ENCLOSURE_1_-_Project_Scenario_-_ARM.pdf

ENCLOSURE 1

PROJECT

SCENARIO: ARM

RFP # NNG16544506R

Enclosure 1

NNG16544506R

Table of Contents

Contents

1 INTRODUCTION

1.1 Sample Problem Overview

1.2 Sample Problem Scope

1.3 Mission Description – ARM

2 SCENARIO TECHINCAL REQUIREMENTS

2.1 ARM Mission Assurance Scoping and Planning

2.2 ARM IV&V Rescope

2.3 Application of Efficiency

3 ACRONYM LIST

1 INTRODUCTION

1.1 Sample Problem Overview

This SAS Scenario for the NASA Asteroid Redirect Mission (ARM) is defined to provide the Offeror an opportunity to illustrate a wide range of experience, skills, and efficient management and development processes for providing assurance services. Additionally, Solar Array System (SAS) bidders can demonstrate an understanding of the complexities of working the many aspects of a mission’s lifecycle within NASA Centers and changes in mission support requirements and schedules that often occur. This Scenario is unique in that it covers two distinct NASA Mission Directorates: Science and Human Exploration.

This Scenario depicts a fictitious ARM mission. The mission was developed to address needs of the SAS procurement activity and is based off the concept of an actual NASA mission. However, no data beyond this document is required or expected to be referenced in regards to mission details.

1.2 Sample Problem Scope

For this Scenario, the Mission Description is the target mission that will receive IV&V assurance services support. Assume that the mission development timeline is 5 years. Assume that the Mission is past Mission Definition Review but prior to System Requirements Review, as defined in NASA Procedural Requirements (NPR) 7120.5E.

1.3 Mission Description – ARM

Objectives Rendezvous with, capture, and redirect an entire near-Earth steroid (NEA), with a mass of up to ~1,000 t, into a stable lunar orbit by the first half of the next decade.

• Demonstrate high--‐power Solar Electric Propulsion (SEP).

• Demonstrate the ability to capture large non--‐cooperative objects in deep space.

• Stimulate the detection and characterization of small NEAs.

• Use assets already under development (SLS and Orion) as a first step to using these new capabilities even deeper in space.

Mission Overview

• Utilizes a 40 kW solar electric propulsion system to rendezvous with a NEA and redirect it to translunar space in the 2021 timeframe.

• Utilizes 50--‐kW--‐class solar arrays current under development by the Space Technology Mission Directorate’s (STMD) Solar Array System program.

• Is capable of launching in 2017 or 2018 on an Atlas V 551, Falcon Heavy, or the Space Launch System

(SLS).

• The storage orbit for the redirected asteroid is a stable distant retrograde orbit (DRO) in the Earth--‐ Moon system with an orbit altitude of ~70,000 km above the lunar surface.

• The DRO is reachable by SLS/Orion with two crew members in a 22 to 25 day mission.

Asteroid rendezvous, capture, and redirect operations build on past mission experience

Asteroid Redirect Vehicle (ARV) The ARV is composed of two modules to enable parallel development, assembly, and testing: a Solar Electric Propulsion (SEP) Module, and a Mission Module. The Mission Module is comprised of an Avionics Module, the sensor suite and the capture mechanism. The SEP Module includes all of the power and propulsion for the ARV. The Avionics Module includes all other spacecraft bus functions.

The SEP Module The asteroid redirect mission is enabled by high--‐power solar electric propulsion. Solar electric propulsion missions always show better performance at higher power levels and this mission is no different. Mission design trade studies indicate that the best combination of asteroid mass and flight times are obtained at the highest power levels. The study, therefore, selected the highest solar array power level that could reasonably be available for launch in this decade. That corresponds to about 50-kW, which is the upper end of the 30--‐kW to 50--‐kW range of power levels currently under development in two Solar Array System (SAS) development activities sponsored by NASA’s Space Technology Mission Directorate (STMD). A 50‐kW solar array beginning-of-life at 1 AU would enable operation of a 40-kW electric propulsion system at end‐of-life with appropriate margins. The 40-kW input power would be processed by multiple, magnetically shielded Hall thrusters operating in parallel with a specific impulse of 3,000 s. Because the asteroid redirect mission is enabled by these technologies it is an ideal platform to meet the needs of STMD’s SEP Technology Demonstration Mission. The asteroid redirect mission would demonstrate deployment and operation of a new class of large lightweight, high-specific-power, flexible‐blanket solar arrays in space along with the operation of a high-power, high-performance electric propulsion system.

