4.1 factors of interest · web viewan iss demonstrated wing would nominally operate at 168 v,...

RFI NNC15ZMX006L

Advanced Solar Arrays for Flight DemonstrationsRequest for Information

It should be understood that there is no explicit or implied commitment for future procurements in this action.

RFI/SSN#: TBD

1. Introduction

A high priority within the Space Technology Mission Directorate (STMD) is the development and demonstration of high-power solar electric propulsion (SEP) technologies required to extend human presence beyond Earth’s orbit. Advancement of solar electric propulsion technology is also expected to benefit a wide range of other commercial and government activities in space. Initial investments by STMD in electric propulsion and advanced solar array systems have raised the technology maturity in these two areas sufficiently to now enable high-power SEP flight demonstration missions. One such demonstration mission concept under consideration is the Asteroid Redirect Robotic Mission (ARRM) to return a small asteroid or boulder from a larger asteroid to a crew-accessible, lunar orbit in the mid-2020’s using SEP, scalable to future missions.

The NASA Technology Demonstration Missions SEP Project, managed by the Glenn Research Center (GRC), is soliciting information from potential vendors regarding the development of 25 kilowatt-class advanced, flexible-blanket solar array wings that could be used for SEP flight demonstration missions such as the ARRM which would use two wings for a 50 kW or higher power configuration. The solar array concept would also require direct extensibility to power levels greater than 100 kilowatts for future applications beyond the initial ARRM mission. Additionally, NASA would like to perform an early flight demonstration of the same solar array wing on the International Space Station (ISS). The demonstration on ISS is intended to be a long-term functional demonstration of the wing and if successful, it would provide power augmentation to the ISS. Additionally, if the ISS demonstration array proves successful, up to three additional identical arrays could be purchased and installed on ISS to provide augmented power.

NASA recognizes that there are significant differences in the design environments for a solar array on a SEP-powered spacecraft operating exclusively beyond low Earth orbit and a solar array attached to the ISS. Despite these differences, NASA is investigating the advantages of an integrated approach for the development of solar arrays for both applications. We are seeking input from industrial solar array providers on the appropriate level of design commonality required for a single combined development of solar arrays for the ISS and deep-space SEP applications to achieve cost savings and efficiencies compared to the independent development and procurement of different solar arrays uniquely designed for each environment.

2. Technical Description

Preliminary requirements and characteristics for the ARRM and ISS solar arrays are listed in the table below. Additional notional information for each row follows in sections 2.1 through 2.14. NASA is soliciting ideas on approaches to meeting all requirements while maximizing commonality.


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To facilitate demonstration of a common solar array for any future NASA program and utilization of power from this demonstration on the ISS, the ISS program will provide the ISS-unique interface hardware built as much as possible to the interface specified by the array provider.

Solar array providers are requested to identify the interfaces to the array, and the ISS program will provide the hardware required to mount to the launch vehicle as well as to the ISS. Hardware provided will include the robotics and EVA interface hardware and the array-to-ISS power system hardware.

Table 1: ARRM & ISS Solar Array Wing Technical GoalsItem ARRM ISS Demonstration Common

Hardware Approach

Proto-flight wings Proto-flight wing(s) Proto-flight wings

Launch Packaging – Stowed Wing Configuration

ARRM with wings in Delta IV, Atlas V, SLS (5 m fairing)

Wing in Commercial Resupply Services (CRS)

Wing for both ARRM in DIV/AV/SLS and wing in CRS

Launch Orientation Vertical, spacecraft bus exterior side wall

CRS unpressurized cargo hold

Vertical, spacecraft bus exterior side wall; CRS unpressurized cargo hold

Release, Phasing, Deployment

Autonomous releases and latching with 28 VDC power

Interfaces for wing tie downs provided by launch vehicle provider, powered at 28 VDC in CRS – power level TBD, 120 VDC at 2B site autonomous and/or manual releases & latching

Fully autonomous releases, powered and/or manual releases and latching with 28 VDC and 120 VDC power available for heaters and motors

Release, Phasing, Deployment Sensors

As needed As needed, defined by array provider and as required for ISS safe deployment

Full sensor set driven by ARRM (nearly all sensors will be inactive for ISS application)

Deployment Redundancy

Single fault tolerant, (for non-structural elements), autonomous

Single fault tolerant, (for non-structural elements), autonomous. EVA crew provides no assistance for wing phasing and deployment

