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."

NASA missions are most often renamed after launch and certification; in this case, given Parker's accomplishments within the field, and how closely aligned this mission is with his research, the decision was made to honor him prior to launch, in order to draw attention to his important contributions to heliophysics and space science.

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.


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).


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.


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


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).


Solar Probe (2005)

Solar Probe Plus (2008)

Minimum perihelion

4 Rs

9.5 Rs


90º from ecliptic

3.4º from ecliptic

Number of solar passes



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


5.5 AU

1 AU

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


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


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).


TRL (Technology Readiness Level)


TPS (Thermal Protection System)


NASA sponsored technology development

Solar array


Combines space heritage and concentrator cells

Cooling system


Adapted from heritage systems

X-/Ka-band transponder


NASA sponsored technology development

LEON3 processor


Qualified product in new JHU/APL application

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




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.


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


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


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.


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.


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.


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)


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



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


Status of the project:

• On 6 November 2017, NASA's Parker Solar Probe spacecraft arrived at NASA/GSFC (Goddard Space Flight Center) in Greenbelt, Maryland, for environmental tests. During the spacecraft's stay at Goddard, engineers and technicians will simulate extreme temperatures and other physical stresses that the spacecraft will be subjected to during its historic mission to the Sun. 25)

- Before arriving at Goddard, Parker Solar Probe was at the JHU/APL (Johns Hopkins University /Applied Physics Laboratory) in Laurel, Maryland, where it was designed and built.


Figure 15: Parker Solar Probe arrives at the integration and testing facility at NASA/GSFC in Greenbelt, Maryland (image credit: NASA/JHU/APL, Ed Whitman)

• 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. 26)


Figure 16: 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. 27)

• 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. 28)



Launch: A launch of the Solar Probe Plus spacecraft is scheduled for July 31, 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. 29)

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).


Figure 17: 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 18 highlights the primary characteristics of each period (Ref.22) .


Figure 18: SPP Orbital Operations Concept (image credit: JHU/APL)