Carolyn R. Mercer 

Glenn Research Center, Cleveland, Ohio

Scott D. Vangen

Kennedy Space Center, Kennedy Space Center, Florida

Julie A. Williams-Byrd

Langley Research Center, Hampton, Virginia

Jonette M. Stecklein, Shamim A. Rahman, Matthew E. Rosenthal, and David M. Hornyak

Johnson Space Center, Houston, Texas

Leslie Alexander

Marshall Space Flight Center, Huntsville, Alabama

David J. Korsmeyer, Eugene L. Tu, and David D. Alfano

Ames Research Center, Moffett Field, California

Craig E. Kundrot

Johnson Space Center, Houston, Texas

Dianne S. Wiley, Stephen C. Davison, and Tibor S. Balint

National Aeronautics and Space Administration, Washington, D.C. 

Critical Technology Determination for 

Future Human Space Flight

NASA/TM—2012-217670

September 2012

GLEX-2012.09.3.3x12551
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 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

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Carolyn R. Mercer 

Glenn Research Center, Cleveland, Ohio

Scott D. Vangen

Kennedy Space Center, Kennedy Space Center, Florida

Julie A. Williams-Byrd

Langley Research Center, Hampton, Virginia

Jonette M. Stecklein, Shamim A. Rahman, Matthew E. Rosenthal, and David M. Hornyak

Johnson Space Center, Houston, Texas

Leslie Alexander

Marshall Space Flight Center, Huntsville, Alabama

David J. Korsmeyer, Eugene L. Tu, and David D. Alfano

Ames Research Center, Moffett Field, California

Craig E. Kundrot

Johnson Space Center, Houston, Texas

Dianne S. Wiley, Stephen C. Davison, and Tibor S. Balint

National Aeronautics and Space Administration, Washington, D.C. 

Critical Technology Determination for 

Future Human Space Flight

NASA/TM—2012-217670

September 2012

GLEX-2012.09.3.3x12551

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Global Space Exploration Conference

sponsored by the IAF and AIAA

Washington, D.C., May 22–24, 2012
Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217670 iii

Contents

Abstract.......................................................................................................................................................................... 1

1.0 Introduction .......................................................................................................................................................... 1

2.0 Architectural Elements and Destinations............................................................................................................. 2

3.0 Technology Selection Method.............................................................................................................................. 3

4.0 Technologies......................................................................................................................................................... 4

5.0 Discussion............................................................................................................................................................. 8

6.0 Summary............................................................................................................................................................... 9

Appendix A.—Acronyms............................................................................................................................................11

Appendix B.—Technology Description Format..........................................................................................................13

Appendix C.—Technology Descriptions.....................................................................................................................15

C.1 TA 01—Launch Propulsion .......................................................................................................................15

C.1.1 Oxygen-Rich Staged Combustion (ORSC) Engine Technology...................................................15

C.1.2 Advanced, Low Cost Engine Technology for SLS .......................................................................15

C.2 TA 02—In-Space Propulsion.....................................................................................................................15

C.2.1 LO2/LCH4 Cryogenic Propulsion System.....................................................................................15

C.2.2 LO2/LCH4 Reaction Control Engines............................................................................................15

C.2.3 Electric Propulsion and Power Processing....................................................................................16

C.2.4 Nuclear Thermal Propulsion (NTP) Engine..................................................................................16

C.2.5 Unsettled Cryogenic Propellant Transfer......................................................................................16

C.2.6 In Space Cryogenic Liquid Acquisition ........................................................................................17

C.3 TA 03—Space Power and Energy Storage ................................................................................................17

C.3.1 300 kWe Fission Power for Electric Propulsion ...........................................................................17

C.3.2 High Strength/Stiffness Deployable 10 to 100 kW Class Solar Arrays ........................................17

C.3.3 Autonomously Deployable 300 kW In-Space Solar Arrays..........................................................18

C.3.4 Fission Power for Surface Missions..............................................................................................18

C.3.5 Multi-MWe Nuclear Power System for Electric Propulsion.........................................................18

C.3.6 Regenerative Fuel Cells ................................................................................................................18

C.3.7 High Specific Energy Batteries.....................................................................................................18

C.3.8 Long Life Batteries .......................................................................................................................19

C.4 TA 04—Robotics, Tele-Robotics, and Autonomous Systems ...................................................................19

C.4.1 Precision Landing and Hazard Avoidance ....................................................................................19

C.4.2 Telerobotic Control of Robotic Systems with Time Delay ...........................................................19

C.4.3 Autonomous Vehicle Systems Management.................................................................................19

C.4.4 Common Avionics ........................................................................................................................20

C.4.5 Automated/Autonomous Rendezvous and Docking, Proximity Operations and Target 

Relative Navigation.......................................................................................................................20

C.4.6 Crew Autonomy Beyond LEO......................................................................................................20

C.4.7 Robots Working Side-by-Side with Suited Crew..........................................................................20

C.5 TA 05—Communication and Navigation ..................................................................................................21

C.5.1 High Data Rate Forward Link (Flight) Communications.............................................................21

C.5.2 High-Rate, Adaptive, Internetworked Proximity Communications..............................................21

C.5.3 In-Space Timing and Navigation for Autonomy...........................................................................21

C.5.4 Quad Function Hybrid RF/Optical Comm, Optical Ranging, RF Imaging System ......................21

C.6 TA 06—Human Health, Life Support, and Habitation Systems................................................................22

C.6.1 Closed-Loop, High Reliability, Life Support Systems................................................................. 22

C.6.2 Deep Space Suit (Block 1)............................................................................................................23

C.6.3 Lunar Surface Space Suit (Block 2)..............................................................................................23

C.6.4 Mars Surface Space Suit (Block 3) ...............................................................................................23

C.6.5 Long Duration Spaceflight Medical Care .....................................................................................23

C.6.6 Long-Duration Spaceflight Behavioral Health and Performance..................................................24

C.6.7 Microgravity Biomedical Counter-Measures for Long Duration Spaceflight...............................24

C.6.8 Microgravity Biomedical Counter-Measures—Optimized Exercise Equipment ..........................24

C.6.9 Deep Space Mission Human Factors and Habitability..................................................................25

C.6.10 In-Flight Environmental Monitoring.............................................................................................25
NASA/TM—2012-217670 iv

C.6.11 Fire Prevention, Detection and Suppression (Reduced Pressure) .................................................25

C.6.12 Space Radiation Protection—Galactic Cosmic Rays (GCR) ........................................................26

C.6.13 Space Radiation Protection—Solar Particle Events (SPE) ...........................................................26

C.6.14 Space Radiation Shielding—SPE..................................................................................................26

C.7 TA 07—Human Exploration Destination Systems ...................................................................................26

C.7.1 Lunar ISRU: Oxygen/Water Extraction From Regolith................................................................26

C.7.2 Mars ISRU: Oxygen from Atmosphere and Water Extraction from Soil .....................................27

C.7.3 Anchoring Techniques and EVA Tools for Micro-G Surface Operations....................................27

C.7.4 Suit Port ........................................................................................................................................27

C.7.5 Surface Mobility ...........................................................................................................................28

C.7.6 Mission Control Automation Beyond LEO...................................................................................28

C.7.7 Dust Mitigation .............................................................................................................................28

C.8 TA 08—Science Instruments, Observatories and Sensors Systems...........................................................28

C.9 TA 09—Entry, Descent, and Landing Systems..........................................................................................28

C.9.1 Entry, Descent and Landing (EDL) Technologies—Mars Exploration Class Missions ...............29

C.9.2 EDL Technologies—Earth Return ................................................................................................29

C.10 TA 10—Nanotechnology ...........................................................................................................................29

C.11 TA 11—Modeling, Simulation, Information Technology and Processing.................................................29

C.11.1 Advanced Software Development/Tools ......................................................................................29