The ARV configuration is dominated by the need to be able to store up to 12 metric tons of xenon. This is significantly greater than the 0.43 t of xenon launched on the Dawn mission, which is the largest xenon propellant load launched to date. Commercial communication satellite manufacturers typically launch only a few hundred kilograms of xenon used for orbit raising and station keeping maneuvers.

Consequently, there are no existing tanks that meet the needs of the asteroid redirect mission, so a new tank development is required. The study team identified a solution that minimizes the development risk and cost and provides the lightest tank mass. This approach uses a composite overwrapped pressure vessel with a seamless aluminum liner design made using existing industry manufacturing techniques.

Capture Mechanism Multiple options for the capture mechanism were considered and evaluated. A non--‐ rigidized, inflatable capture bag approach was selected for this study to assess feasibility. The capture bag approach effectively deals with the range of asteroid mechanical property uncertainties and would work equally well if the asteroid was a rubble pile or solid rock. Importantly, this concept is testable in a 1--‐g environment enabling verification and validation of the system before launch. Other concepts were received from industry that may also be feasible, but these have not yet been examined in detail.

The capture process itself is dominated by the spin state of the target. At relatively slow spin-rates of <

0.2 rpm (periods > 5 minutes) about one or more axes, the problem is relatively straightforward and small forces, < 0.1 g, are transmitted back to the spacecraft. At spin rates an order of magnitude higher, up to about 2 rpm (spin period of 30 seconds), the problem is more challenging. For this case, the study team identified a feasible approach in which the capture bag is closed tightly around the asteroid over a few minute period. Residual cross--‐axis spin is managed by force‐controlled winches to keep accelerations reflected back to the spacecraft to less than 0.1 g while the RCS thrusters are used to de‐tumble and then despin the spacecraft/asteroid combination.

Avionics Module To reduce the cost and schedule for the flight system implementation, flight--‐ qualified, deep--‐space avionics and sensors were identified that would work well for the asteroid redirect mission. This includes avionics and core flight software from the Soil Moisture Active/Passive (SMAP) project, the Mars Science Laboratory, and a sensor suite that could include instruments derived from the Origins-Spectral Interpretation-Resource Identification-Security-Regolith Explore (OSIRIS--‐REx) mission.

Sensor Suite The sensor suite supports optical navigation, asteroid characterization, and asteroid capture. The following sensors are currently in the ARM reference design:

• Wide‐field camera – used for long range optical navigation

• Medium--‐field camera – used for close range mapping of the asteroid

• Visible and IR Spectrometer – used for chemical characterization of the asteroid

• Laser Altimeter – used for asteroid characterization and mapping and for final autonomous control of the capture event

Launch System The study assessed the feasibility and mission performance of three different launch vehicles: Atlas V 551, Falcon Heavy, and the Space Launch System (SLS). The mission is feasible with all three launch vehicles. The lower performance of the flight--‐proven Atlas V can be compensated for with the SEP system by spiraling out from an Earth elliptical orbit and using a Lunar Gravity Assist (LGA) to escape from Earth. The spiral out adds a year to a year and a half to the flight time required to get to the target NEA. The higher performance capabilities of the Falcon Heavy or SLS launch vehicles enable elimination of the Earth--‐spiral phase and would launch the ARV directly to the LGA. This simplifies the flight system design, shortens the overall mission duration, and enables later launch dates.

