AOS-I-ATS-RFP-3.pdf

AOS-I-ATS-RFP-3

Request for Proposal (RFP)

Title: AOS Inclined ATS RFP Mission Target Values and Instrument Accommodations / Allocations

Date: v8/27/21

Reference Number: RFP- # TBD RFP Responses TBD

Table of Contents

1.0 Introduction

2.0 AOS Inclined Instrument Accommodation and Target Values

2.1 Mission Assumptions and Constraints

2.1.1 Orbit Parameters

2.1.2 Operational Modes and Fault Detection

2.1.3 Mission Lifetime

2.1.4 System Risk Tolerances

2.1.5 Launch

2.1.6 Orbital Debris

2.1.7 Propulsion and Delta-V

2.1.8 Spacecraft Availability

2.1.9 Timing and Position

2.1.10 Data Downlink, Communications, and Mission Operations

2.1.11 Onboard Data Storage and Collection Rates

2.1.12 Command Encryption

2.2 Payload Accommodations

2.2.1 Instrument Payloads

2.2.2 Payload Instrument Constraints: Requirements and Targets

2.2.3 Science Operations Pointing

2.2.4 Instrument Alignment Stability

2.2.5 Thermal Interface

2.2.6 Contamination Control

2.2.7 Electronics and Interfaces

2.3 Hosted Payload Approach (Option 2 Specific)

2.4 Development Approach

2.4.1 Systems Engineering

2.4.2 Reviews

2.4.3 Analysis

2.4.4 Documentation

2.4.5 Development Units

2.4.6 Design Rules

2.4.7 Performance and Environmental Verification

2.5 Safety and Mission Assurance Processes

2.6 Access to Space (ATS) Insurance

2.6.1 Insurance for the Traditional Spacecraft Development Approach (Option 1)

2.6.2 Insurance for the Hosted Payload Approach (Option 2 Specific)

2.7 References

3.0 Information Requested

Appendix A - D

1.0 INTRODUCTION

This document provides a mission summary and provides target values (i.e. allocation / accommodation) for instrument accommodation and allocation to the multiple Atmosphere

Observing System (AOS) inclined host spacecraft.

2.0 AOS INCLINED INSTRUMENT ACCOMMODATION AND TARGET VALUES

The following reference parameters should be considered in the responses to this Access to Space (ATS) Request for Proposal (RFP). Detailed instrument allocation and accommodation values are in Appendix B. For a summary of all target values and assumptions refer to Appendix

D.

2.1 Mission Assumptions and Constraints

2.1.1 Orbit Parameters

The planned orbit is 407km circular (+/- 10km), 40-55 degree inclination. Note: The AOS Program is interested in information from the respondent if there are vendor preferences or cost savings enabled by specific orbital choices.

2.1.2 Operational Modes and Fault Detection

The spacecraft should have operational modes that include but are not limited to: operation with instruments in standby, engineering, operation in science mode (instruments operating), safe hold, delta-V, and load shed capability. Any additional spacecraft modes shall be included in the trade space. The spacecraft should be powered during launch. The spacecraft should provide fault detection and safe action for the instruments.

2.1.3 Mission Lifetime

3 years starting at the completion of 3 months of commissioning, with planned consumables for minimum 5 years of operations, plus disposal. For lifetime assessments a launch date of July 2028 can be assumed.

2.1.4 System Risk Tolerances

Each AOS spacecraft and instrument payload complement is currently envisioned to be Risk Class C per NPR 8705.4, Appendix D. Concepts that follow alternative approaches will be considered. Related information is requested per 2.5. Information is requested regarding the reliability of the spacecraft for the mission lifetime per 2.1.3. Depending on the orbital debris approach taken, the spacecraft will have a reliability of deorbit per the orbital debris requirements per 2.1.6.

2.1.5 Launch

The current plan is to launch all spacecraft needed to fly the instrument payload complement quantities per Section 2.2.1 on a single launch vehicle. Information, configurations, and envelopes for launching the respective quantities of systems is requested. The scope of work for the spacecraft vendor should include the system to deploy the spacecraft from the launch vehicle, including cost. It should be assumed that the launch vehicle will provide the signal to deploy the individual spacecraft. The spacecraft vendor will perform all launch site activities, including pre-launch preparation and testing, in coordination with the launch vehicle provider. The spacecraft vendor should assume the launch will occur in the continental United States.

Launch vehicle or launch service costs should not be included in cost estimates as it is envisioned that the launch vehicle will be procured separately by the AOS Program, unless a spacecraft and/or spacecraft deployment solution is inherently linked to a launch vehicle by the solution provider. For this RFP, it is only requested that respondents state whether the spacecraft solution is so linked.

2.1.6 Orbital Debris

The proposed spacecraft must plan for orbital debris and spacecraft re-entry requirements per NASA STD-8719.14, Process for Limiting Orbital Debris.

2.1.7 Propulsion and Delta-V

The proposed spacecraft solution(s) must provide adequate change in velocity (Delta-V) to accomplish the following:

• Orbit insertion (correction of launch vehicle errors) and orbit maintenance for planned mission life maintaining the 407 +/-10 km altitude window as measured at the equator.

• Maintain ground track registration of the radar and lidar spacecraft (if on separate spacecraft) should be maintained within +/- 3 km at the equator.

• Maintain ground track registration of each spacecraft to within +/- 10 km as measured at the equator.

• Maintenance of spacecraft temporal separation to within the values specified in Section

2.2.2.

• Cold-side spacecraft view maintenance over the mission life (see Section 2.2.5).

