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FBO DAILY - FEDBIZOPPS ISSUE OF AUGUST 08, 2018 FBO #6102
SOURCES SOUGHT

A -- NASA-GSFC POD FOR THE HELIOPHYSICS TECHNOLOGY DEMONSTRATION MISSION OF OPPORTUNITY CONCEPT SPACECRAFT AND ON-ORBIT SUPPORT - PDF VERSION OF SYNOPSIS

Notice Date

8/6/2018
 

Notice Type

Sources Sought
 

NAICS

336419 — Other Guided Missile and Space Vehicle Parts and Auxiliary Equipment Manufacturing
 

Contracting Office

NASA/Goddard Space Flight Center, Code 210.S, Greenbelt, Maryland, 20771, United States
 

ZIP Code

20771
 

Solicitation Number

NASA-GSFC-HELIOPHYSICS-TECHNOLOGY-DEMONSTRATION-MISSION-POD
 

Archive Date

9/4/2018
 

Point of Contact

Laura Ottenstein, Phone: (301) 286-4141, Rosa E. Acevedo, Phone: (301) 286-7152
 

E-Mail Address

Laura.Ottenstein-1@nasa.gov, rosa.e.acevedo@nasa.gov
(Laura.Ottenstein-1@nasa.gov, rosa.e.acevedo@nasa.gov)
 

