit we have designed to, but it could be trimmed less, which would move the actual umbra line farther away from ISIS and reduce viewing toward the Sun-Probe line. To prevent this reduction in our FOV in this key region of interest, the ISIS bracket is designed to accommodate a late TPS modification by extending the suite farther away from the spacecraft panel, thus restoring ISIS to the position as close as possible to the umbra.

Mass, Power, and Telemetry: ISIS has been optimized for the SPP mission. Mass is the most constrained resource for the SPP spacecraft, and power is second because of the mass penalty for cooling the solar arrays. The ISIS team has worked diligently to remain within limits set by the mission design. The resource estimates come from a detailed Master Equipment List in which the estimates for each item are carefully maintained and revised over the course of development. Given the heritage of the electronics, the power and telemetry estimates have remained steady during the course of the project development to date. To minimize overall SPP risk, the ISIS team was given a small mass increase to reduce development risks during the risk-reduction SPP activities in Phase B. The CBEs (Current-Best-Estimates) as of the time of the ISIS and mission PDRs, as well as uncertainty estimates are listed in Table 7.

Parameter

EPI-Hi

EPI-Lo

ISIS bracket

ISIS

CBE

Uncertainty

Total

CBE

Uncertainty

Total

CBE

Uncertainty

Total

Total

Mass (kg)

3.628

0.692

4.320

3.435

0.656

4.091

0.817

0.156

0.973

9.384

Power (W)

5.810

0.960

6.770

4.170

0.830

5.000

NA

NA

NA

11.770

Operational heaters (W)

0.480

0.070

0.550

NA

NA

NA

NA

NA

NA

0.550

Survival heaters (W)

3.810

0.570

4.380

2.450

0.370

2.820

NA

NA

NA

7.200

 

EPI-Hi

EPI-Lo

ISIS

Raw

Burst

Total

Raw

Burst

Total

Total

Uncompressed data volume [Gbit/orbit]

3.660

0.000

3.660

11.320

0.200

11.520

15.180

Compressed data volume (75%) [Gbit/orbit]

2.745

0.000

2.745

8.490

0.200

8.690

11.435

Packetized data volume (105%) [Gbit/orbit]

2.882

0.000

2.882

8.915

0.210

9.125

12.007

Table 7: ISIS resource table

Thermal Design: On a spacecraft that is designed to fly much closer to the Sun than any other mission in history, one might reasonably expect that the thermal environment would be extreme and that the limiting hot-case for the thermal design would be at perihelion. However, the limiting hot-case for the thermal design of ISIS is actually near 1 AU, just after launch when the SPP spacecraft has to perform various maneuvers that allow direct solar illumination of ISIS. After those early operations, ISIS remains in the umbra of the TPS at all other times, which provides a stable environment that is "in-family" with past, heritage missions for this type of instrumentation.

The thermal design of the SPP mission requires that the instruments be thermally isolated from the SPP spacecraft. The ISIS bracket configuration provides thermal isolation from the spacecraft deck, from the instruments, between the instruments, and from the bracket, thermally isolating the instruments themselves. This is accomplished by including 1.27 cm ULTEM® spacers at all of the mounting bolt locations on both sides of the bracket. To provide electrical conductivity, while still maintaining thermal isolation, thin straps of copper bridge the gap. These provide a very small cross-section to minimize thermal conductivity, but have a large surface area for good, high-frequency electrical grounding. MLI (Multi-Layer Insulation) covers the majority of ISIS, except for the apertures and thermal radiators, to minimize radiant heat transfer. EPI-Hi and EPI-Lo have independent survival heaters controlled by the spacecraft, and EPI-Hi also has an operational heater controlled by the instrument itself.

Electrical Interfaces: EPI-Hi and EPI-Lo have completely independent electrical interfaces to the SPP spacecraft. Each has separate command and telemetry interfaces to the SPP C&DH (Command and Data Handling) unit as well as separate power interfaces for instrument power to the SPP PDU (Power Distribution Unit). The SPP spacecraft also provides independent temperature sensors and survival heaters that are used to control the EPI-Hi and EPI-Lo temperatures.

Each instrument has separate A-side/B-side command and telemetry interfaces. LVDS (Low Voltage Differential Signaling) is used to establish a high-speed serial communication protocol with ITFs (Interface Transfer Frames) as the packet format. Command ITFs are used to send spacecraft time and status as well as instrument-command packets to each instrument. ISIS uses the spacecraft time and status information to configure itself autonomously. The instruments return instrument-status and telemetry packets in Telemetry ITFs. The instrument-status can be interpreted on-board the spacecraft to make real-time autonomy decisions. The telemetry packets are stored on the SPP data recorder for downlink to ground stations at the appropriate time. To maintain time synchronization, a 1 PPS (Pulse-Per-Second ) signal is provided by the spacecraft. Instead of using a separate line, a "virtual PPS" is provided by carefully controlling the timing of the start-bit of the first byte of the Command ITFs.

EPI-Hi and EPI-Lo have independent power connections to the SPP PDU, each with their own switches. At the SPP mission level (including measurements from the other instruments), mission success can be achieved with either EPI-Hi or EPI-Lo, so additional power redundancy is not required. The spacecraft provides a standard +28 V power bus. ISIS uses a common LVPS (Low Voltage Power Supply) design, which meets the EMI/EMC (Electromagnetic Interference/Compatibility) requirements of the SPP mission. The LVPS design is customized for each instrument and provides reverse voltage protection, in-rush current limiting, EMI filtering, and isolation between the spacecraft power return and the ground returns of the secondary power supply rails that are developed within each instrument.

Integrated Science Team: The ISIS science team includes scientists from the ISIS Leadership, EPI-Hi, EPI-Lo, and SOC groups as well as Senior Science Mentors and theory and modeling team members who are leading experts in energetic particle science (Table 8). These scientists are integrated into the on-going instrument development process as well as the operations planning and data analysis efforts. The ISIS science team works in close coordination with the other SPP instrument investigation teams. It also coordinates data collection and analysis with other missions (e.g. Solar Orbiter) and ground-based observations. This close coordination inside the ISIS science team and cooperation with the broader science community provides a truly outstanding integrated science investigation of the Sun.

Title

Name

Organization

PI (Principal Investigator)

Dave McComas

SwRI (Southwest Research Institute)

DPI (Deputy Principal Investigator)

Eric Christian

NASA/GSFC

Co-PI

Alan Cummings

Caltech

Co-PI

Mihir Desai

SwRI

Co-PI

Joe Giacalone

University of Arizona

Co-PI

Matthew Hill

JHU/APL

Co-PI

Stefano Livi

SwRI

Co-PI

Bill Matthaeus

University of Delaware

Co-PI

Ralph McNutt

JHU/APL

Co-PI

Dick Mewaldt

Caltech

Co-PI

Don Mitchell

JHU/APL

Co-PI

Nathan Schwadron

UNH (University of New Hampshire)

Co-PI

Tycho von Rosenvinge

NASA/GSFC

Co-PI

Mark Wiedenbeck

NASA/JPL

SSM (Senior Science Mentor)

Robert Gold

JHU/APL

SSM

Stamatios Krimigis

JHU/APL

SSM

Edmond Roelof

JHU/APL

SSM

Ed Stone

Caltech

Table 8: ISIS Science Team members

 

EPI-Lo (Energetic Particle Instrument—Low Energy):

EPI-Lo measures energetic particles in the lower portion of the ISIS energy range. The two ISIS instruments complement one another in their energy ranges and their sky coverage in order to obtain the comprehensive set of observations needed to understand solar energetic particle sources, acceleration and transport close to the Sun.

EPI-Lo Overview: EPI-Lo is a novel, light-weight, high-heritage, TOF (Time-of-Flight) based, mass spectrometer that measures energetic electron (25–1000 keV) and ion (~0.04–7 MeV for protons and ~0.02–2 MeV/nuc for heavier ions) spectra and resolves all major heavy ion species and 3He and 4He over much of this energy range in multiple directions. ISIS thus covers the critical energy range from suprathermal energies (∼20 keV/nuc) up to the lower portion of the EPI-Hi energy range with a single instrument. The EPI-Lo characteristics and projected performance are summarized in Table 9.