C.12 TA 12—Materials, Structures, Mechanical Systems and Manufacturing ..................................................30

C.12.1 Structures and Materials for Inflatable Modules...........................................................................30

C.12.2 Lightweight and Efficient Structures and Materials.....................................................................30

C.12.3 Mechanisms for Long Duration, Deep Space Missions................................................................30

C.12.4 Low Temperature Mechanisms.....................................................................................................31

C.13 TA 13—Ground and Launch Systems Processing .....................................................................................31

C.13.1 Low Loss Cryogenic Ground Systems Storage and Transfer........................................................31

C.13.2 Corrosion Detection and Control for Ground Systems ................................................................31

C.13.3 Fault Detection, Isolation, and Recovery for Ground Systems.....................................................32

C.13.4 Wiring Fault Detection and Repair for Ground Systems ..............................................................32

C.14 TA 14—Thermal Management Systems....................................................................................................32

C.14.1 In-Space Cryogenic Propellant Storage—Zero Boil Off (ZBO) LO2 and Reduced/Zero 

Boil Off (RBO) LH2......................................................................................................................32

C.14.2 Thermal Control............................................................................................................................33

C.14.3 Robust Ablative Heat Shield (Beyond Lunar Return Conditions)—Thermal Protection 

System...........................................................................................................................................33

C.14.4 Robust Ablative Heat Shield (Lunar Return Conditions)—Thermal Protection System .............33

References ...................................................................................................................................................................34
NASA/TM—2012-217670 1 

Critical Technology Determination for Future Human Space Flight

Carolyn R. Mercer 

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Scott D. Vangen

National Aeronautics and Space Administration

Kennedy Space Center

Kennedy Space Center, Florida 32899

Julie A. Williams-Byrd

National Aeronautics and Space Administration

Langley Research Center

Hampton, Virginia 23681–2199

Jonette M. Stecklein, Shamim A. Rahman, Matthew E. Rosenthal, and David M. Hornyak

National Aeronautics and Space Administration

Johnson Space Center

Houston, Texas 77058

Leslie Alexander

National Aeronautics and Space Administration

Marshall Space Flight Center

Huntsville, Alabama 35812

David J. Korsmeyer, Eugene L. Tu, and David D. Alfano

National Aeronautics and Space Administration

Ames Research Center

Moffett Field, California 94035

Craig E. Kundrot

National Aeronautics and Space Administration

Johnson Space Center

Houston, Texas 77058

Dianne S. Wiley,* Stephen C. Davison, and Tibor S. Balint

National Aeronautics and Space Administration

Washington, D.C. 20546

Abstract*

As the National Aeronautics and Space Administration

(NASA) prepares to extend human presence throughout the 

solar system, technical capabilities must be developed to enable long duration flights to destinations such as near Earth 

asteroids, Mars, and extended stays on the Moon. As part of 

the NASA Human Spaceflight Architecture Team, a Technology Development Assessment Team has identified a suite of 

critical technologies needed to support this broad range of 

missions. Dialog between mission planners, vehicle developers, and technologists was used to identify a minimum but 

sufficient set of technologies, noting that needs are created by 

specific mission architecture requirements, yet specific designs are enabled by technologies. Further consideration was 

given to the re-use of underlying technologies to cover multiple missions to effectively use scarce resources. This suite of 

 

*

Formerly with NASA Headquarters

critical technologies is expected to provide the needed base 

capability to enable a variety of possible destinations and 

missions. This paper describes the methodology used to provide an architecture-driven technology development assessment (“technology pull”), including technology advancement 

needs identified by trade studies encompassing a spectrum of 

flight elements and destination design reference missions.

1.0 Introduction

NASA is preparing for the next chapter of space exploration 

by developing the capabilities needed to expand human activity 

throughout the inner solar system (Ref. 1). NASA formed the 

Human Spaceflight Architecture Team (HAT) to develop concepts for architectures and vehicle elements, conduct trade studies, and determine the technology and capability requirements 

needed for missions ranging from activities in cis-Lunar space 

to Mars landings. These activities provide cost and feasibility 
NASA/TM—2012-217670 2 

determinations to plan the next series of human exploration 

missions.

As shown in Figure 1, the HAT approach includes several 

processes: design reference missions consistent with NASA’s

investment strategy are proposed; elements needed for the 

missions are conceptualized; schedule and cost estimates for 

each element are developed; integrated schedules and flight 

manifests are determined; and total costs are estimated. A key 

step in this process is the determination of which technologies 

are needed to enable these elements and missions so that full

costs can be estimated. This paper describes the methodology 

used to provide an architecture-driven technology development assessment (“technology pull”) resulting in a list of critical technologies needed to advance human exploration of 

space beyond low Earth orbit.

The HAT created a Technology Development Assessment 

Team to manage the collection and evaluation of these technology needs. This team is comprised of representatives from 

across the Agency, ensuring input from and communication to 

a broad portion of the NASA community.

2.0 Architectural Elements and 

Destinations

Several architectural elements have been conceptualized by 

the HAT team, and many design reference missions have been 

developed to encompass a variety of destinations within the 

inner solar system. While still notional, these elements and 

missions contain enough fidelity to provide a concrete target 

for assessing the likely costs of similar missions, including the 

costs of technology development. The destinations are used to 

drive transportation systems capabilities and assess impacts of 

changes in mission assumptions. The elements and destinations currently under consideration are listed below and notional representations of the elements are shown in Figure 2. 

Architecture Elements (Ref. 2) 

• Space Launch System (SLS) 

• Multipurpose Crew Vehicle (MPCV) 

• Cryogenic Propulsion Stage (CPS) 

• Solar Electric Propulsion Stage (SEP)

• Lander

• EVA Suit (EVA)

• Space Exploration Vehicle (SEV) 

• Deep Space Habitat (DSH)

• Robotics and EVA Module (REM)

• Cargo Hauler

• Surface Elements (lunar, asteroid, Mars, and Mars moons)

Design Reference Missions (DRM)/Destinations (Refs. 3 to 8) 

• Low Earth Orbit (LEO)

• Geosynchronous and High Earth Orbit (GEO and HEO)

• Lunar Vicinity: Earth-Moon Lagrange points one and two 

(E-M L1 and L2)

• Lunar flyby and Low Lunar Orbit (LLO)

• Lunar surface

• Minimum capability, low energy Near Earth Asteroid 

(NEA)

• Full capability, high energy NEA

• Mars moon

• Mars surface

Figure 1.—HAT Cost Analysis Approach (Ref. 3). 
NASA/TM—2012-217670 3 

Figure 2.—Notional HAT Architectural Elements (Ref. 2). 

Note that the multipurpose crew vehicle and space launch 

system are needed for every destination, but the need for the 

other elements are destination and mission specific. Similarly, 

these two elements can be built with existing technology, but 

technology development is required for the other element concepts (dependent on destination).

3.0 Technology Selection Method

Conceptual designs were developed for each of the architectural elements listed in Section 2.0, and notional element performance requirements were determined for each relevant 

design reference mission to determine ballpark cost estimates.

Element conceptual designs also provide a reasonable basis for 

determining whether existing technologies are sufficient to 

provide the expected performance or whether technical gaps 

need to be filled.

The determination of technology adequacy is best accomplished by collaboration between mission planners, spacecraft 

designers and technology developers, as the planners and designers are keenly aware of what is needed and the developers 

know about technologies that may change the way the missions 

and designs are approached. The objective is to determine a 

minimum but sufficient set of technologies, noting that needs 

are created by specific mission architecture requirements, yet 

specific designs are enabled by technologies. 

The technical community submitted a list of technologies to 

the HAT element, architecture/DRM, and destination teams. 