• Momentum unloading, as applicable

• Disposal, based on the orbital debris plan per Section 2.1.6

Note: All maneuvers should be executed in an appropriate timeframe or orientation to meet the spacecraft availability parameter per 2.1.8. If this is not feasible, please describe the system capability.

2.1.8 Spacecraft Availability

The spacecraft should be in an operational mode to support science operations greater than 96% of the time, taking into account times for maneuvers, momentum unloading, or other operations that cause mission and payload accommodation parameters, such as pointing per 2.2.3, to be violated.

2.1.9 Timing and Position

The spacecraft should use GPS to determine position and provide time at the tone and timing pulses to each instrument. UTC registration accuracy should be within 100 msec. Within a given spacecraft, time tag knowledge between instrument measurements should be within 10 msec.

2.1.10 Data Downlink, Communications, and Mission Operations

The AOS Program solicits responses for a ground system solution in support of a constellation of

2-3 spacecraft in Low Earth Orbit (LEO). The six instruments will be distributed across the constellation of spacecraft. The solution should address the ground system solution for both the pre-launch observatory-level integration and test and the post-launch operational phases of the mission. The solution should be an end-to-end ground system, including the use of antenna ground stations and a terrestrial communication network. Traditional Mission Operations Center (MOC) functions, such as mission planning, trending, flight dynamics, and real-time operations should be addressed from a multi-spacecraft operations standpoint. The response should include an approach to optimize/integrate a separate Science Operations Center (SOC) and multiple Instrument Operation Centers (IOC).

Specifically, the responses should include capabilities to enable low latency transmission of the instrument science data to the ground-based science data processing facilities. While science data latencies of less than 1 hour are highly desirable, the AOS Program is interested in solutions with latencies of up to 6 hours. Innovative approaches to data downlink, such as via Optical intersatellite link (OISL) communications, will be considered. If addressed, communication approaches should include downlink, uplink, and commanding concepts.

Reliable and cost-effective solutions should be qualified. The use of industry standards to meet FISMA Medium controls should be addressed, as should any cyber-security solutions. Finally, staffing and automation considerations should be discussed.

2.1.11 Onboard Data Storage and Collection Rates

On board data storage should be sufficient to store at least 72 hours of science data and telemetry in worst case collection rate conditions (typically full sun). Instrument data rates are provided in

Appendix B.

2.1.12 Command Encryption

Commands to the spacecraft should be encrypted.

2.2 Payload Accommodations

The instrument payload complement will be Government Furnished Equipment to the spacecraft vendor. There are several types of instruments planned for inclusion in the inclined orbit project.

The instrument payload configuration can be adjusted to fit with the proposed multi-spacecraft capability and overall mission architecture, as part of the study effort to determine the concepts that feasibly meet science objectives and mission constraints at the lowest cost.

For example, there are six AOS instruments. There is one significant constraint with respect to instrument accommodation: the Tandem Stereographic Camera consists of two units and the two units must be separated on two spacecraft (see the constraints listed in Section 2.2.2). As such, a vendor could accommodate five instruments on one spacecraft with a camera unit on a second spacecraft or the vendor might accommodate the remaining five AOS instruments on two additional spacecraft.

In our proof-of-concept design study, we developed two copies of a rideshare spacecraft and accommodated three instruments on each spacecraft (see Appendix C for a summary of our design study). Our plan was to minimize the spacecraft cost by building two copies of the same bus and minimize launch costs by flying multiple rideshare spacecraft on the same launch vehicle. This deign also allows for a large rideshare spacecraft to ride atop the launch stack.

Responses should examine the required instruments, examine their spacecraft capabilities, and suggest the most cost-effective approach to accommodate our six instruments and meet our science requirements (see Appendix A).

2.2.1 Instrument Payloads

For the purposes of this RFP and to garner consistent responses, five (5 different; 6 total) notional instruments on 2 – 3 spacecraft should be assumed when developing plans and cost estimates to develop, integrate, and test systems to accommodate the instrument payload complement. The notional instruments with Maximum Expected Values (MEV) are summarized in Table 1 with the full list of parameters in Appendix B. For spaceacecraft sizing use MEVs where available. The AOS

Table 1. AOS Instrument Payload MEV Characteristics

Instrument1 Qty Mass

[kg, MEV] Power

[W, MEV]

Data

[Megabits/sec, MEV]

Radar 1 190 336 5 Mbps

Lidar 1 152 336 12 Mbps

Radiometer 1 52 70 0.26 Mbps

Polarimeter 1 45 65 100 Mbps peak (daylight only)

Approx. 50 Mbps (orbit average)

Tandem Stereographic

Camera2

2 (1 per SC) 21 23 20 Mbps peak (daylight only)

Approx. 10 Mbps (orbit average)

Notes:

1. The AOS Program intends to supply notional instrument CAD models to the spacecraft vendor but in the event models can’t be provided the volumes contained in Appendix B can be used as stand-ins.

2. Camera instrument is a collection of 3 cameras with 2 active at a given time while in daylight.

2.2.2 Payload Instrument Constraints: Requirements and Targets

The following are the AOS instrument requirements (r) and targets (t):

1) The spacecraft carrying the Tandem Stereographic Camera unit #1 and the spacecraft carrying the Tandem Stereographic Camera unit #2 must be separated by 45 seconds along-track (+15s / -10s). (r)

2) The six (as above) instruments will be integrated on multiple (2 – 3) rideshare spacecraft and the rideshare spacecraft will launch on a single launch vehicle (Note: 1, 2) (r)

3) All spacecraft are in nearly the same orbit: 407 km circular (+/- 10km), 40-55 inclination. Note: the line of nodes on one spacecraft will be differ by ~0.2 to that of the other spacecraft to account for earth rotation at equator for cross-track spatial co-registration. (t)