Parameter

Required

Goal (expected)

Comment/Heritage

Electron energies

50–500 keV

25–1000 keV

Electron capability from JEDI, RBSPICE

Ion energies

50 keV/nuc–15 MeV Total E

~20 keV/nuc–15 MeV Total E
~85 MeV Total E for Fe

Capability based on that of RBSPICE
Maximum energy ~1.5 MeV/nuc for Fe

Energy resolution

45% for required energy range

11% for required energy range

Telemetry limited

Time sampling

5 s

1 s

Telemetry and/or statistics limited

Angle resolution

<30º x <30º

Ions, ~15º x 12º to <30º x <30º e-, 45º

Varies with elevation

Pitch Angle (PA) coverage

0º–90º or 90º–180º, some samples in both hemispheres

0º–90º or 90º–180º,some samples in both hemispheres

 

Time for full PA

1-5 s

1-5 s

Telemetry limited

Ion composition

H, 3He, 4He, C, O, Ne, Mg, Si, Fe

H, 3He, 4He, C, O, Ne, Mg, Si, Fe

3He/4He ~50 to 1000 keV/nuc

Electron sensitivity, geometric factor, counting rate, and background drivers

<106 cm-2 s-1 sr-1 (G ~ 0.05 cm2 sr; measure event rates to >50 kHz)

102–107 cm-2 s-1 sr-1 (G>0.05 cm2 sr; measure event rates to > 700 kHz; background rate <1 Hz)

G = Geometric factor (cm2 sr) 8 pixels/sensor; background rate is spectral-slope dependent

Ion sensitivity, geometric factor, counting rate, and background drivers

101–106/cm-2 s-1 sr-1 (G ~ 0.05 cm2 sr; measure event rates to >50 kHz; at minimum intensity require accidental rate <~50 kHz)

100–107/cm-2 s-1 sr-1 (G>0.07 cm2 sr; measure event rates to >700 kHz; at minimum intensity require accidental rate <~5 kHz)

80 pixels/sensor

Table 9: EPI-Lo instrument required and projected performance

EPI-Lo (Figures 48 and 49) consists of eight sensor wedges mounted above an electronics box. It has 80 separate entrances (10 on each of eight wedges) densely sampling over half of the sky. This configuration permits full angular distributions without articulation or duty cycle, and allows for measuring the first-arriving, field-aligned ions at the spacecraft for a broad range of vector magnetic (B) field directions.

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Figure 48: EPI-Lo comprises eight wedges mounted on a common electronics box. Ions generate start/position electrons as they transit thin foils in the entrance apertures, and then strike a stop foil and solid-state detector (SSD), yielding angle and E x TOF. Energetic electrons enter the same apertures as the ions, and are detected in a second set of SSDs located behind light- and ion-rejecting cover foils (image credit: ISIS-EPI collaboration)

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Figure 49: Representative species separation for EPI-Lo holes 2 (22.5º from the instrument normal) and 5 (the holes about the instrument perimeter, 90º from instrument normal) based on measured TOF resolution performance on test model. Data taken using an alpha source in hole 2 is overlaid on the hole 2 simulated data. 3He is well distinguished from 4He at 1:100 abundance ratio to ≥1.5 MeV/nuc. The best performance for species identification is in the 64 holes at the 3, 4 and 5 positions, at 45º, 67.5º, and 90º from the instrument normal (image credit: ISIS-EPI collaboration)

Understanding coronal acceleration processes requires ion-mass resolution sufficient to measure separately the intensities of the isotopes 3He and 4He (Figure 49), and the elements C, O, Ne, Mg, Si, and Fe, while simultaneously providing angular coverage with at least 45% energy resolution.

EPI-Lo achieves this resolution, including separation of 3He from 4He, over most of its angular coverage from 200 keV to 4 MeV total particle energy. EPI-Lo also returns high-resolution energy spectra (64 log-spaced energy bins) at reduced time resolution (necessitated by downlink limitations) and measures the required ion composition and pitch angle distributions from 30 keV/nuc to 0.3–1.0 MeV/nuc every 5 or 30 s. In this way ion acceleration histories are completely characterized with no uncertainty owing to an insufficient energy range or composition misidentification. EPI-Lo rejects background by requiring coincidence between the start and stop pulses for the TOF measurement, along with an energy measurement between appropriate TOF time gates. This rejection is non-linear, and is very effective (resulting in inconsequential rates of false valid events) for background counting rates in individual detectors (singles rates) below ~106 s-1; projected singles rates from combined background and foreground sources are ≤3 x 105 s-1. Incoming ion velocities are determined by measuring the TOF between two thin (100 nm Start, 65 nm Stop) carbonpolyimide-aluminum (and palladium, Start only) foils. An ion passing through each of the foils produces secondary electrons, which are deflected toward a MCP (Microchannel Plate) producing "Start" and "Stop" pulses. The ion entrance angles are determined from the position where the Start electrons strike the MCP and are unique for each entrance foil location. The ion energy deposited in the SSD, together with velocity from the TOF, determines species through an onboard table lookup.

EPI-Lo also detects electrons from ~25 keV to 1000 keV. EPI-Lo uses SSDs (Solid State Detectors) shielded by aluminum flashing of ~2 µm thickness as also used in multiple current and upcoming missions, e.g., Cassini-MIMI, MESSENGER EPS, New Horizons PEPSS, Juno JEDI , Van Allen Probes RBSPICE, and MMS EIS. The relatively thick aluminum flashing naturally suppresses light, which is a very important feature for this intrinsically single parameter measurement. While the primary electron measurement does not identify which entrance aperture an electron enters through, each EPI-Lo wedge contains independent electron SSDs, so the sector of the sky is identified over an angular coverage similar to that for the ions. For the small subset of the electrons that produce secondary electrons as they transit the entrance foils, an additional mode that requires a Start pulse along with the SSD pulse identifies the specific entrance aperture that the electron entered through. This mode also provides better background rejection.

MCP detectors and SSD's have been optimized for use on this mission. We also produced a prototype wedge (Figure 50), which has been used for extensive testing and development of the EPI-Lo concept and design. Finally, lead development engineers and instrument scientists on JEDI, RBSPICE, and EIS are all participating in the EPI-Lo effort, as well the corresponding leads that developed and continue to operate PEPSSI.

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Figure 50: EPI-Lo wedge prototype (image credit: ISIS-EPI collaboration)

EPI-Lo Expected Performance: The detectors in EPI-Lo are sensitive to charged particles and to light. The MCPs are sensitive to ultraviolet (UV) light, including both EUV and shorter FUV (Far Ultraviolet) wavelengths whereas the SSDs are sensitive to everything from X-rays to visible wavelengths. However, the mechanisms are different. MCPs respond to charged particles impacting their surface, and UV produces secondary electrons (photoelectrons) on surfaces (either the surface of the MCP itself, or the Start and Stop foils designed to produce secondary electrons as ions penetrate them, or other surfaces inside the sensor volume), which can then enter channels and be amplified similarly to signals from any other particle.

SSDs respond to particles as they deposit energy in the bulk material of the detector. They free bound electrons to create electron-hole pairs and so increase conductivity. In SSDs, light can also energize bound electrons sufficiently that they become free and increase conductivity. Energetic X-rays may deposit enough energy to create free electrons that then raise the energy of sufficient numbers of bound electrons so that the resulting current spike is above threshold. Lower energy photons will not do so, but enough of them arriving within an SSD event time constant (~1 µs can free sufficient numbers of electrons to register as events, or to raise the electronic noise of the detector unacceptably.