Single fault tolerant, (for non-structural elements), autonomous

Deployed Wing Configuration

Mounted to 1-5 m off-set boom/yoke and solar array

Mounted to interface structure. Structural characteristics must be

Wing with dual mounting capability (mount to yoke/boom,


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drive assembly (SADA)

known for all phases of array deployment to inform contingency recovery options (EVA, jettison, etc.)

mount to interface structure)

On-orbit Latency prior to Wing Deployment

None 1-4 weeks (TBR, in CRS, 2 days (at external 2B site)

1-4 weeks (TBR, in CRS), 2 days (at external 2B site)

Wing Power, Voltage, Current

25 kW (TBR), 300 V (TBR), ~1.3 Amp string Isc, 5 Amp channels

20 kW or greater, 168 V, <1 Amp string Isc, <3 Amp channels (generated), Voc< 320 V

25 kW class for ARRM/ ISS, 168 V, <1 Amp string Isc, <3 Amp channels (generated), Voc<320 V

Array Mass 180 kg (TBR) <230 kg for array and <100 kg for 2B SAW interface structure

<230 kg for array and <100 kg for 2B SAW interface structure

Flight Performance and Technology Development Sensors

As needed temperature, current, voltage, accelerometers (TBD)

As needed As needed, full sensor set driven by ARRM (nearly all sensors will be inactive for ISS application)

Mission Earth orbiting, cis-lunar, heliocentric (0.7 to 1.8AU) for 7 years

LEO, 400 km, 51.6 deg inclination, 10 years

10 yr LEO and 7 yr cis-lunar, heliocentric (0.7 to 1.8 AU) operations

Thermal Extremes & Environments

+/-60C deployment, +150C/-200C operating, <100’s cycles, RCS plume heating, free molecular heating with possible thermal shields required

+40C/0C deployment, +/-80C operating, < 100,000 cycles, no plume heating, no free molecular heating, no thermal shields

+/-60C deployment, +150C/-200C operating, < 100,000 cycles, with RCS plume heating and free molecular heating, thermal shields as needed.

Unique Driving Environments

EP plasma, EP plume sputtering ions (level TBR), vacuum charging, 1E14 DENI, 1E20 AO fluence

LEO plasma/charging, no sputtering ions, 1E13 DENI, 2E22 AO fluence

EP & LEO plasma, EP plume sputtering ions, vacuum charging, 1E14 DENI, 2E22 AO fluence

Deployed Strength /Stiffness

>0.1-g / >0.1 hz for wing on off-set boom mounted to a SADA (<TBD m max displacement)

>0.005-g, >0.1 hz, for wing mounted to the interface structure <2 m max displacement

>0.1-g, >0.1 hz, <2 m max displacement for both mounting options

2.1 Hardware Approach


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Proto-flight wings will be utilized for both ARRM and ISS. The hardware may be provided with reduced oversight from NASA provided that the hardware is designed and constructed using best industry standards. The first demonstration of the hardware is anticipated to be on ISS and would be considered a technology demonstration.

No flight spares will be built, and the minimum amount of development and qualification coupons. Components and subsystems will be at TRL 6 by Solar Array PDR (ARO + 5 months).

2.2 Launch Orientation, Packaging and Spacecraft AttachmentThe ARRM arrays are expected to be mounted vertically to the spacecraft bus exterior sidewall such that the spacecraft can fit within the fairing of a Delta IV, Atlas V, or SLS launch vehicle. The bus length available for the array is ~4 m long, with 1 to 2 m (TBR) unsupported additional overhang length.

The initial ISS demonstration array will be transported in the cargo bay of a commercial vehicle. The longest dimension available is <3 m. The space station remote manipulator system (SSRMS) special purpose dexterous manipulator (SPDM) end effector will be used to transfer the wing from the transport vehicle to its intended position over the existing 2B wing. Up to three additional optional arrays would utilize identical transportation and ISS integration procedures and occupy similar solar array wing locations in front of existing ISS arrays.

2.3 Release, Phasing, DeploymentThe ARRM arrays will deploy autonomously using releases and latch securely in place. 28 VDC power required for deployment will come from the ARRM.

The ISS array will deploy autonomously, 28 V power is available in the CRS, and 120 VDC is available at the 2B site for deployment.

Structural/mechanical/electrical interfaces for the SAW tie-downs will be provided by launch vehicle provider.

2.4 Release, Phasing, and Deployment SensorsThe ARRM wing will require sensors to confirm latch release and full deployment.

For the ISS array, sensors as defined by array provider and as required for ISS safe deployment.