These teams reviewed each technology and judged whether 

technology development was required to enable functionality 

of their element or successful completion of a mission to their 

destination, respectively. The teams also identified additional 

technologies, if any, which might be required. The first set of 

technologies was based primarily on work that was already 

under development for a planned lunar mission (as part of the 

NASA Constellation Program), but as the broader technical 

community was educated about the evolving elements and 

destinations, new ideas came forward that might either be required or could substantially change the way the missions 

could be accomplished. 

To simplify the review of the many technologies brought 

forth, each technology was described using a common format 

including these data elements:

• Description: Explanation as to the what and why a 

specific technology development is required

• Performance Characteristics: Details on what advancements beyond SOA is required, including metrics where 

known/applicable

If the technologists and the respective element and destination 

team leads reached consensus on the description and performance characteristics needed, and these team leads agreed that 

the technology was enabling for their vehicle or mission, then the 

technology was added to the list of critical technologies. Some 

technologies that are required for one element are also beneficial 

to others; an example is unsettled cryogenic propellant transfer 
NASA/TM—2012-217670 4 

which is required for the CPS and beneficial for the DSH, 

Lander, and surface elements. Note that the need for some technologies is dependent on specific vehicle configurations, e.g., 

oxygen-rich stage combustion engine technology is applicable to 

the SLS only if liquid strap-on boosters are included in the design. The applicability data was visually mapped by noting 

“drivers” and “beneficiaries” in a spreadsheet containing rows of 

technologies and columns of elements, DRMs, and destinations.

This mapping is notionally shown as the “technology assessment” plot in Figure 1 and the “summary spreadsheet” in Figure 

3. The cost and schedule needed to bring the technology to a 

level where it could be infused into the standard Design, Development, Test and Evaluation (DDT&E) cycle was also collected, 

along with the current technology readiness level (TRL). “Cost 

fidelity” was assessed by determining the knowledge level of 

both the problem being addressed and the costs to develop the 

technology. An example of a technology with high cost fidelity 

is in-space cryogenic propellant storage, where the requirements 

are very well understood and solid plans are in place to develop 

the required technology. A Mars surface space suit is an example 

of a technology whose needs are well understood since the Mars 

environment is well characterized, but costs to develop the technology are not because specific technical concepts have not yet 

been selected. Thermal control is an example of a technology 

whose development costs are well understood for a given thermal system design, but the needs are uncertain because spacecraft designs are not finalized. These two examples yield a 

“medium” cost fidelity. Finally, a long life battery has a “low” 

cost fidelity because technical requirements have not yet been 

determined and a technical concept has not yet been developed.

Cost phasing and “need by” dates were recorded with each 

technology to assist with HAT’s cost estimations for each DRM. 

Technologies were assumed to be matured and available by the 

preliminary design review for the enabled element. “Need by” 

dates for technologies required by multiple elements were based 

on the element that was expected to be completed first.

Finally, each technology was mapped into the technology 

classification system used by NASA’s Office of the Chief 

Technologist (Ref. 9). The full suite of data elements describing each technology is listed in Appendix B. 

Figure 3 summarizes the use of the data collected for each 

technology. Subject matter experts created technology “onepage” descriptions based on the needs of DRMs and architecture elements, and estimated technology development costs and 

the fidelity of those costs. Cost, schedule, and applicability data 

was recorded in a summary spreadsheet for use in the cost 

estimation process depicted in Figure 1. 

4.0 Technologies

Using the methodology described in Section 2.0, a suite of

60 technologies was identified as being critically important for 

at least one mission under consideration by the HAT (“technology pull”). In addition, “common avionics” was identified as a 

technology which could substantially improve system level

affordability, and four ground operations technologies were 

identified as having a similar cost reduction potential. The full 

suite of these 65 technologies is listed in Table 1, and their 

mapping to the architecture elements and destinations is shown 

in Table 2. Summary descriptions of each technology are included in Appendix C. Acronyms used in those descriptions are 

listed in Appendix A. 

Figure 3.—HAT Technology Development Assessment: Data Capture Process (Ref. 3). 
NASA/TM—2012-217670 5 

TABLE 1.—TECHNOLOGIES THAT ENABLE HUMAN SPACE EXPLORATION

Technology Area (TA) TA Breakdown Title

01 Launch Propulsion Systems 1.2 Oxygen-Rich Staged Combustion (ORSC) Engine Technology

Advanced, Low Cost Engine Technology for SLS

02 In-Space Propulsion Technologies

2.1

LO2/LCH4 Cryogenic Propulsion System

LO2/LCH4 Reaction Control Engines 

Non-Toxic Reaction Control Engines

2.2 Electric Propulsion and Power Processing

2.3 Nuclear Thermal Propulsion (NTP) Engine

2.4 Unsettled Cryo Propellant Transfer

2.4 In Space Cryogenic Liquid Acquisition

03 Space Power and Energy Storage 

3.1

300 kWe Fission Power for Electric Propulsion

High Strength/Stiffness Deployable 10-100 kW Class Solar Arrays 

Autonomously Deployable 300 kW In-Space Arrays 

Fission Power for Surface Missions 

Multi-MWe Nuclear Power for Electric Propulsion

3.2

Regenerative Fuel Cell

High Specific Energy Battery

Long Life Battery

04 Robotics, Tele-Robotics and Autonomous Systems

4.1, 4.5 Precision Landing and Hazard Avoidance

4.3 Telerobotic control of robotic systems with time delay

4.5 Autonomous Vehicle Systems Management

Common Avionics

4.6, 4.2, 4.5 Automated/Auton. Rendez. and Docking, Prox Ops, Target Relative Navigation

4.7, 6 Crew Autonomy beyond LEO

4.7, 4.4 Robots Working Side-by-Side with Suited Crew

05 Communications and Navigation

5.2 High Data Rate Forward Link (Flight) Communications

5.4 High Rate, Adaptive, Internetworked Proximity Communications

In-Space Timing and Navigation for Autonomy

5.5 Quad Function Hybrid RF/Optical Comm, Optical Ranging, RF Imaging System

06 Human Health, Life Support and Habitation Systems

6.1 Closed-Loop, High Reliability, Life Support Systems

High Reliability Life Support Systems

6.2

Deep Space Suit (Block 1)

Lunar Surface Space Suit (Block 2)

Mars Surface Space Suit (Block 3)

6.3

Long Duration Spaceflight Medical Care

Long-Duration Spaceflight Behavioral Health and Performance

Microgravity Biomedical Counter-Measures for Long Duration Spaceflight

Microgravity Biomedical Counter-Measures—Optimized Exercise Equipment

6.3, 6.1 Deep Space Mission Human Factors and Habitability

6.4, 11 In-Flight Environmental Monitoring

6.4 Fire Prevention, Detection and Suppression (reduced pressure)

6.5

Space Radiation Protection—Galactic Cosmic Rays (GCR)

Space Radiation Protection—Solar Particle Events (SPE)

Space Radiation Shielding—SPE

07 Human Exploration Destination Systems

7.1

In-Situ Resource Utilization (ISRU)—Lunar: Oxygen/Water Extraction from Lunar Regolith

ISRU—Mars: Oxygen from Atmosphere and Water Extraction from Soil

7.3

Anchoring Techniques and EVA Tools for Microgravity Surface Operations

Suit Port 

Surface Mobility

7.5, 4.7 Mission Control Automation beyond LEO

7.5 Dust Mitigation

09 Entry, Descent and Landing (EDL) Systems 9.1, 9.4 EDL Technologies—Mars Exploration Class Missions

EDL Technologies—Earth Return

11 Modeling, Simulation, Information Technology and Processing 11.2 Advanced Software Development/Tools

12 Materials, Structures, Mechanical Sys. and Mfg.

12.1, 12.2, 7.4.2 Structures and Materials for Inflatable Modules

12.1, 12.2 Lightweight and Efficient Structures and Materials

12.3 Mechanisms for Long Duration, Deep Space Missions

Low Temperature Mechanisms

13 Ground and Launch Systems Processing

13.1 Ground Systems: Low Loss Cryogenic Ground Systems Storage and Transfer

13.2 Ground Systems: Corrosion Detection and Control

13.3 Ground Systems: Fault Detection, Isolation, and Recovery

Ground Systems: Wiring Fault Detection and Repair

14 Thermal Management Systems

14.1 In-Space Cryogenic Propellant Storage (Zero Boil Off LO2; Reduced/Zero Boil Off LH2)

14.2 Thermal Control

14.3 Robust Ablative Heat Shield (Beyond Lunar Return)—Thermal Protection System

Robust Ablative Heat Shield (Lunar Return)—Thermal Protection Systems
TABLE 2.—TECHNOLOGIES MAPPED TO ELEMENTS AND DESTINATIONS

Applicable Capability/Element 

OCT 

TA 

no.