In both cases, light (both UV and visible) must be reduced to a level where these effects cannot degrade the particle measurements. Controlling the amount of light entering the sensor volume involves a variety of techniques: (1) eliminating stray-light leak paths into the instrument; (2) collimating the entrance apertures to eliminate exposure to trajectories that are not useful for particle analysis; (3) coating surfaces from which light may be reflected into the detectors with low-reflectance coatings; (4) using high-work-function surfaces in locations where UV photons may strike surfaces from which photoemission is undesirable; (5) designing the Start and Stop foils using materials and thicknesses that reflect and/or filter the UV and visible light, reducing the light that can directly enter the sensor volume to acceptable levels; (6) taking precautions against failure of this filtering approach as a consequence of small numbers of pinholes in the foils (either from dust impact or launch vibration damage); and, (7) designing the sensor timing constraints such that controlled rates of photoelectron production from UV photons do not impact particle measurements.

Briefly, the instrument design carefully considers: (1) stray light paths and avoids them; (3) coating surfaces as a second-order consideration relative to direct paths for light and UV; and (7) timing constraints fundamental to the instrument design: an approach successfully used on numerous past flight programs (the "hockey pucks"1 on MESSENGER, New Horizons, Van Allen Probes, Juno, and MMS as well as the ENA imagers on Cassini and IMAGE).

All of these programs also rely on collimation and filtering by foils, and all were subject to pinhole degradation at various levels, so these considerations are not new; however, the SPP environment is sufficiently unique that careful consideration to these approaches is required for EPI-Lo.

The environment in which EPI-Lo will operate is one never before experienced. It has been extensively modeled, and there is considerable knowledge of the light environment. The SPP team has provided their best estimate of the worst-case photon environment as a function of solar elongation angle at perihelion. This, and the other expected SPP environments are specified in an internal Project-provided environmental design and test requirements document. Light at the ISIS location is dominated by Thomson-scattered photons, to first order a process that preserves the shape of the solar spectrum. The intensity of the scattered light is a fairly strong function of elongation angle, dominated primarily by the line-of-sight integrated electron density, which varies with elongation (Figure 51). While dust-scattered light dominates at large elongations it is much weaker than the Thomson-scattered light at smaller elongations, so the foil filtering design required by the Thomson-scattered light at smaller elongations is more than adequate for the dust-scattered light.

We know the elongation angles of the various apertures on EPI-Lo. In considering these, we treat the EPI-Lo apertures in two distinct groups: those that contain sky elements within ~12° elongation angle (the edge of the TPS is at about 8°), and those that do not. For those with elongation angles less than 12°, the rough order-of-magnitude (ROM) for the UV and light environment is approximately equivalent to having the disk of the Sun directly in the FOV of that aperture at 1 AU. This then means that the filtering requirements for those apertures are approximately the same as what would be required for EPI-Lo at 1 AU with the Sun directly in the aperture FOV.

Several approaches contribute to determining the required filtering properties of the foils. For visible light (which primarily concerns the SSDs), we have used flight experience from various instruments in Earth orbit. For example, the ISEE-1 EPD (Energetic Particle Detector) had 30 µg/cm2 of Al covering its SSD, and suffered from visible light contamination when the detector viewed the Sun directly. The IMP-8 EPD had 40 µg cm-2 of Al covering its SSD and did not respond to direct Sun in the visible, but it did respond to solar X-ray events. From this we conclude that 40 µg cm-2 of Al or the equivalent would be sufficient for EPI-Lo for elongations <12°. For these small elongation foils we are baselining a Start foil of 24 µg cm-2 of Al plus 18 µg cm-2 of palladium (Pd), which combined with the Stop foil containing 7.3 µg cm-2 of Al, yields a predicted noise level on the SSD of ~0.04 keV µs-1, well below the electronic noise level of ~7 keV µs-1.

For the larger elongation apertures we have baselined a total of 7.3 µg cm-2 of Al plus 18 µg cm-2 of Pd on the Start foil, and 7.3 µg cm-2 of Al on the Stop foil. The predicted noise level for this combination is 0.7 keV µs-1 in the SSD. We plan to test at this level, and increase the thickness of the aluminum layers if necessary. It should be noted that these are ROM estimates and not really directly comparable. The Thomson scattered light is diffuse, whereas sunlight at 1 AU is collimated. Appropriate descriptions in terms of photons at a particular wavelength are number of photons cm-2 s-1 for sunlight at 1 AU but number cm-2 s-1 sr-1 for the scattered light at 10 RS (~perihelion for SPP). The way we employ our model is by using the EPI-Lo geometric factor for the apertures, which lie within a given elongation angle range we are modeling, and calculating the number of photons s-1 that hit the aperture foils, so the calculation is correspondingly conservative. We then run the filter model with a fraction of the full Sun at 1 AU determined such that we match that number of photons/s on the same aperture foils.

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Figure 51: Intensity of Ly-α EUV vs. solar elongation. Most EPI-Lo entrances include angles no closer than 20 degrees elongation (which implies a maximum Ly-α intensity of ~109 cm-2 s-1 sr-1). For four apertures close to the TPS, the maximum is~1012 cm-2 s-1 sr-1, but average only ~1010 cm-2 s-1 sr-1 (image credit: ISIS-EPI collaboration)

EPI-Lo Electrical: The EPI-Lo electronics box contains all the electronics to run the instrument other than the energy and timing preamps, which are located in the sensor wedges. The box contains two octagonal boards mounted into metal frames. The boards stack one on top of the other, with an internal connector providing electrical interconnects between the boards. The functional breakup between the two boards minimizes the number of interconnects needed. See Figure 52 for the block diagram.

The electronics are designed to handle solar energetic particle event intensities up to at least 106 particles cm-2 sr-1 s-1. This includes handling electron counts rates ≥70,000 counts per second and total ion count rates (SSD and MCP valid coincident event) of ≥70,000 counts per second. In both cases, these events can be either evenly distributed over the entire instrument or concentrated in one wedge. The term "handle" is used to mean that the incoming particles are processed in the instrument such that the particle types, directions, and rates can be determined. Note that ground software rate correction will be necessary when rates are sufficiently high (e.g. ≥106 starts per second or ≥40,000 total ion count rates per second). With the current mission and environment assumptions, we expect that rates will be sufficiently high only ~5 % of the time to require such corrections.

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Figure 52: EPI-Lo block diagram (image credit: ISIS-EPI collaboration)

Event Board: An ion entering the sensor through one of the collimator apertures will deposit energy in the SSD and produce secondary electrons in the Start and Stop foils, which are amplified by the MCP and collected in distinct positions on the anode. This collection of time-correlated signals is called an event for the purpose of the description of the electronics that follows. The Event board directly processes the sensor SSD and anode preamp output signals, and contains all necessary analog and digital circuitry to process and store event information on an event-by-event basis. The energy signals from the eight SSD preamplifier sets are processed in parallel peak-detect/discriminator circuit/ADC chains. The MCP anode signals are processed via constant-fraction discriminators (CFDs) and time-to-digital conversion (TDC) circuitry; these measured time differences are converted into event look direction and particle velocity in the FPGA (Field-Programmable Gate Array). FPGA-based event logic also determines which signals comprise valid ion and electron events and coordinates all event hardware processing timing. A soft-core processor (i.e. a processor implemented in VHDL) is also embedded in the FPGA to provide all command, control, telemetry, and data processing functions of the instrument. SRAM, MRAM, and PROM memory storage is provided on the board to support the processor.