4) Spatial co-registration of radar and lidar footprints within 200m cross-track. (t)

5) Data latency of 1 hour or less is desired for the radar and microwave radiometer, from instrument acquisition to the ground data system (GDS). (t)

6) Data latency less than 6 hours is required for all data to be downlinked. (r)

7) The grouping of which instruments on which spacecraft can be optimized for cost, but the following is a desire:

a. The lidar/radar on the same s/c or the lidar/radar can be maintained less than 60 seconds apart on different s/c. (t)

b. The lidar/polarimeter on the same s/c or the lidar/polar can be maintained less than 60 seconds apart (along-track) on different s/c. (t)

c. The radar/radiometer on the same s/c or the radar/radiometer can be maintained less than 60 seconds apart (along-track) on different s/c. (t)

Notes:

(1) Rideshare spacecraft (RSC) include ESPA, ESPA Grande, Propulsive ESPA, Propulsive ESPA Grande, or other spacecraft and adaptors that allow multiple spacecraft on a single launch vehicle.

(2) Implementing AOS with rideshare spacecraft allow other rideshare spacecraft to ride with

AOS and it may allow another a spacecraft to fly on top of the AOS Integrated Payload Stack

(IPS).

2.2.3 Science Operations Pointing

The spacecraft nadir face should be fixed within the control and knowledge parameters below, dictated by the tightest needs of the payload instrumentation (polarimeter and stereographic camera instrument example below) located in Appendix B:

Pointing Control: 43 arc-seconds, 3σ Pointing Stability: 10 arc-seconds, 3σ Pointing Knowledge: 18 arc-seconds per axis, 3σ Jitter: 5 arc-seconds over 0.5 sec RSS, 3-axis

2.2.4 Instrument Alignment Stability

The spacecraft should maintain instrument alignment stability to within the parameter below, dictated by the tightest needs of the payload instrumentation:

Alignment Stability: 20 arc sec per axis 3σ

This includes effects from environments and other disturbances. It can be assumed that bias errors will be characterized on-orbit during commissioning to determine and apply bias corrections.

2.2.5 Thermal Interface

The thermal interface to the instruments, including instrument electronics, will be thermally regulated with the spacecraft providing temperature sensors and heaters to monitor and maintain each instrument mounting interface within the ranges below:

Operational Temperature Range: -10 to +40 degree C Survival Temperature Range: -20 to +50 degrees C

The amount of allowable heat transfer across the mounting interface will be limited in agreement between the instruments and spacecraft. Other than this thermally controlled mounting interface, the instruments are responsible for, and will be delivered with, the necessary subsystems for their own thermal management. This includes internal heaters, heat pipes, and radiator surfaces as necessary. The spacecraft surfaces within view of the instruments will be covered with multi-layer insulation (MLI) to minimize radiative coupling with the spacecraft.

The instruments are designed with the assumption of a cold-sky field-of-view (FOV) available during their operations. As such, the spacecraft will ensure each instrument a view to the cold sky for instrument heat rejection. In addition, the spacecraft will need to ensure a ‘cold-side’ to the spacecraft throughout the annual cycle via mission ConOps or some other means. In the design exercises conducted by NASA for a 65-degree inclination, this entailed approximately 9 (nine) 180-degree spacecraft yaw maneuvers per year to maintain the cold-side of the spacecraft and instruments.

Notes

1. The science instruments can perform in both ram orientations to account for yaw flips.

2. Appendix B contains instrument radiator side information. For the Radar, a 1.5 m^2 radiator panel is required near the vicinity of the solid dish. A notional panel may be designed by the spacecraft vendor and it can be assumed that this radiator would be GFE.

2.2.6 Contamination Control

Some of the instruments are sensitive to contamination by both particles and hydrocarbons.

Adequate precautions must be taken during spacecraft Integration & Testing (I&T) to assure the on-orbit performance of the instrument. Materials used in the spacecraft must be selected to be consistent with meeting low outgassing rates.

During integration and up until launch, some instruments will require continuous purging with dry nitrogen, although brief interruptions on the scale of a few hours in controlled environments are acceptable. Surface cleanliness levels will be monitored through witness samples particle fall-out plates, direct surface cleanliness testing, and UV and white light inspection. Cleaning operations will be performed as necessary. At a minimum, a class 10,000 environment will be required whenever AOS is unbagged for integration and test operations once optical instruments are integrated to the spacecraft.

2.2.7 Electronics and Interfaces

The spacecraft should provide the following electrical interfaces to each instrument:

1. +28 V primary power services sized for instrument power

2. RS-422 communication services or similar for instrument command and telemetry

3. High-speed data interface sized for instrument data rate

4. Pulse Per Second (1 PPS) services for time synchronization

Plans to develop the appropriate spacecraft simulator GSE should be included to enable functional test of integrated payloads in parallel with spacecraft development. Costs associated with supporting GSE should be included.

2.3 Hosted Payload Approach (Option 2 Specific)

The instruments could be hosted on a spacecraft that is part of another system architecture. The highest science value is obtained by instruments at the orbit specified in this RFP. The mission assumptions and constraints in Sections 2.1 and 2.2 still apply to systems that are proposed to host instruments. When developing the Phase A-D cost estimate for hosted payloads (HP), the cost should include the instrument accommodation cost, the launch cost, and any hosting fees, in a single fixed price cost. The Phase E-F cost estimate should include the price per month/year for the ground data system and the mission ops costs.

2.4 Development Approach

The development approach reference parameters described in this section have been defined as a point of departure, based upon processes used to develop similar NASA missions in the past.