HAT Technology Development Entry (Title)

MPCVSLSCPSSEVSEPDSHLanderEVA

Space - In

Robotics

Cargo HaulerSurface ElementsOther

1.2 ORSC Engine Technology D-note 

1.2 Advanced, Low Cost Engine Technology for SLS D-note x x

2.1 LO2/LCH4 Cryogenic Propulsion System note D 

2.1 LO2/LCH4 Reaction Control Engines x D 

2.1 Non-Toxic Reaction Control Engines x x x 

2.2 Electric Propulsion and Power Processing D 

2.3 NTP Engine NTR-D 

2.4 Unsettled Cryo Propellant Transfer x x x (ISRU) Depot-D 

2.4 In Space Cryogenic Liquid Acquisition D x x Depot-D 

3.1 300 kWe Fission Power for Electric Propulsion NEP-D 

3.1 High Strength/Stiffness Deployable 10-100 kW Class Solar Arrays D-Block2 x x x D 

3.1 Autonomously Deployable 300 kW In-Space Arrays D 

3.1 Fission Power for Surface Missions D 

3.1 Multi-MWe Nuclear Power for Electric Propulsion NEP-D 

3.2 Regenerative Fuel Cell x D-note x x D 

3.2 High Specific Energy Battery x D D x 

3.2 Long Life Battery x x x x x x x x D 

4.1, 4.5 Precision Landing and Hazard Avoidance note D x 

4.3 Telerobotic control of robotic systems with time delay x x x D (REM) D-note SEV Rover 

4.5 Autonomous Vehicle Systems Management x x x x D x D x D 

4.5 Common Avionics D-note D-note x x x x D x x x 

4.6, 4.2, 4.5 Automated/Auton. Rendez. and Docking, Prox Ops, Target Relative Navigation x x D x x x x Depot 

4.7, 6 Crew Autonomy beyond LEO x x x D x x x Mission Control-D 

4.7, 4.4 Robots Working Side-by-Side with Suited Crew x x x x x D-note 

5.2 High Data Rate Forward Link (Flight) Communications x x D x In-Space Relay-D 

5.4 High Rate, Adaptive, Internetworked Proximity Communications D D D x x x x x 

5.4 In-Space Timing and Navigation for Autonomy D x D x D-note x x x SEV Rover 

5.5 Quad Function Hybrid RF/Optical Comm, Optical Ranging, RF Imaging 

System

x x D x In-Space Relay 

6.1 Closed-Loop, High Reliability, Life Support Systems x x x D ISS 

6.1 High Reliability Life Support Systems x x D x x D ISS 

6.2 Deep Space Suit (Block 1) x D 

6.2 Lunar Surface Space Suit (Block 2) D D 

6.2 Mars Surface Space Suit (Block 3) D D 

6.3 Long Duration Spaceflight Medical Care x x D x 

NASA/TM—2012-217670 6
TABLE 2.—TECHNOLOGIES MAPPED TO ELEMENTS AND DESTINATIONS

Applicable Capability/Element 

OCT 

TA 

no.

HAT Technology Development Entry (Title)

MPCVSLSCPSSEVSEPDSHLanderEVA

Space - In

Robotics

Cargo HaulerSurface ElementsOther

6.3 Long-Duration Spaceflight Behavioral Health and Performance x D 

6.3 Microgravity Biomedical Counter-Measures for Long Duration Spaceflight x D x Any Habital Volum

6.3 Microgravity Biomedical Counter-Measures—Optimized Exercise Equipment x D x Extended Stay Hab

6.3, 6.1 Deep Space Mission Human Factors and Habitability x x D x 

6.4, 11 In-Flight Environmental Monitoring x x D x x D 

6.4 Fire Prevention, Detection and Suppression (reduced pressure) x x D D D 

6.5 Space Radiation Protection—GCR x D D x D-note Extended Stay Hab

6.5 Space Radiation Protection—SPE x D D x D-note Any Habital Volum

6.5 Space Radiation Shielding—SPE x D D D (GEO) x D-note Any Habital Volum

7.1 ISRU—Lunar: Oxygen/Water Extraction from Lunar Regolith D 

7.1 ISRU—Mars: Oxygen from Atmosphere and Water Extraction from Soil D-note D Return Prop-D 

7.3 Anchoring Techniques and EVA Tools for Microgravity Surface Operations D x x 

7.3 Suit Port D x x D D (REM) D 

7.3 Surface Mobility D SEV Rover-D 

7.5, 4.7 Mission Control Automation beyond LEO x x x x x x Mission Control-D 

7.5 Dust Mitigation x x x x D (Surf) x D SEV Rover-D 

9.1, 9.4 EDL Technologies—Mars Exploration Class Missions x D x 

9.1, 9.4 EDL Technologies—Earth Return x 

11.2 Advanced Software Development/Tools x x x x x x x x 

12.1, 12.2, 7.4.2 Structures and Materials for Inflatable Modules x D-note 

12.1, 12.2 Lightweight and Efficient Structures and Materials x x x x x x x x x x x Gnd Ops (mfg.) 

12.3 Mechanisms for Long Duration, Deep Space Missions x x x x x x x x 

12.3 Low Temperature Mechanisms D-note 

13.1 Ground Systems: Low Loss Cryogenic Ground Systems Storage and Transfer Gnd Ops-D 

13.2 Ground Systems: Corrosion Detection and Control Gnd Ops-D 

13.3 Ground Systems: Fault Detection, Isolation, and Recovery Gnd Ops-D 

13.3 Ground Systems: Wiring Fault Detection and Repair Gnd Ops-D 

14.1 In-Space Cryogenic Propellant Storage (Zero Boil Off LO2; Reduced/Zero 

Boil Off LH2) 

 D x x x Depot-D 

14.2 Thermal Control x x x x x x D-note SEV Rover, Radiato

14.3 Robust Ablative Heat Shield (Beyond Lunar Return)—Thermal Protection 

System

D 

14.3 Robust Ablative Heat Shield (Lunar Return)—Thermal Protection Systems D 

NASA/TM—2012-217670 7
NASA/TM—2012-217670 8 

5.0 Discussion

It is a challenge to identify enabling technologies for missions that are only notionally defined. Primarily this is because 

“enabling” is defined as “the mission cannot be done without 

it” but there is much flexibility on how the missions are 

defined. Conversations can become circular: “Why is this 

technology needed?” gets answered with “What does this 

spacecraft or mission need to do?” Or: “This technology isn’t 

needed because the mission can be done conventionally” is met 

with “But if this particular technology is available, the mission 

can be done differently (e.g., less expensively, more robustly, 

more extensibly).” 

Nonetheless, forward progress is possible for many aspects 

where precursor information is available (for elements like 

SLS and MPCV) or where the need for a technology is evident 

regardless of element definition or DRM (e.g., very long duration missions). A particular benefit of the method described in 

Section 3.0 is that it fosters conversations across organizational 

boundaries and can lead to creative solutions.