EPI-Lo uses a pulse-width technique to determine the energy deposited in each detector for energies above ~1 MeV. This method, used on the Juno JEDI, the RBSPICE, and MMS EIS, allows the energy dynamic range to be extended from ~1.5 MeV to a total energy ~15–20 MeV. In order to cover fully iron composition with no gaps between EPI-Lo and EPI-Hi, the maximum energy measured will be extended to ~85 MeV for iron (1.5 MeV/nuc for Fe nuclei). Three separate approaches to this extension of the iron energy range have been identified, and the preferred approach will be finalized and tested in the engineering model prior to instrument CDR. Each sensor uses an existing, flight-qualified, ASIC (Application-Specific Integrated Circuit) containing preamplifier/shaper circuits to amplify the SSD and APD (Analog Peak Detect) signals, shape the pulse, and generate timing triggers on the rise and fall of each pulse. These signals feed into the EPI-Lo FPGA where the coincidence logic and other digital processing begin. The EPI-Lo FPGA-based processor accumulates events into rates, and packetizes the telemetry products.

Power Board: The Power board contains both the low and high voltage power supplies. The low voltage portion takes spacecraft primary power on a single 9 pin connector and generates 1.5 V (for the FPGA core), 3.3 V (primarily for digital interface logic, memories, and TDCs), and 5 V (primarily for analog functions). A 15 V output powers the high voltage power supply (HVPS). The HVPS generates the necessary high voltage outputs for the sensor MCP and electron optics, with a maximum voltage output of 3300 V. It can independently control high voltage to each of the four quadrants. A bias supply provides up to 250 V for the SSDs.

EMI/EMC Design Considerations: Of principal concern for EMC design are the power supplies. These are controlled to a frequency window centered at n x 50 kHz with n ≥3 and 500 ppm wide overall operating conditions and time. The LVPS is synchronized to 200 kHz by a 400 kHz clock provided by the digital boards. EPI-Lo has a 40 MHz oscillator and EPI-Hi has a 58.8 MHz oscillator; both evenly divide to 400 kHz. To minimize interference, transformers and large inductors are placed as far from the box walls as possible, and stable currents are employed to minimize changes in magnetic emissions. EPI-Lo does have nickel grids, and concerns associated with those are mitigated with careful handling, use of non-magnetic tools, and testing.

All electronic parts are selected for proven radiation tolerance: TID (Total Ionizing Dose) >100 krad, no latch-up, and acceptable SEU (Single-Event Upset) performance. TMR (Triple Module Redundancy) and ECC (Error Correction Code) are used on vulnerable registers inside the FPGA. PROM-based boot code is used to ensure reliable memory loading and checking capabilities. Analog parts are selected with low dose effects in mind. Parts comply with GSFC's (Goddard Space Flight Center) "EEE-INST-002" Level II requirements.

EPI-Lo Mechanical: EPI-Lo uses a symmetric design to enable the wide field of view in a compact, low-mass package. Parts consist of the eight wedge assemblies and their closeouts, a top close out, the "spider" frame, which holds the wedges, and the event- and power-board slices, which comprise the components of the electronics box (Figure 53). The common wedge design is shown in Figure 26. Each contains an MCP assembly, an SSD assembly, and a collimator set. Alignment of all of the pieces enables the coverage and functionality of the detector overall. Preliminary design analysis has been done with a FEM (Finite Element Model) using 89,077 nodes and 70,645 elements.

The preliminary structural analysis of the baseline design was performed using MSC Nastran, MAYA SATK®, and Femap for analysis. The model was simplified wherever possible to reduce solution time. PWAs (Printed Wiring Assemblies) were modeled as plate elements with uniform stiffness, thickness, and density, and the instrument model was oriented to the ISIS bracket configuration in relation to the spacecraft panels.

The project performed a mechanical modal analysis to determine mechanical resonances. The analysis showed the first mode to be 304 Hz (for the Event PWA) and 553 Hz for the overall instrument. The analysis environmental input levels per spacecraft requirements were performed for all three orthogonal axes relative to the spacecraft panel. The 3σ acceleration random response enveloped the static load requirement as desired for EPI-Lo displacements, stresses, and forces. The random-vibration PWA displacement response may be relatively high for EEE (Electrical, Electronic, and Electromechanical) part solder/lead-wire fatigue resistance, and further analysis is planned after EEE parts placement is finalized. All margins of safety are positive for the current model configuration, and a detailed analysis will be carried out for the flight configuration to confirm that the flight design has positive margins and meets the minimum frequency requirement.

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Figure 53: EPI-Lo mechanical assembly (image credit: ISIS-EPI collaboration)

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Figure 54: Common wedge design (image credit: ISIS-EPI collaboration)

 

EPI-Hi (Energetic Particle Instrument—High Energy):

EPI-Hi measures energetic particles in the upper portion of the ISIS energy range. The two ISIS instruments are complementary in their energy range and sky coverage in order to obtain the comprehensive set of observations needed to fully understand solar energetic particle sources, acceleration, and transport close to the Sun.

EPI-Hi Overview: EPI-Hi measures energetic electrons, protons, and heavy ions in the MeV energy range using the dE/dx versus total energy technique in a sensor system based entirely on ion-implanted silicon SSDs. It builds on a heritage of more than 40 years of SSD-based energetic particle instruments, and most directly on the LET (Low Energy Telescope) and HET (High Energy Telescope) that are part of the IMPACT instrument suite on the twin STEREO spacecraft; these instruments have been providing multi-point measurements of energetic particles in the heliosphere since late 2006. The sophistication of SSD instruments has steadily advanced as capabilities have improved with high performance, low power, and miniaturized, front-end electronics in the form of ASICs. The EPI-Hi requirements and expected performance are summarized in Table 10.

Parameter

Required

Goal (expected)

Comment/Heritage

Electron energies

0.5–3 MeV

0.5–6 MeV

STEREO/HET

Ion energies

~1 to ≥50 MeV/nuc

p, He: 1 to 100 MeV/nuc
Z ≥ 6: 1.5 to >100 MeV/nuc

STEREO/LET & HET

Energy binning

≥6 bins per decade

12 bins/decade

Logarithmic bins

Cadence

Fastest: e 1 s, p 5 s

Fastest: e 1 s, p 1 s
most data products: 1 min
angular distributions: 5 min

Large energy bins
best energy resolution

FOVs (Fields of View)

≥π/2 steradians in sunward and
anti-sunward hemisphere

Five 45º half-angle view cones covering sunward and anti-sunward hemispheres

View cones overlap to provide full energy coverage near Parker spiral to within 10º of spacecraft-Sun line

Angular sectoring

e: ≤45º; ions: ≤30º

45º half-angle cones subdivided into 25 overlapping sectors

Overlapping sectors provide improved angular resolution for deriving pitch angle distributions

Elemental composition

H, He, C, O, Ne, Mg, Si, Fe

H, He, C, N, O, Ne, Na, Mg, Al, Si, S, Ar, Ca, Cr, Fe, Ni

STEREO/LET

Isotopic composition

3He/4He

3He/4He

In selected viewing directions STEREO/LET

Intensity range

Up to 3 x 106 protons cm-2 sr-1 s-1

Normal mode: up to ~1x106 protons cm-2 sr-1 s-1 Pixel mode: up to ~4 x 107 protons cm-2 sr-1 s-1

>10 MeV protons

Geometrical factor

N/A

5 view cones, each with AΩ~0.5 cm2 sr

Value at energy with maximum AΩ

Table 10: EPI-Hi instrument required and expected performance

Figure 55 shows a block diagram of EPI-Hi. The sensor system consists of three detector stacks, commonly called "particle telescopes", each controlled by a dedicated electronics board that provides front-end analog signal processing, PHA (Pulse Height Analysis), coincidence determination, data accumulation and analysis, and formatting for readout. We have tailored the telescope designs for measurements in two different but overlapping energy ranges. A pair of LET1 (Low Energy Telescopes) and LET2 measure protons and heavy ions from ~1 to >20 MeV/nuc and electrons from ~0.5 to ~2 MeV while a single HET (High Energy Telescope) extends the energy coverage for electrons up to ~6 MeV and for protons and heavy ions to at least 100 MeV/nuc. LET1 and HET analyze particles incident from both ends of their detector stacks while LET2 analyzes particles incident from only one end because the other end is blocked by the spacecraft. Each of the five active telescope ends has an angular acceptance cone with a half-angle of 45º centered on the symmetry axis of the telescope.