Proposed development approaches that deviate from these reference parameters will be considered. To appropriately evaluate various approaches and relative risks, it is important for the study team to understand how proposed approaches adhere to, and/or differ from, the reference approaches.

2.4.1 Systems Engineering

The spacecraft vendor should perform the necessary systems engineering (SE) required to ensure that the spacecraft meets all of the performance, interface, and implementation requirements of the mission, including the analyses, flow-down of technical requirements, allocation of system budgets, verifications for the spacecraft, definitions of interfaces, technical risk evaluations, system design tradeoff analyses, requirements for GSE, orbital performance analysis, flight software requirements analysis, and lower level requirements (e.g. subsystem, components, assemblies, parts). This includes documenting all information from the design, qualification testing, acceptance testing, and compatibility testing of the hardware and software, together with analysis and assessment of the data with respect to expected performance.

2.4.2 Reviews

As part of the development approach, the spacecraft vendor should conduct and provide technical and programmatic data for the following reviews:

• Spacecraft System Requirements Review (SRR)

• Spacecraft Preliminary Design Review (PDR)

• Spacecraft Critical Design Review (CDR)

• Instrument Integration Readiness Review (IIRR), one per spacecraft

• Full System (Observatory) Pre-Environmental Review (PER), one per spacecraft

• Full System (Observatory) Pre-Shipment Review (PSR), one per spacecraft

• Observatory Acceptance Review (OAR), one per spacecraft

These reviews should meet the NASA/GSFC Criteria for Flight and Flight Support System Lifecycle Reviews, GSFC-STD-1001A.

After each review, the spacecraft vendor will provide formal responses to all request for actions (RFAs) to the AOS Program for approval.

In addition to the above reviews, the spacecraft vendor will provide support to these mission reviews:

• Mission Preliminary Design Review (MPDR)

• Mission Critical Design Review (MCDR)

• Mission System Integration Review (MSIR)

• Mission Operations Review (MOR)

• Flight Operations Review (FOR)

• Flight Readiness Review (FRR)

• Launch Readiness Review (LRR)

The spacecraft vendor will also conduct monthly status reviews (MSRs) at the either virtually or at the Contractor’s facility (as conditions warrant) to review the technical, schedule and programmatic activities. At a minimum, these reviews should include the status of work being performed (e.g., schedule and milestone progress), changes to design parameters and technical performance metrics, and description and status of technical issues, including anomalies and mishaps.

In addition to the meetings and reviews described above, the vendor will support periodic, informal meetings and telecons with the AOS Program.

2.4.3 Analysis

Appropriate analyses should be performed to validate the design will meet requirements with appropriate margins, or verify requirements, including:

• Structural and dynamic analysis

• Functional performance analysis

• Thermal analysis

• Contamination analysis (i.e. plume)

• Analysis necessary to demonstrate margins to GOLD rules per Section 2.4.6

2.4.4 Documentation

The spacecraft vendor will develop, deliver, and maintain all documentation for the observatory and its interfaces, including, but not limited to:

• Spacecraft Performance Specification

• Spacecraft Design and Verification Analyses

• System Performance Verification Plan

• Instrument Interface Control Document (IICD)

• Telemetry and Command Requirements Documentation and Procedures

• FSW Documentation and Procedures

• Test Plans and Procedures

• External Interfaces, Models and Analysis

• Flight Operations Ground System Interface Documentation (Ops ICD)

• Observatory-to-Launch Vehicle Interface Control Documents (LV-ICD) and Launch Vehicle Analysis

• Observatory and GSE Storage, Transportation and Handling Plan

• Observatory Launch Site Operations and Test Plan

• Observatory Launch Site Operations and Test Procedures

• Flight Operations Support Plan and Training

• Spacecraft Operations Description Manual

• Engineering Change Proposals, Deviations and Waivers

2.4.5 Development Units

The development approach for the spacecraft subsystems should include the use of Breadboard

Units, Engineering Units, Engineering Test Units, and Qualification Units to reduce the development risk associated with each subsystem depending on the maturity of the design.

These units can also be used to create a ground test bed to enable functional testing without the need for the Flight unit.

The spacecraft vendor will provide a spacecraft interface simulator of appropriate level of fidelity to enable effective use by mission elements for interface verification, including instruments.

2.4.6 Design Rules

The design of the spacecraft and its subsystems should meet the NASA/GSFC Rules for the Design, Development, Verification, and Operation of Flight Systems (GOLD Rules), GSFC-

STD-1000G.

2.4.7 Performance and Environmental Verification

The full system, after integration of the instrument payload (observatory), test program should include the following tests at a minimum:

• EMI/EMC

• Vibration

• Acoustics

• Shock, both launch vehicle shock environments and self-induced shocks

• Thermal Balance with three thermal cases (hot operational, cold operational, and cold survival)

• Thermal Vacuum testing, 4 thermal cycles

• Comprehensive performance testing (one before any environmental testing, one at hot plateau in TV, one at cold plateau in TV, and one after all environmental testing has been completed)

• Functional tests between all major tests, at the launch site after arrival, and then every two months and on the pad

• Alignment between the spacecraft master cube and the instruments before and after mechanical environments and after thermal vacuum testing

• Deployment testing of any mechanisms before and after mechanical environments

• RF compatibility (NEN and SN)

• Spacecraft or Observatory to Mission Operations Center (MOC) compatibility

• End-to-End testing - There will be two levels of end-to-end testing.

o One level will include all mission elements (e.g., instruments, spacecraft, communication, and ground system, MOC, and the Science Operations Center).

▪ This test will demonstrate that data collected by the instruments can be sent through all mission elements and be processed by the SOC.

▪ The preferred environment for this test is thermal vacuum.