The fidelity of the performance metrics listed in Appendix C

varies among the technologies. The application of some technologies have been well studied for specific missions, and so 

their technical descriptions and performance characteristics are 

well understood. Examples include deep space suit; fire detection and suppression; nuclear thermal propulsion; and 

LO2/LCH4 cryogenic propulsion systems. Other technologies 

are only now being considered because the missions are newly 

developed. Examples include anchoring techniques and EVA 

tools for micro-gravity surface operations. Some technologies 

are discrete and well defined (e.g., high specific energy batteries), some are innovative designs (e.g., suit port), and some are 

families of technology designs (e.g., surface mobility). These 

differences are a consequence of the iterative nature of the 

process—one must start somewhere to create data sets with 

meaning across broad constituent groups.

The original use of this technology data was as input to the 

HAT cost models used to assess mission feasibility as 

described in Section 1.0. The method has also proven useful to 

introduce new ideas into the mission planning activity as 

described above, and to focus technology development programs. For instance, NASA’s Advanced Exploration Systems 

Program looks to the performance metrics generated by this 

process to ensure that their technology development is addressing the most critical problems (Ref. 10). 

The HAT technology development data is also being used as 

a basis of input into the Office of Chief Technologist’s “Strategic Space Technology Investment Plan.” OCT is consolidating 

NASA’s technology needs across the Agency’s directorates, 

and the HAT Technology Development data set is the basis for 

the Human Exploration and Operations Mission Directorate’s 

input. 

In addition, an analysis tool has been developed to provide a 

relative comparison of the HAT identified technologies against 

Figure 4.—Correlation between technologies identified as critically important by HAT and high priority technologies identified by the NRC.

selected Figures of Merit (FOM), DRMs, and mission 

timeframes. Periodically updated HAT technology development data, along with the technology ranking analysis tool, 

provide HEOMD with the ability to weigh the technology 

development portfolio as required within the Capability Driven 

Framework.

This data set also provides a means to clearly communicate 

NASA’s human exploration technology needs to the NASA 

field Centers and potential collaborators. This data has been 

provided to the International Space Exploration Coordination 

Group, which subsequently adopted the framework for assessing technologies for the Global Exploration Roadmap and 

future partnerships (Ref. 11). We expect that it will be similarly useful for discussions with other government agencies and 

industrial and academic organizations. We hope that these 

dialogs provide new ideas regarding the means and opportunities for future human exploration.

Finally, this technology information has been useful for assessing NASA’s strategic technology roadmaps as they pertain 

to human exploration. The National Research Council (NRC) 

recently reviewed those roadmaps and provided prioritization 

recommendations including “High”, “Medium” and “Low” 

priority (Ref. 12). They further identified the “Top 16” technologies from the “High” ranked technologies recommending 

that NASA focus on those in the next 5 yr. The HAT technology listing provides a ready means to compare the NRC’s assessment of critically required technologies to those 

determined by the NASA human spaceflight community, and 

to identify errors and/or omissions from both sides. Figure 4 

shows the HAT identified technologies mapped to the NRC 

rankings. Fully one-third of the HAT technologies were among 

the NRC’s Top 16, and 74 percent were ranked “High.” This 

indicates excellent agreement between the two communities. 
NASA/TM—2012-217670 9 

About one-fifth of the HAT technologies were ranked lower 

than “High” by the NRC. Primarily this occurred because long 

duration reliability was not a ranking factor for the NRC while 

HAT placed high importance on this, some HAT technologies 

are design specific and did not get captured by the NRC process, and the NRC did not consider ground operations to require technology development. The NRC’s “Top 16” listing is 

shown in Table 3, with those that overlap with the HAT listing 

shown in bold. Note that in some cases several HAT technologies group to form one NRC technology; for example, the 

NRC “Fission (Power)” encompasses three HAT technologies:

fission power for surfaces, for 300 kW-class electric propulsion systems, and for multi-MW electric propulsion systems.

TABLE 3.—NRC’S “TOP 16” TECHNOLOGIES, WITH HAT 

HUMAN EXPLORATION FOCUSED CRITICAL 

TECHNOLOGIES SHOWN IN BOLD.

NRC’s “Top 16” Technologies for NASA

2.2.1 Electric Propulsion 

2.2.3 (Nuclear) Thermal Propulsion 

3.1.3 Solar Power Generation (Photovoltaic and Thermal) 

3.1.5 Fission (Power) 

4.2.1 Extreme Terrain Mobility 

6.3.2 Long-Duration (Crew) Health 

8.1.1 Detectors and Focal Planes 

8.1.3 (Instrument and Sensor) Optical Systems 

8.2.4 High-Contrast Imaging and Spectroscopy Technologies 

8.3.3 In Situ (Instruments and Sensor) 

14.1.2 Active Thermal Control of Cryogenic Systems

X.1 Radiation Mitigation for Human Spaceflight 

X.2 Lightweight and Multifunctional Materials and Structures 

X.3 Environmental Control and Life Support System

X.4 Guidance, Navigation, and Control

X.5 Entry, Descent, and Landing Thermal Protection Systems 

The method described in this paper is adaptable and allows 

for additions and deletions as DRMs evolve. We continue to 

learn more uses for the data and consider it a rich set of information to guide plans for the next chapter of human spaceflight.

6.0 Summary

NASA’s HAT Technology Development Assessment Team 

has developed a method to identity technologies that will enable 

the next chapter of human space exploration. This method relies 

on dialog between system designers and technology developers 

to determine a minimum but sufficient set of technologies, noting that needs are created by specific mission architecture capabilities, yet specific designs are enabled by technologies. 

The list of technologies that have been identified by this process was presented, along with brief descriptions and performance goals for each technology. These performance goals and 

descriptions may change as future exploration missions become more developed, noting that those missions will be influenced by the availability of these technologies.

In addition to contributing to cost assessments for a variety 

of design reference missions, this information is being used to 

foster collaborations with the international community through 

the ISECG and the Global Exploration Roadmap, to focus 

technology investments within NASA, and to assess the technology roadmaps being developed within NASA.

NASA/TM—2012-217670 11

Appendix A.—Acronyms

AAES Aeroassist, Aerocapture, and Entry Systems

AR Atmospheric Revitalization

AU Astronomical Unit

BAC Broad Area Cooling

BPP Bubble Point Pressure

CDR Critical Design Review

CFM Cryogenic Fluid Management

CH4 Methane

COTS Commercial Off the Shelf

CPS Cryogenic Propulsion Stage

CTV Crew Transport Vehicle (aka MPCV)