Small Business Set-Aside

N/A
 

Description

PDF VERSION OF SYNOPSIS Partnership Opportunity Document (POD) for NASA's Goddard Space Flight Center (GSFC) Heliophysics Technology Demonstration Mission of Opportunity Concept Spacecraft and On-Orbit Support dated August 6, 2018 1.0INTRODUCTION/SCOPE NASA recently released a draft Announcement of Opportunity (AO), Heliophysics Technology Demonstration SALMON-3 Program Element Appendix (PEA), NNH18ZDA009J (hereafter referred to as the "Tech Demo AO"). The final Tech Demo AO is expected to be released in August 2018. NASA's Goddard Space Flight Center (GSFC) is developing a mission concept to be proposed in response to this Tech Demo AO. The purpose of issuing this Partnership Opportunity Document (POD) is to select a teaming partner to support the proposal effort. The mission comprises two Spacecraft (SC) called the Optical SC (OSC) and the Detector SC (DSC). Each SC consists of a GSFC-provided Instrument Payload (PL) and a vendor-provided SC bus that supports the PL. The mission concept development effort is currently in pre-Phase A. This phase ends with the submission of a Step 1 proposal, which is due 3 months after the AO is released. If the proposal is down-selected, the next step in the proposal process will be a 9-month mission concept study culminating in a Concept Study Report, which is the Step 2 proposal, and a Site Visit. The following schedule should be used as a basis for responses to this opportunity: Partnership Opportunity Document releasedAugust 6, 2018 Responses dueAugust 20, 2018 Partner Selection announcedAugust 23, 2018 Step 1 Proposal in response to SALMON-3 AO AO release date + 3 months Phase A Mission Concept Study April TBD, 2019 Mission Concept Study Report dueJanuary TBD, 2020 Down-Selection announcedJune TBD, 2020 Begin Phase B June TBD, 2020 Launch Readiness No later than December, 2024 1.1COST Total cost and cost credibility are important issues for the proposed mission, and trade studies are expected to be required in this regard. The cost cap for this AO is $65 million in Fiscal Year (FY) 2019 dollars. This cost includes the instruments, spacecraft, integration and test, and mission and science operations, as well as a required 30% minimum unencumbered cost reserve on Phases A-D and a 20% minimum unencumbered cost reserve on Phases E-F. Reserves will be held at the Project level, not with the partner. Each Phase A mission concept study has a total cost cap of $0.4 million. There will be no exchange of funds between the teaming partners for the portion of this partnership opportunity dealing with the preparation of the initial submission of the Step 1 Proposal to the Tech Demo AO. Funding will be available for subsequent phases should the candidate mission concept be competitively selected for those additional phases. 1.2DESIRED MISSION SERVICES GSFC is interested in formally selecting a partner to provide the following products and services for the mission: •The OSC and DSC spacecraft buses •Support of the integration of the GSFC PLs to the OSC and DSC buses •Integration of the resulting OSC and DSC spacecraft to the launch vehicle •On-orbit mission operations for the duration of the mission (a minimum of 6 months and nominally 12 months). All interested parties are required to respond to this POD in accordance with Section 5 below. 1.3PROPOSAL SUPPORT It is expected that the selected POD respondent will provide support using their own resources to help develop the proposal in the areas of well-defined and documented spacecraft, instrument accommodations, instrument-to-spacecraft and spacecraft-to-launch vehicle integration support, and mission operations. Partner responsibilities include collaborating with the Principal Investigator (PI) and other proposal team members to develop: end-to-end performance requirements, system architecture, well-defined interfaces to the spacecraft in the form of Interface Control Documents (ICDs), and predicted flight performance. Partner responsibilities also include participation in trade studies and cost estimation for mission phases B-F. The period of performance for this phase of proposal support is expected to last approximately 3 months, starting in August 2018. If the proposal is down-selected for a 9-month Mission Concept Study (Phase A), partial funding will be provided to the partner to support the Study, the writing of the Mission Concept Study Report, and the Site Visit. If the mission is selected for development, launch, and operations (Phases B-F), the partner will be responsible for the design and fabrication of the OSC and DSC buses, instrument accommodations and spacecraft integration support, integration support for integrating the spacecraft with the launch vehicle, and mission operations. The period of performance for this interval is expected to last approximately 5 years, starting mid-2020. These dates and times may change depending on selection timelines and budget allocations or phasing. The respondent to this POD shall state the timeframe in which their services will be available and the extent to which the respondent can meet the timeline requirements set forth above. 