In addition to the three telescope boards, EPI-Hi includes a DPU (Data Processing Unit) board that is responsible for the overall control of the instrument, including managing the boot-up of processors on the telescope boards, collecting data from those boards and merging them into a single data stream that is passed to the spacecraft, and receiving commands from the spacecraft and passing them to the processor for which they are intended. The remaining two electronics boards contain the LVPS that produces the required regulated, filtered voltages needed by the instrument and the bias supply that produces programmable voltages for biasing SSDs.

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Figure 55: EPI-Hi block diagram (image credit: ISIS-EPI collaboration)

A metal enclosure, designated the "E-Box", holds the electronics boards and provides a degree of shielding against radiation and high-speed dust impacts. We also use the E-Box, which is attached to the ISIS bracket, as a platform on which we mount the detector telescopes in the orientations needed to achieve the required viewing directions, as illustrated in Figure 56.

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Figure 56: EPI-Hi mechanical configuration including E-Box and telescopes (image credit: ISIS-EPI collaboration)

HET (High-Energy Telescope): Figure 57 shows cut-away views of the telescopes. HET has a thick central stack of detectors together with a single detector spaced farther away at each end. A coincidence between this front detector and the central stack defines the FOV of the telescope. The front detectors (H1) and the outermost detectors at each end of the central stack (H2) have five main active elements, consisting of a central bull's-eye and four quadrants of a surrounding annular region, with all elements having equal areas, as illustrated in Figure 30a. Identifying which detector segment was hit by an incident particle in each of the two detectors allows us to subdivide the FOV into 25 overlapping sectors, thus enabling measurements of the distribution of particle incidence directions.

The detectors other than H2 in the central HET stack have a central segment with an active area equivalent to the combined areas of the five segments in H1 and H2. The combination of the thickness of the central stack and the active diameter of the detectors is such that some particles incident at large angles from the telescope axis will exit through the sides of the stack. To identify and reject these side-exiting particles, the detectors include an annular "guard" segment that is used as anti-coincidence in the analysis. The guard also allows identification of particle trajectories that clip the detector edge resulting in reduced energy-loss signals. This feature is important because the detector thickness is a non-negligible fraction of the detector diameter and thus edge-clipping trajectories are not improbable. The HET detectors (as well as the LET detectors other than L0 and L1) also include one additional segment in the form of a small pixel that is used to extend the dynamic range in particle intensities to values above which the larger detector segments will saturate.

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Figure 57: Cut-away illustration of the EPI-Hi telescope designs and detector naming (image credit: ISIS-EPI collaboration)

The individual silicon detectors used in the HET telescopes have thicknesses of either 500 µm (H1) or 1000 µm. However, for H3, H4, and H5, we use two successive 1000 µm detectors connected to the same front-end electronics to produce what is effectively a 2000 µm detector without requiring the high bias voltage that an individual detector of this thickness would need.

LET (Low-Energy Telescope): The conceptual design of the LETs is similar to that of HET, using a central detector stack in conjunction with position-sensitive detectors that define the FOV and subdivide it into a number of sectors that we use for measuring particle angular distributions. Similar to HET, the LET detectors, other than L0 and L1, have thicknesses of either 500 µm (L2) or 1000 µm. In order to achieve a low threshold energy for EPI-Hi and to minimize the energy gap between EPI-Hi and EPI-Lo, we have developed a process for fabricating very thin silicon detectors. At the front of LET we have L0 and L1 detectors with thicknesses of 12 µm and 25 µm, respectively (Figure 57). A 1 MeV proton has energy just sufficient to penetrate the L0 detector and the thin windows in front of the telescope (~3 µm silicon equivalent) and provide the 2-parameter measurement required for particle identification. The thin detectors are sufficiently uniform in thickness to allow the required species separation, including distinguishing between the isotopes of He. Like the H1 and H2 detectors, L0, L1, and L2 are all position sensitive with a central bull's-eye surrounded by an annulus subdivided into four quadrants. Figure 58b shows a photograph of a prototype L1 detector.

Protons with energy greater than ~8 MeV will have energy losses in the L0 detector that fall below the detection threshold of the front-end electronics. For this reason, in LET we identify events either by a coincidence between L0 and L1 or a coincidence between L1 and L2. It is possible for an event to satisfy the L1•L2 coincidence without triggering L0 either because the signal fell below threshold in L0 or because the particle trajectory did not pass through the active area of this device. We have designed the L0 detector to have its 1 cm2 active area located at the center of a large (~9 cm2) thin silicon membrane that covers the entire field of view defined by the L1•L2 coincidence. This ensures that the analyzed particles will have passed through a consistent amount of material, thereby enabling accurate calculations of their incident energies. The L2 detector has a large annular guard region so that we can determine whether an event defined by a coincidence between L0 and the center element of L1 but lacking a signal in the central five segments of L2 is due to a particle stopping in L1 or to a trajectory passing outside the central part of L2.

Unlike in HET, the central stack detectors in LET (L3 and L4) have an active diameter large enough to intercept even the widest-angle trajectories defined by the L1•L2 coincidence. Thus, the LET detectors do not require guard segments to identify side-exiting particles. We have also segmented these detectors into a central bull's-eye surrounded by a wide annular region so that we can dynamically adjust the LET geometrical factor and thereby increase the dynamic range in particle intensities that EPI-Hi can measure.

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Figure 58: (a) Illustration of the detector segmentation. Detectors L0 through L3 are used in LET; detectors H1 through H3 are used in HET. The dashed line in the L0 detector drawing indicates the diameter inside which the silicon thickness is 12 µm. The L4, L5, and L6 detectors are identical to L3 while the H4 and H5 detectors are identical to H3. (b) Photograph of a prototype L1 detector (image credit: ISIS-EPI collaboration)

Species and Energy Coverage: Figure 59 shows the energy range covered by LET and HET for elements with atomic numbers in the range 1 ≤ Z ≤ 30. The primary measurements involve the analysis of particles that trigger two position-sensitive detectors and then stop in the central detector stack, as indicated by the absence of a signal in the detector at the opposite end of the stack. In the case of HET we extend the coverage to higher energies by also analyzing events due to particles that penetrate the entire thickness of the stack.

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Figure 59: Species and energy coverage by LET and HET telescopes (image credit: ISIS-EPI collaboration)

Viewing Directions and Angular Sectioning: Figure 46 shows the locations of the EPI-Hi 45º half-angle conical FOVs in the sky. Although these fields are largely unobstructed, there are several places where portions of the spacecraft or of other instruments block them. The most significant blockage is by the spacecraft's TPS. The forward view cones of HET and LET1 are intentionally oriented so that the TPS cuts through them, thus allowing detection of particles arriving from as close as possible to the radial direction over the full EPI-Hi energy range. The combination of the EPI-Hi telescopes will allow measurement of nearly complete pitch angle distributions for nominal Parker spiral magnetic fields over the entire orbit except for blockage by the TPS, which affects less than 10º even near perihelion. A number of the EPI-Lo apertures (Figure 47) are also oriented for viewing within the overlap region between the HET and LET1 forward fields of view, thus allowing excellent spectral coverage in the region where the average Parker spiral magnetic field will often be located when the spacecraft is near the Sun.

We assign detected particles to one of 25 sectors based on which of the 5 main detector segments they hit in two position-sensitive detectors (Figure 58). By orienting one of the two detectors with its quadrant sectors rotated 45° relative to those on the other detector we are able to identify 25 unique mean viewing directions, as illustrated in Figure 60a. The angular width of the individual sectors is broad, extending as much as 25° from the mean direction in some cases. However,≥ 80 % of the geometrical acceptance of the sectors fall within 15º of the mean. We rely on the significant overlaps among the sectors to provide information about particle angular distributions, as illustrated in Fig. 60b, where we show the accuracy with which the arrival direction of a narrow beam of particles can be determined as a function of that direction and of the number of particles counted.