CxP Constellation Program

DDT&E Design, Development, Test and Evaluation

DRA Design Reference Architecture

DRM Design Reference Mission

DSH Deep Space Habitat

DSN Deep Space Network

ECLS Environmental Control and Life Support

EDL Entry, Descent, Landing

E-M Earth-Moon

EVA Extravehicular Activity

FDIR Fault Detection, Fault Isolation, and Recovery

GCR Galactic Cosmic Rays

GEO Geosynchronous Earth Orbit

HAT Human Space Flight Architecture Team

HEFT Human Exploration Framework Team

HEO High Earth Orbit

HEOMD NASA’s Human Exploration and Operations Mission Directorate

HRP NASA’s Human Research Program

HSF Human Space Flight

IMM Integrated Medical Model

ISCEG International Space Exploration Coordination 

Group

ISRU In-Situ Resource Utilization

ISS International Space Station

IV Intravenous

IVA Intravehicular Activity

L1 Lagrange Point 1

L2 Lagrange Point 2

LAD Liquid Acquisition Device

lbf Pound-force

LCH4 Liquid Methane

LEE Latching End Effectors

LEO Low Earth Orbit

LH2 Liquid Hydrogen

LLO Low Lunar Orbit

LO2 Liquid Oxygen

MCC Mission Control Center

MLI Multilayer Insulation

MMOD Micrometeorites and Orbital Debris

MPCV Multipurpose Crew Vehicle

NASA National Aeronautics and Space Administration

NEA Near Earth Asteroid

NEO Near Earth Object

NRC National Research Council

NTP Nuclear Thermal Propulsion

OCT NASA’s Office of the Chief Technologist

ORSC Oxygen-Rich Staged Combustion

PDR Preliminary Design Review

PLSS Portable Life Support System

PMAD Power Management and Distribution

PSR Perennially Shadowed Regions

RBO Reduced Boil Off

RCS Reaction Control System

REM Robotics and EVA Module

RF Radio Frequency

RFC Regenerative Fuel Cell

RP Rocket Propellant

SARJ Solar Alpha Rotary Joint

SEP Solar Electric Propulsion stage

STS Space Transportation System

SLOC Source Line of Code

SLS Space Launch System

SOA State of the Art

SRR Strategic Readiness Review

SPE Solar Particle Events

SSRMS Space Station Remote Manipulator System

TA Technical Area

TBD To Be Determined

TRL Technology Readiness Level

TPS Thermal Protection System

WR Water Recovery

ZBO Zero Boil Off 

NASA/TM—2012-217670 13

Appendix B.—Technology Description Format

The following information was collected for each technology listed in Table 1. 

• Title: Brief descriptive title of the technology development

• Discipline: HAT categorization of the technology (e.g., 

Chemical Propulsion, Advanced Propulsion, etc.)

• OCT TA no.: Cross reference to the NASA OCT Technology Breakdown Structure

• Version/Date: Used for tracking revisions

• Description: Explanation as to the what and why a specific technology development is required

• Performance characteristics: Details on what advancements beyond the current SOA is required, including metrics where known/applicable

• Applicability to Capabilities/Elements; Destinations/Con-Ops: Identifies what Architecture Elements 

and DRMs are applicable to the technology entry. “Driving” indicates a technology required (i.e., technology 

pull) by a capability/element, (per element overview)

and/or destination/con-op (per DRM), and “Beneficiary” 

similarly indicates those that benefit from the technology 

if available and matured/advanced to a sufficient degree 

(e.g., technology push)

• Current TRL: Estimate of the current technology readiness level.

• ISS Demonstration: Identifies International Space Station utilization for technology demonstration of the applicable technology entry (‘Planned’ = integral to the 

technology development plan; ‘Candidate’ = potential 

ISS utilization; ‘None’ = no ISS utilization identified). 

• Cost to infuse into standard DDT&E cycle: Cost and

time estimate for the technology development. The monies and time typically required for advancing the technology to TRL 6, at which time a standard DDT&E 

production process would integrate the technology into 

the flight assembly/system. Additional information is also provided on how the cost/time estimate was determined, and also any significant demonstrations included 

in the cost are listed.

• Cost Fidelity: A 5x5 matrix tool used to help understand 

the fidelity of the cost estimate. The two axes are “How 

well do we know what is required?” and “How well defined is the plan?” Requirement definition ranges from 1 

(very poorly understood) to 5 (well known), and the plan 

ranges from 1 (little or no plan) to 5 (full, multi-year project plan).

• Cost Phasing: An estimate of the profile type to be used 

by the HAT Cost Assessment Team (Standard Phasing or 

Level of Effort) 

• Need By Date: Identifies the Program or Element 

DDT&E milestone (e.g., reviews such as SRR, PDR, 

CDR) at which the technology is targeted to be matured 

to TRL 6 for infusion into the Program or Element.

• Architecture Elements Mapping: matrix shows each 

identified Architecture Element mapped across the full 

range of technology development entries. A green highlighted field with a “D” indicates that particular technology development is required for that Element 

(technology pull). An “x” in the matrix represents benefits of the technology developments mapped against the 

Elements (technology push). 

• Destination/DRM Mapping: matrix shows each HAT 

Cycle 2011-C destination DRM’s mapped across the full 

range of technology development entries. A green highlighted field with a “D” indicates that particular technology development is required for that particular DRM 

(technology pull). An “x” in the matrix represents benefits of the technology developments mapped against the 

DRMs (technology push). 

• HAT Technology Development Assessment Team 

Point of Contact

• Subject Matter Expert Point of Contact

• Notes: Any additional information applicable to the 

technology development entry

NASA/TM—2012-217670 15

Appendix C.—Technology Descriptions

The following list summarizes the technical description and

performance characteristics for each of the technologies listed 

in Table 1. The other data entries listed in Appendix B are not 

included in this list. Acronyms are spelled out in Appendix A. 

The technologies are grouped according to the technology 

areas determined by NASA’s Office of the Chief Technologist.

In some cases a single HAT technology maps to multiple OCT 

technical areas. Technology Area TA 08 Science, Instruments, 

Observatories and Sensor Systems, and TA 10 Nanotechnology, are not used by the HAT. 

Note that this summary is still a work in progress, and differences exist in the level of detail provided. Some of these 

differences are caused by differing levels of fidelity currently 

existing between the reference missions; others are caused by 

differing TRL levels of the technology.

This data is based on information collected during 2011 CycleC of the HAT process, and is current as of December 8, 2011.

C.1 TA 01—Launch Propulsion 

Enhance existing solid or liquid propulsion technologies by 

lower development and operations costs, improved performance, availability and increased capability.

C.1.1 Oxygen-Rich Staged Combustion (ORSC) Engine 

Technology

Chemical Propulsion (OCT TA 1.2) 

Description

• ORSC LO2/RP engine supports the Space Launch Systems (SLS) evolution to larger Payloads. ORSC LO2/RP 

engine has significant DDT&E risks associated with 

closed cycle and high system pressures (combustor/turbomachinery.

Performance Characteristics

• Improved capability for Hydrocarbon performance 

+17 percent Isp, +60 percent T/W, –30 percent cost, and 

–75 percent failure rate over SOA

C.1.2 Advanced, Low Cost Engine Technology for SLS

Chemical Propulsion (OCT TA 1.2) 

Description

• By mitigating the largest contributor to recurring costs, 

low-cost advancements in near- and far-term engine 

manufacturing technology would enable exploration systems to meet affordability and sustainability requirements. This includes innovative approaches to metals and 

non-metals engineering and fabrication that significantly 

minimize operations and reduce touch labor with minimal to no compromise in material properties (e.g., metal 

stereo lithography, chemical etching, and advanced welding/joining techniques). 

Performance characteristics 

• Subscale demonstration for booster application with fullscale benefits for in-space technology

• Develop and demo technologies to reduce touch labor 

during manufacture (number of welds, machining, joints, 

parts, assembly w/o sacrifice to performance) 

• Reduce chemical propulsion recurring cost of near-term 

engines suitable for SLS

• Provide acceptable performance at half the price of currently available systems 

C.2 TA 02—In-Space Propulsion

Advancements in conventional and exotic propulsion systems, 

improving thrust performance levels, increased payload mass, 

increased reliability, and lowering mass, volume, operational 

costs, and system complexity. 

C.2.1 LO2/LCH4 Cryogenic Propulsion System

Chemical Propulsion (OCT TA 2.1) 

Description

• An In-Space Stage, powered by a demonstrated workhorse 

engine, intended for mission applications beyond LEO.

• The oxygen and methane propellant combination has the 

potential for good engine performance, which can result in 

lower vehicle mass and greater payload-carrying capability.

Performance characteristics 

• Improved handling and non-toxicity benefit of the 

LCH4/oxygen combination (hours rather than days 

ground operations) 

• Approximately 10 percent specific impulse performance 

improvement relative to hypergolic systems. 

C.2.2 LO2/LCH4 Reaction Control Engines

Chemical Propulsion (OCT TA 2.1) 

Description 

• The oxygen and methane (LCH4) propellant combination 

has the potential for greater engine performance, which 

can result in lower vehicle mass and greater payloadcarrying capability. 