Safe Lunar Storage Orbit Retrieval of the asteroid is accomplished by using the SEP system to make small adjustments to the asteroid’s orbit in order to target a lunar gravity assist maneuver enabling capture into the Earth--‐Moon system. The 40--‐kW SEP system provides a few hundred meters per second velocity change to the 1000--‐t asteroid over the three to five year return trip. This redirection is only possible for asteroids that naturally return close to the Earth--‐Moon system in the timeframe of interest. The lunar gravity assist provides a velocity change of ~1,600 m/s, which is sufficient to capture the asteroid into the Earth--‐Moon system. An additional ~60 m/s is provided by the SEP system to transfer the asteroid to the long--‐term--‐ stable, lunar distant retrograde storage orbit. The asteroid size and composition would be selected such that it could not survive an Earth atmospheric entry. The long--‐term stable orbit is targeted to make the asteroid available as a destination for multiple future missions.

2 SCENARIO TECHINCAL REQUIREMENTS

The following sections identify the separate/individual technical scenarios that the offeror shall need to address. Note that the offeror is not to perform any actual mission assurance analysis work in response to the RFP.

2.1 ARM Mission Assurance Scoping and Planning

In this sample problem, the Offeror is presented with the new Mission (ARM) for an upcoming Government Fiscal Year.

The Offeror is responsible for scoping the effort required to provide Mission Assurance services to the ARM. This scenario cites scoping by element for simplicity; however, scoping may be performed by element, behavior, risks to mission success, critical events, or any other discriminator deemed appropriate by the Offeror.

The Offeror shall provide the following responses in their proposal:

1. Identify the critical system and software elements of the ARM and develop a prioritization of those critical elements.

2. Specify the system and software elements requiring Missions Assurance based on their criticality.

3. Assess and identify the required analysis tasks, and resulting IV&V products, to provide the necessary Mission Assurance and reduce risk to Mission Success. It should be clear how the proposed tasks directly contribute to reducing risk to Mission Success and providing the necessary Mission Assurance.

2.2 ARM IV&V Rescope

Occasionally, during the course of an IV&V Project, re-scoping of the work may be warranted due to events outside of the IV&V Project’s control. In this sample problem, the offeror is asked to re-scope the ARM Mission Assurance support based on each of the following events separately.

The Offeror shall provide the following responses in their proposal:

1. Determine and identify what risks and impacts would be experienced by the IV&V project in the event that the ARM Preliminary Design Review (PDR) is slipped approximately one year. Summarize how the IV&V plan will be adjusted to respond to the delay.

2. Through your Mission Assurance activities and information received from the Mission Project, you have determined that an in-scope critical element will require more assurance than originally planned. Identify how you will adjust your plan to provide the necessary assurance for the element.

3 For each of these events, the Offeror shall provide a summary of what steps will be taken to address the changes and how the existing IV&V plan will be affected. The Offeror shall also provide information regarding any coordination with other offices and/or supporting functions within the IV&V Program if applicable.

2.3 Application of Efficiency

In this sample problem, the offeror is asked to address a situation in which the planned Mission Assurance has been performed to a high degree of efficiency, leading to a projected cost underrun (surplus) of approximately $100,000.

The Offeror shall provide the following responses in their proposal:

1. Provide a recommendation as how to best address the budget surplus. The Offeror shall summarize what actions would be taken to address the surplus and how existing plans will be affected.

2. Provide information regarding any coordination with other offices and/or supporting functions within the IV&V Program if applicable.

3 ACRONYM LIST

ARCM Asteroid Redirect Crewed Mission ARM Asteroid Redirect Mission ARV Asteroid Redirect Vehicle AU Astronomical Unit DRO Distant Retrograde Orbit IV&V Independent Verification & Validation LGA Lunar Gravity Assist NEA Near Earth Asteroid NPR NASA Procedural Requirements OSIRIS-REx Origins, Spectral Interpretation, Resource Identification, Security, Regolith Explorer PDR Preliminary Design Review SAS Solar Array System SEP Solar Electric Propulsion SLS Space Launch System SMAP Soil Moisture Active/Passive STMD Space Technology Mission Directorate