2.0MISSION PROGRAMMATICS OVERVIEW 2.1GENERAL The Tech Demo PEA solicits Small Complete Mission (SCM) proposals for spaceflight demonstration of innovative medium Technical Readiness Level (mid-TRL) technologies that enable significant advances in NASA's Heliophysics Scientific Objectives and Goals. Spacecraft bus technologies proposed may have a TRL of less than 6 when proposed, but any such technologies must be accompanied by a roadmap for advancing them to TRL 6 by Preliminary Design Review (PDR) as specified in Requirement 35 of the SALMON-3 AO. The NASA Solar Terrestrial Probes (STP) Program Office will manage the Heliophysics Tech Demo Mission of Opportunity (MO) investigations under the requirements of NPR 7120.5E, NASA Space Flight Program and Project Management Requirements for Class D missions, as described in Section 4.1.2 of the SALMON-3 AO and as modified by the NASA SMD Tailoring/Streamlining Decision Memorandum (issued Dec. 7, 2017). These and other documents related to Class D missions are available at https://soma.larc.nasa.gov/stp/tdmo/tdmo-library.html. The mission will be managed by GSFC in partnership with the Partner selected through this POD. 2.2LAUNCH VEHICLE Access to space for the solicited Tech Demo investigation(s) will be provided by NASA in the form of a secondary payload accommodation on the Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA) planned for NASA's Heliophysics STP-5 mission, Interstellar Mapping and Acceleration Probe (IMAP), projected for launch in 2024. Integration costs to the IMAP ESPA will be funded by NASA. Except for a battery trickle charge until the moment of launch and the signal to release the secondary payload, the ESPA will not provide propulsion, power or other spacecraft support beyond the standard ESPA accommodations. The launch vehicle and launch services will be provided by NASA as part of the IMAP mission. The selected partner must demonstrate that the proposed spacecraft are compatible with the ESPA Grande system interface specifications for the Tech Demo AO, available at https://soma.larc.nasa.gov/stp/tdmo/pdf_files/2018-07-10_IMAP_ESPA_SIS.pdf. 2.3LAUNCH MANIFEST The partner must commit to providing the solicited services on a schedule commensurate with the requirements of the Tech Demo AO and the IMAP mission. The mission is required to be launch ready no later than June 30, 2024, or 6 years after the contract is in place, whichever is earliest. Earlier launch readiness dates may be considered, pending the AO final guidance. 3.0MISSION DESCRIPTION The following mission description is conceptually accurate but not final in its specifics. In particular, the description of the SC buses and their subsystems provides in various places the specific capabilities of certain components. These notional characteristics are the result of a concentrated but limited GSFC study that produced an initial point design that was not optimized for Size, Weight, Power and Cost (SWaP-C). It is important to stress that the specifics of this point design, including illustrations, should not be interpreted as placing any design requirements on the proposed OSC and DSC buses; they are included only to help explain what the buses need to accomplish to support the payload and mission. Respondents to this POD are encouraged to propose more efficient OSC and DSC designs. After a partner is selected, GSFC expects to work with the partner to refine the mission design further to minimize SWaP-C and take full advantage of heritage from existing partner SC buses. Capability descopes that preserve the threshold mission objectives (which will be known to the partner) will be jointly examined. POD respondents can suggest potentially high-value targets for future trade studies associated with bus design. The overview is organized in sections describing: •The GSFC-supplied payloads •The SC buses that will accommodate these payloads •Mission operations An organization responding to this POD needs to describe: •The proposed design of its OSC and DSC buses •How the buses will accommodate and interact with the GSFC-supplied instrument payloads described below •How the buses will be operated and controlled during the mission 3.1MISSION OVERVIEW The mission comprises two 3-axis stabilized spacecraft (SC), the Optical SC (OSC) and the Detector SC (DSC), each having approximately 100 kg of total mass in the initial point design. The OSC and DSC will together carry out a formation flying mission in an orbit around the Sun-Earth L1 Lagrange Point (SEL1) point. Figure 1 (figures are collected in Section 9 of this document) shows the OSC and DSC in their stowed configurations when they are attached to each other and ESPA ring with two Releasable Clamp Band (RCB) assemblies. The DSC and OSC are shown in their separated and deployed configurations in Figure 2. The mission will be a secondary payload (PL) on the NASA IMAP Mission and use one of the 24-inch diameter ports on the ESPA Grande Ring that will be on the IMAP LV. Within about an hour after launch, and minutes after the IMAP SC has separated from the