• Demonstrated performance of a TRL 6 engine including: 

○ Specific impulse of 317-s; Impulse bit of 4 lbf-s; 

50,000 cycles with a cryogenic valve; 

○ Ignition and operation over a range of propellant inlet 

conditions (liquid/liquid to gas/gas)

Performance characteristics 

• Improved handling and non-toxicity benefit of the 

LCH4/oxygen combination (hours rather than days 

ground operations). 

• Approximately 10 percent specific impulse performance 

improvement relative to hypergolic systems. 
NASA/TM—2012-217670 16

Non-Toxic Reaction Control Engines

Chemical Propulsion (OCT TA 2.1)

Description 

• Propulsion system technologies for non-toxic or “green” 

propellants for use in reaction control systems.

• Non-toxic technologies for RCS engines over the thrust 

range of 25 to 1000 lbf. Propellant options include hypergolic ionic liquids and nitrous oxides monopropellants, 

both of which can be easily stored in space and on the 

ground.

Performance characteristics 

• Improved handling and non-toxicity benefit of hours 

rather than days ground operations. 

• Non-toxic bipropellant or monopropellants that have 

higher specific impulse (greater than hypergolic) and/or 

high specific impulse density (greater than hypergolic) 

with better safety and reduced handling risks

C.2.3 Electric Propulsion and Power Processing

Advanced Propulsion (OCT TA 2.2) 

Description

• Solar electric vehicles are required for the economical 

transport of equipment from LEO to HEO and crew from 

HEO to a NEO, as well as cargo from LEO to Mars, because they can significantly reduce the number of heavy 

lift launches required and can decrease sensitivity to 

mass growth of other in-space elements. They also increase crew safety by providing multiple engines for 

more robust off-nominal operations. 

• A propulsion system requiring nominally 300 kW of 

electrical power is required for these missions; likely an 

array of 30 to 50 kW thrusters will be used. In addition to 

designing, building, and testing high power thrusters, 

technology development is required for power processing, power distribution, and propellant storage. Determining the performance of the integrated power and 

propulsion system is needed to design the subsystems 

since the required performance represents such a large 

increase relative to the SOA. Data of interest include the 

interaction of the thruster plumes with the high-voltage 

solar array, array degradation from the Van Allen belts, 

and guidance, navigation and control of the SEP vehicle 

with large, flexible solar arrays.

Performance characteristics 

• High power (~400 kW power at beginning of life)

• High specific impulse (~2000 s) 

• Low mass (< ~45 mt wet mass with mass growth allowance to fit within a 100 mt launch vehicle)

C.2.4 Nuclear Thermal Propulsion (NTP) Engine

Advanced Propulsion (OCT TA 2.3) 

Description 

• Nuclear thermal propulsion (NTP) was identified by 

NASA’s DRA 5.0 as required for economical transport 

of crew to Mars because it provides the high thrust and 

high specific impulse needed to significantly reduce 

launch mass for the heavy payloads identified (Ref. 13). 

The NTP system would also reduce the cost of transits to 

the Moon, E-M L1, NEOs, and orbital missions to Mars 

and its moons. The NTP system consists of two principal 

components. The first component is the primary NTP 

stage that includes the nuclear thermal rocket engines, 

RCS, avionics, auxiliary power, long duration CFM for 

the LH2 propellant and docking capability. The second 

component is an integrated saddle truss and LH2 drop 

tank assembly connecting the NTP stage to the mission 

payload that provides additional propellant storage for a 

wide range of mission and payload needs.

• NTP has strong synergy with chemical rocket hardware 

and can use the same LH2 tanks in the launch vehicle. It 

can be developed in a timely manner at reasonable cost 

and can service both NEOs and Mars with same vehicle 

components helping to reduce overall cost.

Performance Characteristics

• High thrust (10’s of klbf) 

• High Isp (~900 s) propulsion 

• Three 25,000 lbf nuclear thermal rocket engines with either NERVA-derived or ceramic-metallic (cermet) fuel. 

• Long duration cryogenic fluid management.

C.2.5 Unsettled Cryogenic Propellant Transfer

Cryogenic Fluid Systems (OCT TA 2.4) 

Description

• Efficient transfer of cryogenic fluids in-space is required 

for propellant resupply to a Cryogenic Propulsion Stage 

(CPS) and/or oxygen resupply to a Deep Space Habitat 

(DSH) and has direct planetary application to ISRU Surface Systems. The SOA for propellant transfer in cryogenic upper stages requires the use of an ancillary propulsion 

system to settle the cryogenic propellants at the tank outlets and a helium pressurant system to maintain a constant 

tank pressure (LO2 only) during propellant transfer. After 

engine start up the thrust generated by the propulsion system maintains the propellants at the tank outlet and the 

LH2 tank uses an autogenous gaseous hydrogen pressurant 

system. This is not possible for tank-to-tank transfers; “unsettled” transfer is also beneficial for propellant resupply 

of large tank-to-propulsion systems.

• A pumped transfer at unsettled conditions and without a 

liquid acquisition device in the storage requires a twophase fluid tolerant pump for liquid transfer. A transfer 

process requires a robust leak-free fluid transfer coupling 

to mate the storage tank and the propulsion system receiver tank, an efficient transfer line chill down technique to minimize the liquid used to chill down the 

transfer line and a micro-g gauging concept to verify the 
NASA/TM—2012-217670 17

high fill fraction of the propulsion system receiver tank. 

An automated propellant leak detection system would 

ensure the safe in-space transfer operation.

• These technologies are also directly applicable to LO2

/Methane propellant systems.

In-Space Performance Characteristics (< 0.00003 g).

• Two Phase Fluid Tolerant Transfer pump: operation to a 

vapor fraction of ~ 0.8 with cryogenic fluids

• Automated Fluid Coupling: leakage < 10-3 sccs gHe after 1000 cycles

• Leak Detection: TBD

• Mass gauging: < 2 percent uncertainty of measurement

• Fill Fraction of propulsion system receiver tank: > 0.9

• Minimum Fluid used to chill transfer lines: <1 percent of 

transfer line mass

C.2.6 In Space Cryogenic Liquid Acquisition

Cryogenic Fluid Systems (OCT TA 2.4) 

Description

• Cryogenic liquid acquisition technology is needed for 1) 

unsettled tank-to-tank propellant transfer, 2) unsettled 

tank-to-engine propellant transfer, and 3) propellant 

transfer into heat exchangers needed to maintain propellant tanks at required temperature and pressure. It is important to transfer only cryogenic liquids for these 

applications, without transferring ullage gas. Propulsive 

maneuvers can be used to settle the cryogens to ensure 

liquid-only transfer, but this parasitic propellant burn increases system mass, particularly for the frequent transfers needed for the thermodynamic vent system for tank 

pressure and temperature control.

• In micro- and reduced-gravity, liquid tends to cling to the 

walls of the tank, making it difficult to sufficiently cover 

the tank outlet during fluid outflow.

• An in-space liquid acquisition device (LAD) is required to 

acquire vapor-free liquid from a propellant tank in 

micro-g. LADs represent the first stage in successful fluid 

transfer from a tank to a propulsion system (or another 

tank). LADs rely on surface tension forces to separate liquid and vapor in the tank and capillary flow to maintain 

communication between liquid and the outlet during 

expulsion.

• A second system required for in-space liquid acquisition 

for large propellant storage and long duration missions is 

an autogenous pressurant system. Helium pressurant 

supply is impractical for these missions due to the helium 

mass required and the large launch mass penalty. An alternative to helium pressurization would be to extract a 

small amount of liquid or two phase fluid and feed it 

though a heat exchanger to vaporize the liquid and return 

it to the tank as a pressurant.

• These technologies are directly applicable to LO2/CH4

propellant systems.

• LADs have a proven flight heritage when using higher 

surface tension storable liquids (e.g., hydrazine), but 

have not yet been tested in cryogenic liquids (H and O) 

in low-g environments. 