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Figure 60: (a) Mean viewing directions of each of the 25 sectors. Dashed circles indicate angles of 15º, 30º, and 45º from the telescope axis. (b) Simulated distributions of derived incidence directions for parallel beams incident at 5º, 10º, 15º, 20º, 25º, and 30º from the telescope axis when determining the direction using 200 detected particles (image credit: ISIS-EPI collaboration)

Front-End Electronics: The telescope boards process the signals from the silicon detectors using an ASIC called PHASIC (Pulse Height Analysis System Integrated Circuit). A first generation version of this part is flying in the STEREO/LET and HET instruments. Solar Probe Plus requires a higher radiation tolerance than STEREO, so we have made several modifications to the PHASIC design to facilitate the production of a version that can tolerate a dose of at least 100 krad and also to improve several key aspects of its performance.

The PHASIC contains 16 complete dual-gain PHA chains. Figure 61 shows a schematic diagram of one of these chains. It consists of a charge-sensitive preamplifier that drives two independent post-amplifiers that shape the pulse and provide additional gain. By having two parallel analysis chains we are able to make precise pulse height measurements over the preamplifier's entire dynamic range. The shaped pulse, which is bipolar with a peaking time of 1.9 µs, drives a peak detector that contains a linear gate and discriminator. The discriminator controls the linear gate and prevents subsequent pulses from contaminating the detected peak level until it has been digitized. The PHASIC uses Wilkinson rundown ADCs, which determine the pulse height by measuring the time interval required for discharging a hold capacitor using a constant current. The PHASIC's on-chip digital circuitry also counts the number of triggers of the high- and low-gain analysis chains and provides several programmable-input OR-gates that are used to perform the lowest level of event coincidence logic and thereby enable rapid rejection of some uninteresting triggers without requiring readout of the digitized pulse heights. The PHASIC developed for EPI-Hi also incorporates shift registers to record the time-history of the discriminator outputs and aid in identification of crosstalk. Each dual-gain PHA includes a precision test pulser for use in the functional testing and calibration of the circuit.

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Figure 61: Schematic diagram of one PHASIC channel (image credit: ISIS-EPI collaboration)

The PHASIC design includes several parameters that are programmable by setting bits in a digital control register. This programmability provides sufficient flexibility so that the PHASIC can meet the analysis requirements of all of the different detectors used in EPI-Hi. By programming the feedback capacitance in the preamplifier we select the gain appropriate to the range of signals for a given thickness of detector. We can also select between two options for the gain ratio between the high- and low-gain analysis chains. Other programmable parameters include the discriminator thresholds and the value of the capacitor used by the test pulser for injecting a calibrated charge into the preamplifier.

Table 11 summarizes the PHASIC specifications. As illustrated in Figure 62, the ASIC is installed in a hybrid circuit containing a number of passive components that support the operation of the chip. These include a precision resistor for each PHA (Pulse Height Analysis) chain that sets the ADC's rundown current and provides excellent stability over temperature. The design of the ASIC itself relies on ratios of component values for setting key parameters, since such ratios are much more predictable and stable than absolute component values. We have fabricated and tested a prototype version of the EPI-Hi PHASIC using the commercial ON-Semi C5N CMOS process that will allow the addition of Aeroflex processing steps to improve the total tolerance to >100 krad.

Number of dual-gain PHAs

16

Chip size

7.4 mm x 7.4 mm

Power

11 mW per active PHA

Dynamic range

Up to 23000 (full scale/trigger threshold)

Integral non-linearity

<0.05% of full scale

Differential non-linearity

<1%

High/Low gain ratio

68 or 40, configurable

ADC type

Wilkinson

ADC resolution (each gain)

11 bits, 12th bit overflow

Shaping

Bipolar, 1.9 µs to peak

Preamp feedback capacitance

5–75 pF, programmable in 5 pF steps

Preamp full scale output swing

4.0 V

Cross-talk between PHAs

<0.2%

Radiation tolerance

>100 krad, no latchup below 80 MeV/(mg/cm2)

Gain temperature coefficient

<50 ppm/ºC

Offset temperature coefficient

<0.1 channel/ºC

Operating temperature range

-30 to +50ºC

Leakage current balancing

Up to 32 µA with 10-bit resolution

Threshold programmability

Up to 6% of full scale (each gain) with 10 bit resolution

Deadtime per trigger

4–67 µs (pulse height dependent)

Table 11: PHASIC chip specifications

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Figure 62: Photo of the PHASIC chip installed in hybrid (image credit: ISIS-EPI collaboration)

 



FIELDS (Electromagnetic Fields Investigation)

The FIELDS instrument is being developed at UCB/SSL (University of California, Berkeley/Space Sciences Laboratory), PI: Stuart Bale. The SPP/FIELDS instrument suite was developed under NASA contract NNN06AA01C. Development of the Search Coil Magnetometer sensor is funded by CNES (French Centre National d'Études Spatiales). FIELDS is utilizing four thin long whip antenna and three magnetometers placed on a common boom, the FIELDS instrument will measure: 41)

• Full DC and AC electric and magnetic fields Alfven waves and turbulence, Poynting flux, whistlers, electrostatic waves, solitary waves, shocks and upstream waves, reconnection jets, vperp and more

• Very accurate plasma density (< 1%) and electron temperature (< 5%), strahl diagnostics

• Fast (1 kHz) spacecraft potential and density fluctuation measurements - compressive waves, shocks, current sheets, probe-to-probe timing

• Radio emission (remote) from flares, (Parker) microflares, CME-driven shocks (shock speed/strength)

• Dust and nanodust diagnostics - occurrence rates and mass/speed diagnostics.

FIELDS has significant advantages over previous instruments:

• First-ever ‘double probe' electric field instrument designed and flown for solar wind studies. The instrument suite will provide:

- Low frequency/DC plasma wave electric fields and phase speeds

- Poynting flux

- Rapid density fluctuations

- Spacecraft floating potential (for SWEAP electrons)

• Front-mounted design avoids enormous plasma wake signatures

- Design optimized to measure true solar wind LF/DC fields

- Design minimizes heat input to the SPP spacecraft

- Longer antenna baseline = better sensitivity for HF/radio/QTN

• FIELDS measures dust impact rate and mass/charge

- More sensitive than traditional dust detectors, larger collection area

- Sensitive to ‘nanodust' picked up by solar wind

• Dual fluxgate magnetometers reduce spacecraft magnetics challenges.

 

Instrument design:

The FIELDS instrument combines magnetic and electric field measurements into a single, coordinated experiment. Magnetic fields are measured using both fluxgate and search-coil (induction) magnetometers mounted on a deployable boom in the spacecraft umbra. FIELDS will make electric field measurements both as a current-biased resistively-coupled double-probe instrument and as a capacitively-coupled radio and plasma wave instrument.

This places several constraints on the geometry and surfaces of the antenna system, and upon the design of the preamplifier and receiver electronics. In addition to the primary measurement objectives of electric and magnetic fields and waves, the FIELDS measurements will provide very accurate electron density and temperature measurements, density and velocity fluctuation measurements, and signatures of dust impacts on the SPP spacecraft.

Figure 63 shows the general layout of the FIELDS sensors on the spacecraft. The V1–V4 electric field probes are mounted at the base of the SPP Thermal Protection System (TPS) or heat shield and deploy out into full sunlight. At the SPP perihelion altitude of 9.8 RS these antennas will reach temperatures of more than 1300 °C. Another simple voltage sensor V5 is mounted on the magnetometer boom in the umbra of the spacecraft. Two fluxgate magnetometers (MAGi and MAGo) and a SCM (Search Coil Magnetometer) are also mounted on the boom.