LV, the two spacecraft will be released from the ESPA as a single connected unit into a transfer trajectory to an orbit about the SEL1. The OSC and DSC will separate from each other when the Lagrange Orbit Insertion (LOI) maneuver has been completed, approximately 45 days after launch. Figure 3 illustrates their approximate ephemeris they will follow and where the separations occur. Figure 4 summarizes the operations over the course of the mission, which will last 6 to 12 months. The single ephemeris in Figure 3 covers both the minimum 6-month on-orbit operating time and the nominal total mission lifetime of 12 months. Because the SC will not approach the vicinity of Earth after completing their operations in SEL1 orbit, there is no requirement for end-of-mission decommissioning, although spacecraft safing measures such as propellant evacuation are expected be required. Both spacecraft require propulsion to attain and maintain formation. Figure 5 summarizes the layout of the thrusters on the DSC and OSC. In formation, the OSC and DSC act as a distributed Photometric Solar Telescope (PST), with the OSC about 150 m nearer the Sun than the DSC. As shown in Figure 2, the OSC carries the PST optical devices that focus light from the Sun onto the PST detectors on the DSC. To maintain the formation required by the PST, both the OSC and the DSC will also carry a set of emitters and detectors, called the Precision Formation Flying System (PFFS) to sense the alignment and separation between the OSC and DSC. 3.2INSTRUMENT PAYLOAD GSFC will provide two technology demonstration instruments, the Photometric Solar Telescope (PST) and the Precision Formation Flying System (PFFS). The PST and the PFFS constitute the payload; other elements of the OSC and the DSC are designated collectively as the spacecraft bus, for which the Partner is responsible. Both the PST and the PFFS have components on both spacecraft, as illustrated in Figure 6. The total OSC payload mass is about 7.5 kg CBE and 10 kg MEV. The OSC payload requires an estimated total nominal power of 26 W CBE/34 W MEV, with an estimated total peak power of 30 W CBE/40 W MEV. The total DSC payload mass is about 24 kg CBE and 31 kg MEV. The DSC payload requires a total nominal power of 115 W CBE/150 W MEV with an estimated total peak power of 120 W CBE/155 W MEV. Taken together, the PST and PSS elements will produce an average of about 1 kbps of housekeeping telemetry. This will result in less than 1 Gb of housekeeping data per week that will have to be stored and then downlinked weekly. 3.2.1PST Using the formation enabled by the PFFS, the PST will acquire images of the Sun. The PST data, which account for about 90% of the total data volume, are acquired only during fine alignment mode of the two SC. PST operation is expected to be limited by collected data downlink capacity to a duty cycle of less than 10%. In other words, the mission will acquire PST image data in bursts, and only a ground-commanded or payload-selected subset of the data will be downlinked. The result is that PST is expected to produce on the order of 80 Gb per week of solar observations that need to be downlinked weekly. The OSC carries the focusing elements of the PST, each of which acts like a lens with a focal length of 150 m (the working separation of the precision formation). The PST focusing elements, which need to be as close as possible to the c.g. of the DSC, are housed in open-ended cylindrical enclosure with a diameter of about 0.5 m that passes all the way through the OSC and is located inside the RCBs on the OSC. As also shown in Figure 2, the DSC carries the uncooled PST detectors in a single vacuum enclosure. For the PST to operate as intended, the OSC and DSC need to be 150 meters apart, with an allowable error of ±5 mm. Also, both the OSC and DSC need to be pointed at a defined area of the Sun with an accuracy of about 20 arcsec. Roll about the boresight of the PST needs to be controlled to about 20 arcsec. The OSC and DSC also require a transverse alignment error of no more than about ±3 mm. 3.2.2PFFS The PFFS comprises a suite of hardware and software components required for the acquisition and alignment of two free-flying spacecraft, as shown in Figure 6. Specifically, a set of sensors measures the relative position (range and transverse relative position) between the OSC and the DSC, and a fine guidance sensor measures the inertial direction of the spacecraft orientation. These sensor data are fed into the PL computers where the PFFS navigation software produces a best estimate of the SC position relative to what is required to observe the science target on the Sun. The PFFS control software then generates, and sends on to the spacecraft bus computer, the thruster commands to maintain the OSC and DSC in a relative position that allows the science observations to be carried out. The spacecraft bus computer will pass the required thruster commands to the thrusters. Note that the spacecraft bus is responsible for the attitude control of both spacecraft. To support these modes, the PFFS navigation software also relies on sensor data that