2.5 Deployment RedundancyBoth the ARRM wings and ISS wing will be single fault tolerant (for non-structural elements) for deployment and fully autonomous. For the ISS wing, assume EVA crew provides no assistance for wing phasing and deployment.


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2.6 Deployed Wing ConfigurationThe ARRM arrays will be deployed such that the nearest point on the wing stays out of a 60 degree (TBR) keep-out zone along the axis of the electric thruster plume(s) with other driving factors for the deployed wing configuration including: managing bending moments from inertial loading and Orion RCS plume moments, minimizing SEP module RCS heating & pressure loads, ability to meet SEP module deployed wing KOZ*, ability to meet SEP module deployed wing LOS**, minimizing deployed wing radiator view factor blockage, minimizing deployed wing shadowing (from the asteroid, from the deep space habitat module), minimizing spacecraft MOI (to enhance ACS, GNC designs), minimize local EP induced charge-exchange plasma density, minimize the asteroid dust flux source, minimize boom mass and stowed volume. Structural characteristics must be known for all phases of array deployment to inform contingency recovery options (EVA, jettison, etc.).

*KOZ (deployed wing keep-out-zone requirement considerations): avoid wing to vehicle or asteroid contact including protuberances under worst case appendage deflections, avoid EP plume impingement on wing for spacecraft momentum management, avoid EP plume scattered Xe+ ions with sputter energy levels, standoff distance for EVA translation corridors, standoff distance for asteroid capture bag and asteroid surface.

**LOS (deployed wing impact for line-of-sight requirements), having wing avoid LOS for spacecraft communication antennas, star trackers, other navigation sensors; providing for wing-mounted camera view angles and LOS for proximity operations at asteroid and with Orion; providing more favorable viewing field with less solar array shadowing for final approach to asteroid.

For ARRM, a 1 to 5 m (TBR) off-set boom/yoke will be required, and a single DOF SADA will be required to track the sun. A reference configuration is shown in Figure 1.

The ISS array will fit closely in front of the existing ISS 2B wing and be removable in case of failure to restore the functionality of the existing wing. Additional optional arrays would similarly fit closely in front of other existing ISS wings. The existing Solar Alpha Rotary Joint (SARJ) will be used for sun tracking. A reference configuration is shown in Figure 2 and Figure 3.

2.7 On-orbit LatencyThe ARRM wings will deploy immediately following spacecraft orbit insertion.


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The ISS wing may remain stowed for 1 to 4 weeks in the CRS at the ISS before being scheduled for transfer and 2 days at the external 2B site before deployment. During this time there will be <17 eclipse and shadow-based thermal cycles per day of +TBD C/ -TBD C

2.8 Wing Power, Voltage, CurrentThe nominal power output of the ARRM wings will be 25 kW, operating at 300 V (TBR), generating 5 Amps per channel with ~1.3 Amp per string Isc using XTJ Supercells.

The ISS can also accept a 25 kW wing and would require no less than a 20 kW wing. An ISS demonstrated wing would nominally operate at 168 V, generating >.75 Amp IMP, <1 Amp per string Isc, <3 amp channels (generated) and VOC <320 V. The new wing will shadow some of the current wings, reducing their output. Greater power levels up to 33.8 kW and with different current and cells to match the nominal ARRM wings are acceptable provided savings in recurring engineering and costs are justified through commonality.

The 20 kW estimate is based on the need to provide a net 10 kW power increase to ISS relative to the current ISS 2B Solar Array Wings. The new wing will shadow some of the current wings, reducing their output approximately 20 kW.

Greater power levels up to 33.8 kW and with different current and cells to match the nominal ARRM wings are acceptable provided savings in recurring engineering and costs are justified through commonality.

The ARRM-sized common wing operated on ISS is expected to provide slightly less than 25 kW (about 23 kW). This is due to mounting the ARRM solar array in front of the ISS 2B SAW and operations in LEO, which will lead to a 15-20 degree C hotter operating temperature and resulting lower power output. This assumption may be design specific.

2.9 MassThe ARRM wings are expected to have a mass allocation of nominally 180 kg per wing.

The ISS wing(s) are expected to have a mass allocation of no more than 230 kg to fit within the ISS structural mass limit. An additional 100 kg is allocated for the interface structure.

2.10 Flight Performance and Technology Development Sensors


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ARRM per wing sensors to include measurement of 3 PV temperatures, 2 structure temperatures, PV segment current and voltage, 3 accelerometers.

For ISS the required instrumentation will be supported as defined by the array provider.