▪ The test typically lasts approximately 2 days (not including setup time).

o The other level will include these mission elements: instruments (may not be included in every test), spacecraft, communication, and ground system, and MOC.

▪ There will be approximately 3 of these tests.

▪ This testing will occur at ambient.

▪ The testing typically lasts approximately 3 days each (not including setup time).

All tests should be in accordance with the General Environmental Verification Standards (GEVS), GSFC-STD-7000A, proto-flight test program. Considerations for possibly reducing tests given the quantities of systems are welcome.

2.5 Safety and Mission Assurance Processes

The AOS Program is interested in information on how the vendor would implement Safety and Mission Assurance, whether through a conventional approach with AOS Program-levied Safety and Mission Assurance requirements or through alternative approaches that still guarantee mission life requirements and satisfy mission Risk Class C per NPR 8705.4.

2.6 Access to Space (ATS) Insurance

Given the perceived risk of rideshare and hosted payload missions and given the dependence of mission success on multiple rideshare/hosted spacecraft, commercial insurance may help mitigate the mission’s risk. The insurance market for the commercial space industry is global with approximately 40 insurers worldwide including four markets in the US. Within the space insurance market there are many different types of coverage available addressing all aspects of satellite and vehicle manufacture, transportation, launch and in-orbit operation. Typical coverages commence at lift-off and extend for one year in orbit and cover the loss of a satellite during launch and early operation. At the end of this period insurance can be purchased to cover in-orbit technical issues. Pre-launch and third-party liability coverages are also available.

Insurance can also be limited to certain aspects of the satellite such as bus or platform only, payload only, launch only etc.

The process of purchasing insurance requires the involvement of spacecraft manufacturers, launch services providers, insurance brokers, underwriters, financial institutions, reinsurers, and AOS Program agents cooperating to coordinate an insurance for any given commercial satellite launch. This process provides the insurance company with a level of oversight and technical understanding of the underlying risks associated with the satellite and/or launch vehicle.

2.6.1 Insurance for the Traditional Spacecraft Development Approach (Option 1)

The perceived rideshare risk could be transferred to the spacecraft manufacturer by procuring the bus with a replacement guarantee in the event of failure. The on-orbit delivery of the spacecraft would occur following the successful completion of on orbit testing. If the spacecraft manufacturer is unable to deliver the spacecraft or meet agreed performance milestones due to loss or serious anomaly, then they would be required to build a replacement spacecraft. The replacement guarantee could be amended in the procurement contracts to become NASA credits in the event NASA elected not to have the mission repeated. The satellite manufacturer will be responsible for procuring insurance to cover the contractual replacement obligations. The cost of the insurance will increase the cost of the procurement to NASA. However, this cost impact will be significantly less than the cost of NASA replacing the spacecraft.

2.6.2 Insurance for the Hosted Payload Approach (Option 2 Specific)

The hosted payload (HP) risk can be transferred to the HP service provider by procuring the hosted payload service with a re-launch guarantee in the event of failure. Like Option 1, the on-orbit delivery of the HP would occur following the successful completion of on orbit testing. If the hosted payload provider is unable to deliver the payload to orbit or meet agreed performance milestones due to loss or serious anomaly, then they would be required to provide a replacement hosted payload opportunity. The replacement guarantee could be amended in the procurement contracts to become NASA credits in the event NASA elected not to have the mission repeated.

2.7 References

The following files are references for, and posted with, this RFP:

Doc # Title

GSFC-STD-1001A

NASA/GSFC Criteria for Flight and Flight Support System Lifecycle Reviews

GSFC-STD-1000G

NASA/GSFC Rules for the Design, Development, Verification, and Operation of Flight Systems (GOLD Rules)

GSFC-STD-7000A General Environmental Verification Standards (GEVS)

SP-20205003605 NASA Technology Readiness Assessment Best Practices Guide

NPR 7123.1C NASA Systems Engineering Processes and Requirements

NPR 8705.4 Risk Classification for NASA Payloads

Doc # Title

NASA-STD-8719.14 Process for Limiting Orbital Debris

GPR 7120.4D GSFC Risk Management

AOS-I-ATS-RFP-1 AOS Inclined ATS RFP Cover Letter

AOS-I-ATS-RFP-2 AOS Inclined ATS RFP Statement of Work

N/A Master Equipment List (MEL) Template w/ Instructions

3.0 INFORMATION REQUESTED

The information requested in this RFP along with expected deliverables and a complete Contract

Data Requirements List (CDRL) is contained in AOS-I-RFP-2, AOS Inclined ATS RFP Statement of Work.

Appendix A-1 - Spacecraft Capabilities vs. Instrument Assumptions (worksheet 1 of n)

Appendix A-2 - Spacecraft Capabilities vs. Instrument Assumptions (worksheet 2 of n)

S/C No.

Spacecraft Class /

Rideshare Adaptor

Bus Name /

Number

P/L Mass

Capability

P/L Power

(OAP/Peak)

Pointing

Control

(arcsec)

Pointing

Knowledge

(arcsec)

Pointing

Stability

(arcsec/sec)

Mounting

Area Avail.

Data

Storage

(Mbytes)

Data

Downlink

(Mbps)

Spacecraft

Design Life

Spacecraft

Dimensions

Delta V /

Prop

System

2 of n

No. Instrument Type

Instrument

Name P/L Mass

P/L Power

(OAP/Peak)

Pointing

Control

(arcsec)

Pointing

Knowledge

(arcsec)

Pointing

Stability

(arcsec/sec)

Mounting

Area Req.

Data

Storage

(Mbytes)

Data

Downlink

(Mbps)

Instrument

Design Life

Instrument

Dimensions Notes (#)

Total

Delta

Spacecraft Capabilities: Worksheet 2 of n

Instrument Requirements:

S/C No.