is produced by the spacecraft bus, such as coarse Sun sensor measurements, star tracker measurements, and gyro measurements. In addition to the control data that flow from the payload to the bus, some data will flow from the OSC payload to the DSC payload and vice-versa. Thus, the SC buses will provide a radio-frequency (RF) crosslink system, including antennas, that provides intersatellite communications, as well as a relative navigation RF system that determines the coarse range and bearing between the OSC and DSC. 3.2.3Payload Computer Both the OSC and DSC carry a payload computer that will communicate when necessary through the SC computer and radio crosslink provided by the SC buses. The OSC and DSC payload computers will support the PFFS. The DSC payload computer will also accept data from the PST detectors and store them in a solid-state buffer. 3.3SPACECRAFT BUS 3.3.1SC Bus Mechanical The proposed bus (comprising two connected spacecraft and their instrument payloads when mounted to the ESPA) must fit within a volume 42 x 46 x 38 inches, as shown in Figure 1. The ESPA Ring coordinate system and further information are available at https://soma.larc.nasa.gov/stp/tdmo/pdf_files/2018-07-10_IMAP_ESPA_SIS.pdf. As shown in Figure 1, the OSC and DSC buses include two Releasable Clamp Band (RCB) assemblies, one for separating the OSC from the ESPA Ring, and another that includes a connector to provide power and data feedthroughs, for separating the OSC from the DSC. It is expected that the LV will provide the means to actuate the OSC-ESPA RCB at the correct time, but the SC will have to activate the RCB to separate from each other. As shown in Figure 2, the OSC SC bus needs to deploy a 2.3-m diameter Sun shade to prevent the desired emission within the very narrow PST field of view from being overwhelmed by out-of-field light. The material that forms the Sun shade (which will presumably deploy with the OSC solar panels to block the open space between them) can be chosen freely. The SC buses also need to mechanically support the OSC and DSC payload masses given above. 3.3.2SC Bus Thermal The SC buses shall maintain the PST and PFFS payload components within their nominal temperature ranges, typically -10 to +40 °C operational and -35 to +70 °C survival. There are no cryogenic or unusually temperature-sensitive PL elements. 3.3.3SC Bus Propulsion The DSC bus shall incorporate thrusters to: •Be used early in the mission to maneuver the mated OSC-DSC pair into the desired trajectory toward an orbit about the SEL1 point •Face toward the OSC bus and control the relative separation of the two buses along the inter-spacecraft line. •Unload reaction wheel momentum •Maintain the SEL1 orbit The OSC bus shall incorporate thrusters to: •Unload reaction wheel momentum •Control the relative translational position of the two buses perpendicular to the inter-spacecraft line •Maintain the SEL1 orbit Figure 5 shows examples of the thruster units selected during the GSFC study. The Ibit values shown in Figure 5 reflect the estimated thrust control resolution that is required to maintain the fine formation with sufficient precision. The selected thrusters used non-contaminating cold gas propellant. The partner is not bound by this choice but must ensure that all propellants are compliant with ESPA requirements. The payload requirement for on station propellant will be less than 5 kg for the OSC and 2 kg for the DSC, with an assumed Specific Impulse (ISP) of 40 sec, based on the point design for the SC. 3.3.4SC Bus Attitude Control Each spacecraft shall provide pointing with control error not to exceed (NTE) 20 arcsec (3σ). Control of roll about the Sun-spacecraft line may be relaxed by up to a factor of 2 relative to control of pitch and yaw. Weekly command uploads will tell the OSC where to point on the Sun and for how long. Note that the pointing will not be static in the sidereal reference frame because the OSC and DSC, which rotate about the Sun once per year, will have to yaw by about a degree per day to stay aimed at the same spot on the Sun. Moreover, the OSC and DSC will be moving in their SEL1 orbit and following solar features as the Sun slowly rotates, so the command uploads will have to include specifications for these finer motions. In addition to supporting the PFFS modes involving attitude control, the DSC ACS is required to point the DSC so that the high-gain antenna (HGA) is aimed at the Earth during periods of ground communications. The ACS on both SC will also be used to aim the SC for momentum unloading and orbit maintenance burns, as well as for orientations needed by the thrusters to acquire or re-acquire the precision formation. 3.3.5SC Bus Power The OSC and DSC spacecraft shall have a power generating capability during all mission phases sufficient to power themselves and the PST and PFFS payloads. As shown in Figure 2, the OSC SA may support the sunshade (although this is not required), and the DSC SAs need to extend out far enough to get sufficient sunlight outside the shadow of the sunshade. The SAs selected during the GSFC study provided about 250 W End-of-Life (EOL) for the OSC and about 200 W EOL for the DSC. Each of the two SC have an ~ 420Wh battery The study-selected Electrical Power Supply (EPS) electronics were based on GSFC-developed units, but it is anticipated that the POD respondents will be able to specify EPS electronics based on the above SA and battery information. 