2.11 MissionThe ARRM mission is Earth orbiting, cis-lunar, heliocentric (0.7 to 1.8 AU) for 7 years (TBR).

The ISS wing will fly in LEO at 400 km and 51.6 degree inclination. The expected lifetime for this wing to provide power augmentation is 10 years.

2.12 Thermal Extremes and EnvironmentsThe ARRM wings:+/-60C deployment, +150C/-200C operating, <100’s cycles, RCS plume impingement heating while stowed and/or deployed, ascent free molecular heating while stowed with possible thermal shields required

The ISS wing:+40C/0C (TBR) deployment, +/-80C (TBR) operating, up to 100,000 cycles, no plume heating, no free molecular heating, no thermal shields

2.13 Unique Driving Environments

The values below for ARRM and ISS are engineering estimates:

The ARRM wings:EP plasma, EP plume sputtering ions, vacuum charging, 1E14 DENI, 1E20 AO fluence

The ISS wingLEO plasma/charging, no sputtering ions, 1E13 DENI, 2E22AO fluence from sweeping ram

2.14 Deployed Strength and StiffnessThe ARRM wings: >0.1-g / >0.1 hz for wing on off-set boom mounted to a SADA (<TBD m max displacement)

The ISS wing:> .03 psf plume pressure/.005 g > 0.2 hz minimum for 1st mode, <2 m max displacement

3.0 General


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This RFI is used solely for information planning purposes and does not constitute a solicitation. In accordance with FAR 15.201(e), responses to this RFI are not offers and cannot be accepted by NASA to form a binding contract. NASA is under no obligation to issue a solicitation or to award any contract on the basis of this RFI. The information provided in response to this RFI will not be made public in an effort to protect any propriety company information. Nonetheless, respondents should clearly and properly mark any propriety or restricted data contained within its submission so it can be identified and protected. Respondents are solely responsible for all expenses associated with responding to this RFI. Responses to this RFI will not be returned, and respondents will not be notified of the result of the review.

4.0 Information Requested

Information is requested to better understand the possible approaches to a potential RFP and possible contract award for ARRM and ISS solar array wings prior to finalizing the system requirements.

NASA is seeking industry’s perspective on potential efficiencies/inefficiencies that may accrue by developing, building and testing three units (with an option of three additional units) of a common design as compared with two ARRM wings and a separate ISS wing (or wings).

NASA requires input to understand what performance and schedule impacts, if any, this common design would have on the ARRM mission.

To determine these impacts, we would like to identify expected changes to both recurring and nonrecurring costs caused by a common design based on the consideration of each of the factors listed in listed in Section 2, as well as any other parameters that may have a strong effect. The requirements and characteristics listed for each mission are the minimum required, so a common array would have to meet both.

Of particular interest are comments about how designing to one set of requirements that encompasses both applications may impact nonrecurring and recurring costs, as compared to two independent designs.

Quantities of interest include one solar array for ISS (with an option for three additional arrays) and two solar arrays for ARRM, each sized for nominally 25 kW class at BOL at 1 AU for ARRM mission, (ISS power output would be lower due to installation at the 2B site). Insight into how any cost savings or penalties would scale with additional unit purchases is also of interest.

Elements of interest to be included in that assessment are listed below. Responses will be used for NASA’s preparation of a potential future RFP or RFPs.


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4.1 Factors of Interest

A. Solar array design(s) considered, including integration with the ISS and cargo delivery vehicle and with the ARRM vehicle. Include a design summary (mass, stowed/deployed configurations, key design features, performance) of the common wing.

B. Summary comparison of the features of a common design, unique designs

and/or a hybrid option from perspective of the design and development approach, including: increases or reductions in non-recurring engineering cost for uniqueness/commonality; increases or reductions in recurring cost of multiple copies of common arrays versus a similar number of arrays of unique design; and implementation risk factors associated with commonality versus uniqueness.

C. Assuming a common array strategy, identify the relationship and potential cost saving regarding the number of solar arrays procured for both ARRM and ISS. For example a common array solicitation may request two ARRM arrays and one ISS demonstration array with an option of up to three additional ISS demonstration arrays.

D. Address if the described common array may have any commonalities with future commercial advanced arrays that could reduce the unit costs and if any adjustments to the common array specifications could be made to better align commercial and government applications.

E. Driving requirements that need further definition, evaluation, and rationale if a contract were to be let before the ARRM system requirements review.

F. Potential benefits, issues, or requirements divergence caused by including ISS solar array requirements in designing, developing, integrating, and operating the ARRM mission.