In addition to izts sensor complement, FIELDS consists of three digital signal processing boards, a computer/processor unit, two boards to control antenna biasing, magnetometer formatting electronics, and two low noise power supply units. FIELDS was originally proposed with single string architecture for the processing computer and power supply. A failure mode analysis of measurement requirements performed in Phase B, showed that a single failure in FIELDS could lead to the loss of an unacceptable number of mission-level science requirements. As a result, the FIELDS instrument architecture was split into two halves such that no single failure results in a loss of all measurements.

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Figure 63: FIELDS uses 5 voltage and 3 magnetic sensors to measure electric and magnetic fields. The four V1–V4 sensors are deployed into full sunlight near the base of the SPP heat shield (TPS). The SCM (Search Coil Magnetometer) is mounted at the end of the instrument boom. Two fluxgate magnetometers (MAGi and MAGo) and a simple voltage sensor V5 are also mounted on the boom (image credit: FIELDS collaboration)

The FIELDS instrument is very sensitive to conducted and radiated noise sources from other instruments and spacecraft subsystems; at high frequencies (~MHz) the instrument is sensitive at the level of nV/√Hz. To accommodate this, a spacecraft level electromagnetic cleanliness (EMC) program has been established. A fundamental requirement of the SPP EMC (Electromagnetic Compatibility) program is that all DC–DC power converters be operated at fixed frequencies in 50 kHz intervals, beginning at 150 kHz (i.e. 150 kHz, 200 kHz, 250 kHz, etc.) and that these chopping frequencies be crystal-controlled. This picket-fence approach concentrates power-supply noise and harmonics into well known and narrow frequency bands providing ‘clean' regions of spectral density in which to make sensitive measurements.

Further, the FIELDS instrument team made the decision to synchronize its internal sampling clocks to multiples of 150,000 Hz with its master clock operating at 150,000 Hz x 256 = 38.4 MHz. To accommodate digital signal processing algorithms, which prefer power-of-two data blocks, FIELDS uses a rescaled timebase that is called a ‘New York second' (NY sec) and define 1 NYsec as a convenient power-of-two number of clock cycles (217) of the standard 150 kHz power supply chopping frequency. Thus 1 NYsec is defined as being 217/150,000 ~ 0.873813 . . . sec.

While the FIELDS DFB ‘burst mode' operates in sync at 150,000 Sa/s (samples/second), which is 217 samples/NYsec, the lower cadence data is sampled at rates that are 150,000 Sa/s divided by further powers-of-two. For example, the fluxgate magnetometers operate in synchronization at 150,000/29 ~292.969 Sa/s, which is exactly 256 Sa/NYsec. This allows FIELDS to sample in synchronization with the EMC prescribed frequency of 150 kHz avoiding large noise signals from power converters and to maintain the power-of-two data format desired by FFT (Fast Fourier Transform) algorithms. Note that the FIELDS and SWEAP instruments use a master/slave clock configuration over a dedicated interface: SWEAP uses the FIELDS clock signal when available. This will maintain phase coherence between the FIELDS and SWEAP measurements, enabling both accurate high cadence data processing and the removal of deterministic noise signals.

Sensor and Preamp Design: The FIELDS suite uses five voltage probes and three magnetometers to make measurements over 20 MHz of bandwidth and 140 dB of dynamic range. The ‘V1–V4' voltage probes function both as current-biased double probe electric fields sensors (as on the THEMIS or Van Allen Probes satellites) and as capacitively-coupled radio and plasma wave sensors (as on Wind, Cassini, STEREO). The V5 sensor makes a simple voltage measurement near the magnetometer boom and may be used to infer the sunward electric field for plasma waves. The fluxgate and search coil magnetometers are standard devices for measuring low frequency and wave magnetic fields.

The V1–V4 Electric Antennas: Four voltage sensors (V1–V4) are deployed in nearly orthogonal, co-linear pairs slightly behind the plane of the spacecraft heat shield (the Thermal Protection System or ‘TPS'), as shown in Figure 63. To function properly as a double-probe electric field measurement, these sensors must be coupled to the plasma through a photoelectron current; i.e. they must be in sunlight. This orientation also keeps the V1–V4 antennas out of the spacecraft wake and ensures a minimal perturbation to the spacecraft interaction with the solar wind.

Figure 64 shows a CAD drawing of one of the units; each unit is identical, with some small differences in the mounting hardware. The primary sensor consists of a 2 m long, 1/8" diameter Niobium C-103 thin-walled tube (called the ‘whip'). All but the last 8 cm of the whip are exposed to full sunlight and will reach high temperatures (>1300ºC) at SPP perihelion. The whip is clamped to a 30 cm Molybdenum ‘stub' element which acts as an electrical and thermal isolator and the signal from the whip is fed through with a molybdenum wire to the hinge and preamp below. A chevron shaped C-103 heat shield covers the final 8 cm of the whip and the entire stub. This creates a substantial shadowed area that radiates excess heat into space. Additionally, the shield, whip, and stub are all isolated from each other with sapphire, a good thermal insulator. Both of these features greatly reduce the thermal input into the base, where conductors and other materials must be below 230ºC. This design has been modeled and verified in the laboratory. The preamp sits near the bottom of the antenna mechanism, to minimize stray capacitance from the cables. The antenna system is stowed back against the spacecraft for launch and deploys by spring force and is rate-limited by a flyweight brake system. After release, it will take a few seconds to deploy to the final configuration. Each sensor will be deployed individually, with the FIELDS electronics operating, to aide in calibration and characterization of the measurement.

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Figure 64: A CAD drawing of a V1–V4 antenna unit. The whip (colored green) is the sensor and extends 2 meters beyond the end of the 30 cm stub. The stub acts as an electrical and thermal isolator. The niobium C-103 whip signal is carried back through a small, pure niobium wire contained in the stub to the preamplifier at the base. A heat shield shadows the stub, allowing it to radiate excess heat from the whip, while another shield supports blanketing that blocks heat radiating from the TPS (image credit: FIELDS collaboration)

The V5 Voltage Sensor: A simple voltage probe (V5) will be mounted on the SPP magnetometer boom, deployed in the umbra behind the spacecraft (and is therefore coupled to the plasma through thermal electrons rather than photoelectrons). While this sensor sits in the spacecraft plasma wake and will see the low frequency signatures of that interaction, it will provide good capacitively-coupled voltage measurements of the radial electric field E|| present in plasma waves and will help constrain the knowledge of the electrostatic center of the spacecraft.

The V5 design is shown in Figure 65. Two short sensor elements extend from a preamplifier box and are electrically isolated from the box by PEEK fittings. The two tube elements are tied together at the preamp (i.e. it is not a differential measurement).

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Figure 65: A CAD drawing of the V5 voltage sensor. Two short tubes act as a single (electrically tied) sensor to measure the plasma voltage. A simple preamplifier is housed in an attached enclosure. The tubes can be current-biased and the preamp enclosure can be voltage-biased (image credit: FIELDS collaboration)

Electric Preamplifiers: The four main antennas are connected individually to the V1–4 preamps and the fifth antenna on the mag boom is connected to the V5 preamp. The preamps provide low-noise highimpedance inputs, voltage gain, and low impedance outputs. As shown in Figure 66, the V1–4 preamps provide three outputs: the HF 20 MHz bandwidth output to the RFS (Radio Frequency Spectrometer), the MF 1 MHZ output to the TDS (Time Domain Sampler), and the LF 64 kHz output to the AEB (Antenna Electronics Board) and DFB (Digital Fields Board). The HF amplifier chain is a new design consisting of a FET input buffer followed by a wide bandwidth op amp providing gain and driving a 50 ohm terminated coax output. The LF and MF signals are provided by a second unity gain op amp, as used in legacy designs (e.g. THEMIS, RBSP). This op amp is powered by a supply referenced to a "floating ground driver" on the AEB and provides a signal range of ±70 V from DC to 300 Hz and ±10 V from 300 Hz to 1 MHz. The V5 preamp does not include the HF chain and consists of the single legacy op amp design providing LF and MF outputs.