3.3.6SC Bus Communications Figure 7 illustrates the overall mission communications architecture. The design study assumed that there will be two radio transceivers on each of the SC. One of these radios on each SC will support high-rate HGA and low-rate patch antennas communications with the ground; in the GSFC study, this was an X-Band radio with 7 W RF power. The radio shall be compatible with the NASA Deep Space Network (DSN) and shall be capable, in conjunction with the HGA, or providing an 8 Mbps downlink rate from the SEL1 orbit. This will allow about 80 Gb of data to be downlinked during one weekly DSN pass. The mission will also use the HGA to receive ground commands, which it will relay to the OSC and DSC payload computers as appropriate. Omni antennas shall be incorporated to support a 32 kbps downlink rate to the DSN with the same radio. The OSC and DSC buses shall also incorporate patch antennas and associated radios to provide cross-link communications between the DSC and OSC at 2.5 Mbps for separation distances up to about 50 km. The same radios will support relative navigation transmissions between the two SC. Relative navigation will enable the DSC and OSC to determine their approximate range and bearing when beyond the range of the medium alignment mode of the PFFS. 3.3.7SC Bus C&DH The DSC bus will provide a Command and Data Handling (C&DH) computer that will control all ground and inter-SC communications. It will accept subsets of the data buffered by the DSC payload computer and prepare these subsets for downlink to the DSN network, The C&DH computers shall provide data interfaces to the DSC payload computers capable of both sending and receiving data. The preferred data interface between the spacecraft and the instrument is not defined. Interfaces that can be supported by the spacecraft shall be described in the response to this POD. This data handling interface must be able to receive data from the DSC payload computer at a rate no lower than 40 Mb/s. The bus C&DH computer must be able to store at least 20 GB (160 Gb) of data for downlink. The spacecraft C&DH shall also supply a clock pulse with a maximum 1 sec cadence and an accuracy of 0.01 sec to the payload computers. The C&DH hardware used during the GSFC study of the mission was based on in-house designs, but it is thought that there are commercially available units that can provided the needed functionality. 3.4RADIATION TOLERANCE The mission components are expected to absorb a Natural Total Ionizing Dose (TID) of approximately 10 krad assuming 2.5 mm of aluminum shielding. Note that 2.5 mm is a reference value and not binding on a POD respondent. The respondent should document the radiation tolerance of key bus components (including a Radiation Design Margin of 2) as well as, in the case of electronic components, their susceptibility to Single Event Effects (SEEs), including latch-up immunity. 3.5MISSION OPERATIONS The Partner shall be responsible for establishing and operating a Mission Operations Center (MOC) for a minimum of 6 and a maximum of 18 months (12 months nominal) after launch. The MOC shall coordinate with the Science Operations Center (SOC), operated by GSFC. The SOC will generate commands that will be transmitted to the MOC no more than once daily during the operational phase of the mission, and not necessarily on weekends. Command generation and MOC operations will be more intensive during the commissioning phase. 4.0STATEMENT OF WORK During the proposal preparation period, the partner will participate as part of the mission proposal team. Statements of Work (SOWs) are not required to be submitted with the Step 1 proposal. However, they are required before Phase A (Step 2 proposal) work can begin. Therefore, the partner shall provide a draft statement of work during the Step 1 proposal effort that specifies through task statements what the partner is proposing to provide to the mission in Phases A-F. SOWs will include the following as a minimum: Scope of Work, Deliverables (including spacecraft performance data), and Government Responsibilities (as applicable). SOWs need not be more than a few pages in length. 