G. Potential benefits, issues, or requirements divergence caused by including the ARRM requirements in designing, developing, integrating, and operating the ISS solar array(s).

H. A list of technology developments and associated complexity and cost that would be needed to meet ISS requirements that have not already been included in existing solar arrays that have been designed for the nominal ARRM. Example ISS requirements driving technology development include CRS launch packaging, EVA/robotic interfaces for wing handling and deployment and potential jettisoning, high count thermal cycles and high atomic oxygen (AO) fluence erosion of organic resins/polymers and chemical conversion of coverglass optical coatings.


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I. Identification of any requirements either stated above or assumed for a common array that would preclude a response from your company as an otherwise interested industry partner. This might include items such as pre-deployed volume, power or voltage levels, or environmental limits such as thermal cycles.

J. Identify any changes to the stated or assumed common requirements from

Factor A, above, that would allow the responder’s products to better meet the intent and goals of the requirements.

K. While NASA’s intention is to consider a common array design, identify if there are recommended minor design variations that properly address disparate uses, but also manage to avoid the possible additional incurred costs associated with offering two separate and unique array requirements. Examples of such minor design variation might include how structures and materials are protected from atomic oxygen in a LEO environment or how the voltage ratings could be varied between arrays of different applications. This input does not require design details and should be limited to those areas of design that;

i. are expected to reduce overall costs from a single, unchanging common design,

ii. may be accommodated without requalification of the array.

L. Identify any stressing limitations of the ISS launch packaging envelope and provide what increases in the stated bounding envelope would lessen the impacts on array design.

M. How schedule flexibility and impact of multiple deliveries affects both the overall cost and schedule.

N. Provide an evaluation of whether a cost-effective hybrid strategy exists that would rely upon some common elements (e.g. structure and/or mechanism) and some mission unique elements (cells, wiring, voltage).

O. Cost and risk of increasing the power per wing (for ARRM) from 25 kW up to a maximum of 50 kW/wing.

P. An engineering cost estimate for the three common design solar array wings and support hardware (launch costs not included) as compared to two unique or hybrid designs.


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Figure 1. Representative ARRM vehicle showing generic stowed solar array wing options.


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Figure 2. ISS demonstration solar array will be mounted on the ISS 2B SAW left and right blanket box trunnion fittings at the wing base.

Figure 3. Notional concept for ISS demonstration wing integration at 2B site


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5. Formatting Instructions for SubmittalResponses should be limited to the information requested in Table 1 below.

Table 1. Volumes Requested

Volume I Advanced Solar Arrays for Flight Demonstrations Response

1 hard copy, 2 electronic copy

Responses to this RFI shall be submitted in writing and electronic media postmarked no later than 5:00 PM EST on January 30, 2015.

Request for Information: Responses must be submitted to:Leahmarie Koury, Contracting OfficerGlenn Research CenterMS 60-121000 Brookpark RoadCleveland, Ohio [email protected]

The response must be sent as one printed hardcopy and electronically as single Microsoft Word .docx and/or Microsoft EXCEL.xlsx file for each response on a CD or DVD.

For the purposes of this RFI, an Engineering Cost Estimate (ECE) is defined as a high-level estimate without detailed line item breakouts or rates that provides an educated financial estimate to be reported in FY15 dollars.


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mailto:[email protected]

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Acronym List

ARO – After Receipt of OrderARRM – Asteroid Redirect Robotic MissionAO – Atomic OxygenAU – Astronomical UnitBOL – Beginning of LifeCRS – Commercial Resupply ServicesDENI- Damage Equivalent Normally IncidentDOF – Degree of FreedomEP – Electric PropulsionEVA – Extravehicular ActivityFAR – Federal Acquisition RegulationKOZ – Keep-Out-ZonekW – KilowattsLEO – Low Earth OrbitLOS – Line-of-SightMOI – Moment of InertiaIsc – Short Circuit CurrentISS – International Space StationPDR – Preliminary Design ReviewPV - PhotovoltaicRCS – Reaction Control SystemRFI – Request for InformationSADA – Solar Array Drive AssemblySARJ – Solar Alpha Rotary JointSAS – Solar Array SystemsSAW- Solar Array WingSEP – Solar Electric PropulsionSLS – Space Launch SystemSPDM – Special Purpose Dexterous ManipulatorSSRMS – Space Station Remote Manipulator SystemSTMD – Space Technology Mission DirectorateTBD – To be determinedTBR – To be revisedV - Volts


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4.1 factors of interest · web viewan iss demonstrated wing would nominally operate at 168 v,...

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