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Figure 66: A simplified schematic of the V1–V4 electric preamplifier circuit. A signal from the antenna whip is fed into 3 separate channels that feed the DFB, TDS, and RFS receivers. The LF side using a floating voltage system to accommodate the large expected plasma voltage variations. The grey box in the upper left represents the plasma voltage signal and sheath impedance, and some estimated values of the sheath resistance are shown in the table within the figure (image credit: FIELDS collaboration)

Fluxgate Magnetometers: The two fluxgate magnetometers (MAGs) for SPP are similar to the triaxial, wide-range, low-power, and low-noise magnetometers built by GSFC (Goddard Space Flight Center) for MAVEN, Van Allen Probes, STEREO, etc. This line of flight magnetometers totals 79 instruments to date starting with IMP-4, launched in 1966. For SPP, the MAGs will provide data with bandwidth of ~140 Hz, sampling at 292.97 Sa/s, as part of the FIELDS instrument suite. The measurement dynamic range is ±65,536 nT with a resolution of 16 bits. The primary science objectives addressed by the magnetometers are determining the structure and dynamics of the magnetic fields at the sources of the fast and slow solar wind, contributing to the study of the coronal processes that lead to heating of the solar corona, and exploring the roles of shocks, reconnection, and turbulence in accelerating energetic particles.

The sensor design for the SPP MAGs (Figure 67) will provide maximum thermal isolation from the boom, which will undergo considerable temperature variation, depending on its location in the umbra or in sunlight. Kinematic mounts limit the heat transfer across the feet of the sensor. The heritage of the kinematic mounts derives from the Juno magnetometers, where they were used to ensure that the sensor temperature variation was limited. Heater power is provided by a proportional AC heater to reduce the temperature variations of the sensors and to provide survival heating. The heater is synchronized to a frequency provided by theMEP. The sensor mass requirements have lead to use of a lightweight composite base.

Each MAG sensor has a corresponding electronics board in the SPP FIELDS MEP. For enhanced reliability, the outer MAG board is controlled by the DCB, and the inner MAG board by the TDS. The control component on each board is the rad-hard Aeroflex FPGA that contains all of the MAG logic functions and on-board SRAM. These functions include command handling, telemetry packet formatting, ADC readout, auto-ranging algorithm, DRIVE clock generation, and housekeeping readout. MAG produces a data product at the required cadence of 1 message per 0.874 seconds. The ranging algorithm selects one of the four ranges (±1024, ±4096, ±16,384, ±65,536 nT) based on the ambient magnetic field. The sensor connection to the electronics boards is accomplished by a tuned circuit, including the harness, which is calibrated at the GSFC Acuña Magnetometer Test Site. Because the sensor temperatures will be substantially lower than standard GSFC magnetometer sensors and will vary significantly due to changes in the spacecraft Mag boom orientation relative to the Sun, calibration will be performed over a wide range of temperatures.

PSP_Auto4

Figure 67: CAD drawing of a SPP MAG sensor, showing the composite structure supporting the two bobbins and electronics board (green) inside the composite cover. Also seen are two of the three kinetic mounts supporting the sensor on the composite 4-hole mounting plate, the alignment cube, and the pigtail harness that connects to the spacecraft harness (image credit: FIELDS collaboration)

TheMAG sensors are mounted on the SPP magnetometer boom (Figure 68), where their relative separation and close proximity to the spacecraft compromise their gradiometric functionality, and thus the ability for an accurate removal of any spacecraft field contamination at the outer MAG. This enhances the importance of the magnetic cleanliness testing being implemented for the spacecraft and payload.

PSP_Auto3

Figure 68: A schematic of the spacecraft magnetometer boom and sensors, shown deployed. Two fluxgate magnetometers are located at 1.9 m (MAGi) and 2.72 m (MAGo) from the rear deck of the spacecraft. The V5 voltage sensor is at 3.08 m and theSCM is located at the end of the boom: 3.5 m from the spacecraft. This is a relatively short boom, constrained to remain in the spacecraft umbra at perihelion. SCM data will require special processing to remove the drive signal from the fluxgates (image credit: FIELDS collaboration)

SCM (Search Coil Magnetometer): The SCM will measure all three components of the AC magnetic signature of solar wind fluctuations, from 10 Hz up to 50 kHz and a single component from 1 kHz to 1 MHz. The wide bandwidth and dynamic range allows FIELDS to investigate transients caused by interplanetary shocks and reconnection, the turbulent cascade beyond the electron kinetic scale, but also numerous plasma wave modes.

The SCM instrument consists of a triaxial search-coil that has a solid technical heritage in several past missions. Nearly identical instruments are being built for the Taranis and Solar Orbiter missions. Each sensor consists of a magnetic core with a winding whose voltage is proportional to the time-derivative of the magnetic field. Two sensors of SCM cover the ELF/VLF frequency range from 10 Hz–50 kHz. The third one is a dual-band sensor that covers both the ELF/VLF and the LF/MF (1 kHz–1 MHz) ranges. The three sensors, each of which is 104 mm long, are mounted orthogonally on a non-magnetic support (Figure 69).

Two challenges raised by Solar Probe Plus are the low temperature environment on the magnetometer boom, and the unusually large dynamic range of the instrument. The latter is needed to accommodate both small-amplitude fluctuations of solar wind turbulence, and large transients near the Sun. Peak values, as scaled from observations made by Helios at distances from the Sun down to 0.29 AU, may reach 3000 nT in the ELF/VLF range. Thanks to a careful design, the dynamic range of the instrument has been increased from past models by several tens of dB, now reaching 160 dB in the ELF/VLF range, and 130 dB in the LF/MF range. SCM will be located in the shade of the spacecraft, at the end of the magnetometer boom, and thus needs a heater to keep it above deep space temperatures. To mitigate thermal losses, the instrument will be wrapped in an insulatingMLI layer, with very compact design. The SCM design is very compact; in particular, the preamplifier has been miniaturized by 3D Plus, and will be located inside the foot of the instrument.

PSP_Auto2

Figure 69: SCM engineering model for SPP (image credit: FIELDS collaboration)

The sensitivity and instrument response of the SCM are illustrated in Figure 70. The sensitivity is sufficient to observe small-amplitude solar wind turbulence in the inner heliosphere, and properly distinguish Elsässer variables, while also capturing large transients (hence the low gain -50 dBV/nT in the ELF/VLF range). The analog signals in the ELF/VLF range will be processed by the Digital Fields Board (DFB), which will deliver either spectra or continuous waveforms up to 150,000 Sa/s. The LF/MF signal will be processed by any of the RFS, DFB, or TDS receivers. The survey data products will be spectral matrices, which give access to the polarization, and waveforms of up to 293 Sa/s for all three components. The latter will be merged with the DC magnetic field, as measured by MAG, into one single composite magnetic field product.

PSP_Auto1

Figure 70: Measured sensitivity (in red) and frequency response (in blue) of SCM. The curves on the left are for the ELV/VLF antenna and the curves on the right for the LF/MF one. The highest measurable levels are 3000 nT in the ELF/VLF range, and 100 nT in the LF/MF range (image credit: FIELDS collaboration)

MEP (Main Electronics Package): The MEP is the stack of FIELDS receivers, computers, and power supplies and is mounted within the SPP spacecraft structure. Figure 71 shows the engineering model (EM) MEP on the bench in Berkeley.

PSP_Auto0

Figure 71: A photograph of theEM (Engineering Model) of the FIELDS main electronics package and V1–V4 preamps. The individual boards are labeled (image credit: FIELDS collaboration)

 

HeliOSSP (Heliospheric Origins with Solar Probe Plus)

Marco Velli of JPL (Jet Propulsion Laboratory) will serve as the Observatory Scientist (OS) in support of the NASA Science Mission Directorate. The OS investigation addresses SPP science objectives using the SPP system of measurements and provides theoretical input and independent advice to maximize the scientific return from the mission. 42)


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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 (herb.kramer@gmx.net).

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