Spacecraft Class /

Rideshare Adaptor

Bus Name /

Number

P/L Mass

Capability

P/L Power

(OAP/Peak)

Pointing

Control

(arcsec)

Pointing

Knowledge

(arcsec)

Pointing

Stability

(arcsec/sec)

Mounting

Area Avail.

Data

Storage

(Mbytes)

Data

Downlink

(Mbps)

Spacecraft

Design Life

Spacecraft

Dimensions

Delta V /

Prop

System

1 of n

No. Instrument Type

Instrument

Name P/L Mass

P/L Power

(OAP/Peak)

Pointing

Control

(arcsec)

Pointing

Knowledge

(arcsec)

Pointing

Stability

(arcsec/sec)

Mounting

Area Req.

Data

Storage

(Mbytes)

Data

Downlink

(Mbps)

Instrument

Design Life

Instrument

Dimensions Notes (#) m

Total

Delta

Spacecraft Capabilities: Worksheet 1 of n

Instrument Requirements:

Appendix B – AOS RFP Instrument Assumptions Summary

For spaceacecraft sizing use MEVs where available.

Appendix B.1 – Notional Radar Instrument Assumptions

Assumption Type Value

Dimensions

L x W x H

Electronics Box A: 30 cm x 30 cm x 30 cm

Electronics Box B: 30 cm x 15 cm x 15 cm

Solid Antenna: Diameter: 210 cm, Height: 70 cm

Deployed Mesh Antenna: Diameter: 210 cm, Height including fixed feed horn: 180 cm

External Radiator Area Requires radiator area of approximately 1.5 m^2 near solid antenna and electronics boxes for heat pipe transfer. For ATS RFP can assume radiator tailored to specific spacecraft and will be GFE.

Mass Properties 120 kg CBE; 190 kg MEV

Box A: 30 kg CBE; 40 kg MEV

Box B: 10 kg CBE; 15 kg MEV Solid Dish: 50 kg; 75 kg MEV

Mesh Dish: 30 kg; 60 kg MEV

Data latency < 1 hour

Data interface(s) 4 Mbps CBE; 5 Mbps MEV

Timing 1 pps signal

Pointing

Accuracy/Knowledge

Changing off-nadir pointing to match cross-track location of lidar beam

(other spacecraft) to within 0.025 degrees (90 arcsec), maintaining 2 degree off-nadir along-track pointing angle. Co-aligned with Lidar-09R

(scene registration, not temporal)

Pointing Stability 0.01-degree (36 arcsec) RMS 3 sigma over 2 seconds

Jitter 0.01-degree (36 arc sec) RMS 3 sigma over 0.2 seconds

Horizontal Spatial

Resolution

5 km Ku-band, 1 Km W-band

Field of Regard No mechanical scanning, +/- 6 degrees cross-track and +/- 6 degrees along-track.

Field of View Near-nadir, single beam, 0.1 degree (W-band, fixed antenna), and 0.7 degree (Ku-band, both antennas).

Viewing Restrictions W-band beam nominally points 2-5 degrees forward 0 cross-track; Solid and Mesh dishes aligned along S/C velocity vector

Operational Temp -10 to +45 deg C

Survival Temp TBD (N/A for ATS RFP Study)

Thermal Stability N/A

Thermal Gradients N/A

Power 230 W CBE; 336 W MEV Box A 220 W CBE; 316 W MEV

Box B 10 W CBE; 20 W MEV

Appendix B.2 – Notional Lidar Instrument Draft Assumptions

L x W x H

Small Dedicated Satellite Version (only use of dedicated S/C):

99 cm x 92 cm x 88 cm (radiator on 3 sides requires S/C to rotate once per orbit about nadir axis)

Shared Payload Version (more common):

Envelope: 148 cm x 99 cm x 102 cm Radiator Side: 148 cm x 102 cm

Mass Properties 125 kg CBE, 152 kg MEV

Data latency <6 hours (<2 hours preferred)

Data interface(s) 2.87 Mbps CBE, 12.0 Mbps MEV

Timing 1 PPS with 0.005 ms accuracy

Pointing < 170 arcsec

Pointing Stability < 25 arcsec

Jitter Jitter < 20 arcsec max, < 0.33 arcsec/second

Field of Regard 50 degrees (keep out zone)

Field of View 0.0066 degrees half angle

Viewing Restrictions Nominal pointing 2-5 degrees off-Nadir, pointed with W-band of radar; sunshield around Ø60cm receiver; Do not point telescope within 60 degrees of the sun

Operational Temp -10 to +30 C

Survival Temp -20 to +60 C

Thermal Stability N/A

Thermal Gradients N/A

Power Peak: 309 W CBE; 336 W MEV Orbit Avg. 296 W CBE

Standby: 140 W standby Survival: 86 W CBE; 104 W MEV (survival heater)

Appendix B.3 – Notional Polarimeter Instrument Draft Assumptions

L x W x H

Envelope: 60 cm x 40 cm x 45 cm

Radiator Side: 60 cm x 45 cm Mass Properties 35 kg CBE; 45 kg MEV

Data latency 3-6 hours

Data interface(s) 100 Mbps peak MEV (daylight only), 0 Mbps during eclipse Orbital Average: Approx. 50 Mbps

Timing 1 PPS (1 ms accuracy) via Spacewire

Pointing 43 arc-sec (0.012 degrees) Accuracy; 18 arc sec (0.005 degrees)

Knowledge - all 3 axes

Pointing Stability 10 arc-sec (3 sigma)