5.0POD RESPONSE INSTRUCTIONS, FORMAT, AND SELECTION CRITERIA 5.1INSTRUCTIONS The respondent shall: •Provide demonstrated flight heritage and associated mission operations of spacecraft similar to the ones proposed. •Demonstrate understanding and quantified experience in the design, fabrication, integration, and testing of the spacecraft system proposed. The response shall describe how the proposed spacecraft and mission operations meet the requirements given in Section 3. •Describe the capabilities of the spacecraft as they apply to the requirements given in Section 3 and provide supporting performance data. •Provisionally define the spacecraft to instrument interfaces. Provide information on the maturity of these interfaces and indicate if the latest configuration has flight heritage and demonstrated on-orbit performance. Provide the total spacecraft uplink and downlink capability. •Provide the on-orbit consecutive life-time capability of the proposed spacecraft (planned and demonstrated). •Identify the technical maturity/qualification of the proposed spacecraft and mission operations concept. If the spacecraft has not already demonstrated the required mission life, the respondent shall describe how these items will be demonstrated, including a timeline for this demonstration, which will not be funded under this effort. •Describe the approach for supporting the Tech DemoAO Step 1 proposal, the Phase A Mission Concept Study (Step 2 proposal), the Phase A Site Visit at a TBD location, and mission development, including the level of support that the partner plans to make available for each activity. •Provide an estimated Rough Order of Magnitude (ROM) cost (in FY19 dollars) from Step 2 selection (Phase B) onward for the all spacecraft activities including design, fabrication, integration and testing, and mission operations. The response shall include a brief discussion of the uncertainty in the cost estimate. 5.2FORMAT The response to this partnership opportunity is limited to 15 pages in not less than 12-point font. Excluded from the page count are the cover letter, title pages, table of contents, and acronym list. Partners may attach additional appendices that further describe their capabilities, although GSFC is under no obligation to include the contents of such appendices in the evaluation of the offer package. The entire offer package, including any cover letter, title pages, and other supporting material, shall be formatted as a Portable Document Format (PDF) file delivered to the E-mail address below. 6.0EVALUATION FACTORS AND CRITERIA The evaluation team will use the following factors in selection and award: 1.Technical Approach (35%). Offerors will be evaluated on their ability to meet the technical requirements given in Section 3. This includes demonstrated understanding of the requirements and proposed approach to meet those requirements. 2.Cost (40%). Offerors will be evaluated on their overall cost and on the reasonableness of cost and schedule estimates. It will be important for the offerors to demonstrate through their proposal an understanding of cost efficiencies available to a Streamlined Class D mission as described in documents available at https://soma.larc.nasa.gov/stp/tdmo/tdmo-library.html. 3.Relevant Experience and Past Performance (25%). Emphasis will be given to demonstrated experience with similar missions. 7.0POINT OF CONTACT Questions about this POD should be directed to Laura Ottenstein (Phone: 301-286-4141, Email: Laura.Ottenstein-1@nasa.gov). 8.0FINAL DUE DATE OF POD RESPONSE The response to the POD is due no later than 5 p.m. EDT on the "Reponses due" date given above in Section 1.0. The electronic PDF document shall be sent to Laura Ottenstein (Email: Laura.Ottenstein-1@nasa.gov). It is the responsibility of potential respondents to monitor the NASA Acquisition Internet Service (NAIS), GSFC Procurement Site http://fbo.gov for information concerning this POD. 9.0FIGURES PLEASE SEE PDF VERSION OF THIS DOUCMENT ATTACHED TO SYNOPSIS!! Figure 1. Stowed DSC and OSC mated to each other and ESPA Grande Ring (left) and shown apart from ESPA Grande Ring (right).   Figure 2. DSC and OSC Deployed Configuration Figure 3. Mission trajectory. Figure 4. Mission timeline. Figure 5. DSC and OSC thrusters.   Figure 6. GSFC-Provided PL Elements on OSC and DSC. Figure 7. Block Diagram of Mission Communications Architecture 10.0ACRONYMS AOAnnouncement of Opportunity CBECurrent Best Estimate C&DHCommand and Data Handling DSCDetector Spacecraft EELVEvolved Expendable Launch Vehicle EOLEnd of Life EPSElectrical Power Supply ESPAEELV Secondary Payload Adapter EUVExtreme Ultraviolet FYFiscal Year GNCGuidance, Navigation and Control GSFCGoddard Space Flight Center HGAHigh Gain Antenna I&TIntegration & Test ICDInterface Control Document IMAPInterstellar Mapping and Acceleration Probe ISPSpecific Impulse MATMedium Acquisition Telescope MEVMaximum Expected Value NASANational Aeronautics and Space Administration NLTNo Later Than NPRNASA Procedural Requirements NTENot to Exceed OSCOptics Spacecraft PDFPortable Document Format PDRPreliminary Design Review PFFSPrecision Formation Flying System PIPrincipal Investigator PLPayload PODPartnership Opportunity Document PSTPhotometric Solar Telescope RCBReleasable Clamp Band ROMRough Order of Magnitude SASolar Array SALMON-3Third Stand Alone Missions of Opportunity Notice SCSpacecraft SCMSmall Complete Mission SEESingle Event Effects SEL1Sun-Earth L1 Lagrangian Point SMDNASA Science Mission Directorate SOWStatement of Work STStar Tracker STPSolar Terrestrial Probe SWaP-CSize, Weight and Power + Cost TASTransverse Alignment System TBDTo Be Determined TIDTotal Ionizing Dose TRLTechnology Readiness Level
 

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