Jitter 226.8 arcsec (0.063 degrees) /sec jitter

Field of Regard 144 degrees

Field of View Along track 114 degrees; cross track 94 degrees

Viewing Restrictions Nadir at center of FOR; Length aligned along S/C velocity vector

Operational Temp -20 to +35 C

Survival Temp -40 to +70 C

Thermal Stability N/A

Thermal Gradients N/A

Power 50 W CBE; 65 W MEV; survival power 20 W

Appendix B.4 – Notional Radiometer Instrument Draft Assumptions

L x W x H

Envelope: 80 cm x 45 cm x 45 cm

Radiator Side: 80 cm x 45 cm Mass Properties 40 kg CBE; 52 kg MEV

Note: Contains spinning mirror (inertia 0.007 Kg*m^2) that rotates up to 150 RPM with peak angular moment CBE of 0.15 N*m*s

Data latency <1 hour

Data interface(s) 0.2 Mbps CBE; 0.26 Mbps MEV

Timing TBD – future work

Pointing Nadir, (cross-track scanning), 0.1-degree accuracy (360 arc sec) Knowledge, 0.05 degree (180 arc sec)

Pointing Stability TBD – future work

Jitter 144 arcsec within 1 ms

Field of Regard Nadir (along-track), +/- 43 degree cross-track., 72-88 degree off-nadir cross-track (cold sky calibration – instrument has a scan mirror & view port to accommodate – radiator is on this face)

Field of View TBD – future work

Viewing Restrictions Cold view for radiator; Length aligned along S/C velocity vector

Operational Temp -10 to +45 deg C

Survival Temp TBD (N/A for ATS RFP Study)

Thermal Stability N/A

Power 55 W CBE; 70 W MEV; 45 W survival heater power

Appendix B.5 – Notional Tandem Stereographic Camera Draft Assumptions

L x W x H

Envelope: 75 cm x 20 cm x 40 cm

Radiator Area: 75 cm x 40 cm Note: Contains 3 cameras (Nadar and +/- 38 degrees)

Mass Properties 18 kg CBE; 21 kg MEV

Data latency < 6 hours

Data interface(s) 16 Mbps CBE; 20 Mbps MEV (daylight only); 0 Mbps during eclipse

Note: Data rate for two (2) cameras with third camera inactive;

Orbital average approx. 10 Mbps MEV.

Timing 1 ms accuracy

Pointing 0.25 degrees accuracy (900 arc sec)

Pointing Stability 100 arcsec (3 sigma) over 30 sec

Jitter 5 arcsec over 0.5 sec RSS of all three axes

Field of Regard Cross track 15 degrees;

Along track 15 degrees

Field of View 15 deg full angle per camera head cross track; 12.5 degrees along track

Viewing Restrictions One camera at nadir; second camera head: 38 deg aft, third camera:

38 deg forward; Length aligned along S/C velocity vector

Operational Temp -20 to +55 C

Survival Temp -30 to +85 C

Thermal Stability N/A

Power 19 W CBE; 23 W MEV; 4 W survival heater power

Appendix C – A Summary of the Goddard Proof of Concept Design Study

Appendix C.1 – Proof of Concept: Mission and Observatory Assumptions

Owner Assumption

Mission LRD July 25, 2028 (ATP ~ now)

Mission 2-year mission life, 5 years consumables

Mission Mission Class C

Mission 407 km orbit (+/- 10 km), 65-degree inclination

Mission S/C-1 & S/C-2 separated by 45 seconds (+/- 15s), S/C-1 trails.

Mission Constellation alignment 3km cross-track

Mission

Cameras co-registered temporally between the 30s and 60s spacing (simultaneous overlap); spatial co-registration of radar and lidar footprints within 100m cross-track.

Three cameras used on the S/C to account for yaw flips.

Mission End of Mission disposal required

Mission ELV-Class LV

Mission The L/V accommodates two contributed s/c: one ESPA Grande S/C and one s/c on top the Integrated P/L stack (IPS).

Owner Assumption

Observatory 3 axes stabilized (+Z points to Nadir, +Y cold, +/- X RAM depending on yaw flips)

Observatory Mission Phase based Modes (See Modes Table; MEL)

Observatory On-board propulsion

Observatory

Pointing:

43 arc-sec (0.012 degrees) Accuracy; 18 arc sec (0.005 degrees) Knowledge - all 3 axes:

Stability: 20 arcsec (3 sigma) all 3 axes

Jitter: 10 arcsec/10 msec all 3 axes

Observatory Absolute timing 1 PPS with 0.005 ms accuracy

Observatory Accommodate 54 Mpbs CBE science data from SSG-2 and 6.8 Mbps CBE science data from SSG-1

Observatory 1-hour data latency

Observatory 28V service to payloads

Observatory 72 hours onboard storage

Observatory Accommodate payload door and other deployments

Appendix C.2 – Proof of Concept: Ground Systems and Operations Assumptions

Owner Assumption

Ground

Systems

Provide support for 2 observatories simultaneously

Ground Systems

Receive housekeeping & science data telemetry

Ground Systems

Record/Archive science data

Ground Systems

Provide critical event telecom coverage: Launch Sep, S/A Deployment, Instrument Deployments

Owner Assumption

Operations Provide an operations environment compliant appropriate mission class (Class C)

Operations Provide an IT Security plan for the MOC

Operations Provide an FOT Staffing Profile supporting Operations Concept

Operations Develop Ground Data Systems Architecture

Operations Develop Communication links that support MOC Architecture

Operations Provide Operations & other required tests & simulations

Operations Provide any required Instrument Calibration support

Operations Develop Activity Scheduling protocols

Operations MOC/SOC acquire Level 0 data, process to Level 1 and Level 2.

Perform science data trending.

Appendix C.3 – Proof of Concept: Instruments on a Notional S/C in a Notional L/V

Appendix C.4 – Proof of Concept: S/C-1 Instrument Layout

Appendix C.5 – Proof of Concept: S/C-2 Instrument Layout

Appendix D: AOS Inclined RFP Target Value Summary

Category Parameter

RFP

Section(s) ATS RFP Target Values

Instrument Suite

Radar (1x) 2.2.1; B.1 MEV: 190 kg, 336W, 5 Mbps

Tandem Stereographic Camera #1 (On S/C #1)

2.2.1; B.5 MEV: 21 kg, 23 W, 16 Mbps Daylight Only; Approx. 8 Mbps Orbit Average

Tandem Stereographic Camera #2 (On S/C #2)

2.2.1; B.5 MEV: 21 kg, 23 W, 16 Mbps Daylight Only; Approx. 8 Mbps Orbit Average

Backscatter Lidar (1x) 2.2.1; B.2 MEV: 152 kg, 336 W, 12 Mbps

Radiometer (1x) 2.2.1; B.4 MEV: 52 kg, 65 W, 0.26 Mbps

Polarimeter (1x) 2.2.1; B.3 MEV: 45 kg, 65 W, 100 Mbps Peak Daylight Only, Approx. 50 Mbps Orbit Average

Programmatic

Risk Class 2.1.4 Class C per NPR 8705.4

Response Option Selected 1.0 Option 1 or Option 2

Number of Observatories 2.2 2 (minimum) - 3 (maximum)

Bus Type 1.0 Any

Risk Tolerances / Reliability 2.1.4 Typical for Class C

Access to Space Insurance 2.6.1, 2.6.2 Desired for Option 1 & 2

Room for Large Ride Share SC on Top Launch Stack

2.2, C.1 Evaluate ability for architecture to accommodate ride share SC on top launch stack

ESPA Ride Share Capable 2.2.2 Desired

Launch Year 2.1.3 July 2028

Environmental Requirements 2.4.7 Test program in accordance to GSFC-

STD-7000A (GEVS)

Design & Analysis (Mechanical, Thermal, Electrical)

2.4.6 Provide per GSFC GOLD Rules:

GSFC-STD-1000G

Instrument Operating Temperature 2.2.5 Temp: -10 to +40°C

Instrument Survival Temperature 2.2.5 Temp: -20 to +50°C

Cold Sky Field of View 2.1.7, 2.2.5 Made available to instrument radiators to maintain temperatures.

Section(s) ATS RFP Target Values

Spacecraft Launch Dimensions 2.1.5 All SC able to fit in 1x ELV Class LV along with identified rideshare SC.

Propulsion and ACS Actuators

Mission Life 2.1.3 3 years, 5 years consumables + disposal

Delta-V Capability (MEV Mass) 2.1.7 Enough to satisfy 5-year life

Orbital Debris & Controlled Entry 2.1.6 Controlled Reentry per

NASA-STD-8719.14

Instrument Voltage 2.2.7 28 V

Operational Modes 2.1.2 Science, Safe, Load Shed, Eng., Etc.

Powered at Launch 2.1.2 Yes

Communication and Data

Comm Type 2.1.10 Terrestrial & Space

Data Downlink 2.1.10 Terrestrial & Space

Command Encryption 2.1.12 Yes

Data Latency 2.2.2, B. 1 hour preferred else 3-6 hours

Communication Type 2.2.7 RS-422 or Similar

Data Storage 2.1.11 Enough for 72 hours no eclipse (peak instrument data collection rates)

Timing and Position 2.1.9 Position: Use GPS Timing: GPS UTC Registration <100 msec

Pointing Control Accuracy 2.2.3; B 43 arcsec (3σ) / 3-axis for polarimeter SC, see Appendix B for other instruments

Pointing Stability 2.2.3; B 10 arcsec (3σ) / 3-axis for polarimeter SC, see Appendix B for other instruments

Pointing Knowledge 2.2.3; B 18 arcsec (3σ) / 3-axis for polarimeter SC, see Appendix B for other instruments

Alignment Stability 2.2.4 20 arcsec (3σ) / 3-axis

Jitter 2.2.3; B 10 arcsec (10 ms) / 3-axis for stereographic camera SC, see Appendix B for other instruments

Science Availability 2.1.8 >96%

Flight Dynamics

Momentum Management Maneuver Frequency

2.1.7, 2.1.8 Specify as applicable

Orbit Parameters 2.1.1, 2.2.2 407 +/- 10 km SMA; 40-55° inclination

Constellation Alignment 2.2.2 <3 km Cross-Track (Radar & Lidar SC) Else <10 km Cross-Track

Constellation Co-Registration 2.2.2 ≤200m cross-track (Radar/Lidar)

Section(s) ATS RFP Target Values

Constellation Spacing (Stereographic Camera SC)

2.2.2 45 +15 / -10 Seconds

Const. Lidar / Radar Spacing (If no Stereographic Camera SC #3)

2.2.2 < 60 seconds

Yaw Flips Required 2.2.5 Yes, and specify number and timing

De-Orbit 2.1.6 Required

Safety & Mission Assurance

SMA Approach 2.5 Specify SMA approach

Integration & Test

Contamination Control 2.2.6 Class 10k when optical instruments integrated

Development Approach 2.4 Specify approach and note any differences from what’s listed in subsections 2.4.1 through 2.4.7

Mission Operations

Mission Operations Center 2.1.10 Address MOC ConOps

Science Operations Center 2.1.10 Address SOC ConOps

Instrument Operations Center 2.1.10 Address IOC ConOps