In-Space Performance Characteristics (< 0.00003 g)

• Ratio of LAD delivery system pressure drop to BPP drop 

at maximum outflow rate - < 0.75 to 0.5

• Percent of LAD residual LH2 mass to total tank LH2

mass (Expulsion efficiency) - < 1 to 3 percent 

• Ratio of total autogenous pressurant system mass to the 

mass of equivalent helium pressurant system - < 0.8 to 1.0 

C.3 TA 03—Space Power and Energy Storage 

Improvements to lower mass and volume, improve 

efficiency, enable wide temperature operational range and 

extreme radiation environment over current SOA space photovoltaic systems, fuel cells, and other electrical energy generation, distribution, and storage technologies.

C.3.1 300 kWe Fission Power for Electric Propulsion

Power Systems (OCT TA 3.1) 

Description

• Fission power systems being developed for surface applications can be used to power electric propulsion vehicles.

Performance characteristics

• Moderate power, low mass (<30 kg/kWe) power system 

for Nuclear Electric Propulsion

• 1200 K Li-cooled unfueled reactor, 2 x 340 kWe Brayton 

power conversion, 500 V power management and distribution

C.3.2 High Strength/Stiffness Deployable 10 to 100 kW 

Class Solar Arrays

Power Systems (OCT TA 3.1)

Description

• High power, high voltage, autonomously deployable surface solar arrays in 1/6th to 1/3rd gravity environments are 

needed to generate reliable electric power for surface 

outpost elements over the mission duration. In addition, 

applications for in-space use with flight elements 

requires operations during low-g accelerations under 

propulsion (0.1g). Enabling features include compact 

stowage, reliable deployment in partial gravity, on an 

irregular surface and dusty environment, Martian wind 

load strength, EVA compatibility, dust mitigation to limit 

photovoltaic power degradation and robust to surface 

arcing environment (Martian surface triboelectric charging). Few options exist and only at the conceptual level.

These include mast deployed vertical, Sun-tracking blanket solar arrays for lunar polar surface mission and horizontally deployed, fixed tent like solar arrays. Solar array 

panels would employ low mass, flexible panel substrates 
NASA/TM—2012-217670 18

populated with advanced photovoltaic cells, like inverted 

metamorphic triple junction solar cells, with bandgap 

tuning for the Martian surface solar spectrum substrates.

• These solar arrays would power outpost surface elements 

(e.g., habs/labs, rovers, ISRU, lander/ascent stages, etc.)

• These solar arrays would power in-flight space elements 

(e.g., CPS, DSH)

Performance characteristics

• High power (10-100 kW), 

• High voltage (<~200 V) 

• Autonomously deployable surface solar arrays in 1/6th to 

1/3rd gravity environments

• Operational under low-g propulsion accelerations (0.1g)

C.3.3 Autonomously Deployable 300 kW In-Space Solar 

Arrays

Power Systems (OCT TA 3.1)

Description 

• High power, high voltage, autonomously deployable solar arrays are required to generate reliable electric power 

for the SEP Stage over its mission duration. Enabling 

features include compact stowage, reliable deployment, 

~0.1-g deployed strength and robust performance 

through the mission end-of-life. Leading options include 

large, dual-wing structures (2 x 200 kW) and modular, 

sub-wing structures (20 x 20 kW) employing advanced 

photovoltaic cells on flexible substrates. Fine pointing 

requirements for concentrator-based arrays may limit 

functionality for some missions, so both planar and concentrator architectures should be considered.

Performance characteristics 

• High power (~400 kW at beginning of life)

• High voltage (~ 350 V)

• Low mass and low stowed volume (TBD W/kg and 

W/m3

) 

• Cost (2X reduction)

C.3.4 Fission Power for Surface Missions

Power Systems (OCT TA 3.1) 

Description

• Abundant power for surface missions is enabled by a surface-emplaced fission reactor. The availability of substantial amounts of continuous power provides 

opportunities for significant science, exploration, and engineering activities on Mars and the Moon.

Performance characteristics

• 40 kWe Fission Power System (reactor, power conversion, heat rejection, PMAD)

• 900 K reactor, 10 kWe Stirling convertors, 400 K radiators, 400 V PMAD

• 150 kg/kWe for surface missions 

C.3.5 Multi-MWe Nuclear Power System for Electric 

Propulsion

Power Systems (OCT TA 3.1) 

Description

• Nuclear power system development for very high power 

electric propulsion vehicles to deliver cargo and/or crew 

to Mars. Once built, this system would also reduce the 

cost of transits to the Moon, E-M L1, NEOs, and the 

Martian moons. 

Performance characteristics

• High (>1 MWe) power, low mass (<15 kg/kWe) power 

system for nuclear electric propulsion.

• Flight power system development and qualification

C.3.6 Regenerative Fuel Cells

Power Systems (OCT TA 3.2) 

Description

• Long duration energy storage is required for extended 

surface missions to store solar energy and provide power 

during low insolation. Applicable to Lunar or Mars surface applications requiring high power and/or long sortie 

durations. 

• RFC system includes a fuel cell and an electrolyzer, each 

of which can be used independently for power/water 

generation and H2/O2 generation, respectively. Electrical 

power can be used for any vehicle. Water and O2 can be 

used for life support for crewed vehicles. Also applicable 

to ISRU.

• Technology development includes reducing the number 

of ancillary components to increase reliability and operational lifetime, and reduce parasitic power losses, mass, 

and volume.

Performance characteristics

• Power generation >10 kWe for 8 hr or more

• Operable with reactants at >2000 psi to reduce tank 

volume

• Round trip energy conversion efficiency > 50 percent

• Minimize mass (TBD Wh/kg)

• Operational life >10,000 hr

C.3.7 High Specific Energy Batteries

Power Systems (OCT TA 3.2) 

Description

• Batteries with very high specific energy and energy density are required to enable untethered EVA missions lasting 8 hr within strict mass and volume limitations.

Batteries are expected to provide sufficient power for life 

support and communications systems, and tools including video and lighting. Advanced batteries are enhancing 

for every other vehicle.
NASA/TM—2012-217670 19

Performance characteristics

• Battery-level specific energy > 325 Wh/kg and energy 

density > 540 Wh/liter 

• 8 hr operation per mission over an operating temperature 

of 10 to 30 °C.

• Nominally 100 cycles and 5 yr calendar life 

C.3.8 Long Life Batteries

Power Systems (OCT TA 3.2) 

Description

• Long life and low temperature survivable batteries will 

enable lunar night survival and operations. Polar Craters 

Ops will require batteries that can survive a cryogenic 

thermal environment.

Performance characteristics

• Battery-level specific energy > 220 Wh/kg and energy 

density > 410 Wh/liter at a C/10 discharge rate

• Operate at lunar night temperatures for 14 d 

• Operate in a perennially shadowed region such as a polar 

crater

C.4 TA 04—Robotics, Tele-Robotics, and Autonomous Systems 

Improvements in mobility, sensing and perception, manipulation, human-system interfaces, system autonomy are needed. 

Advancing and standardizing interfaces for autonomous rendezvous and docking capabilities will also be necessary to 

facilitate complex in-space assembly tasks.

C.4.1 Precision Landing and Hazard Avoidance

EDL (OCT TA 4.1, 4.5)

Description

• Need autonomous landing and hazard avoidance systems, including terrain relative navigation, that operate in 

all lighting conditions, including darkness. Autonomous 

Landing and Hazard Avoidance Technology would enable a first of a kind development for planetary precision 

landing and hazard avoidance. 

Performance characteristics

• The components and techniques have been simulated and 

tested to TRL 5 but a full set of integrated field test is 

needed to show TRL 6 and applicability to future missions

• Need 90-m accuracy at 3-σ uncertainty relative to premission identified landing location. Need 0.3 m hazard 

recognition and avoidance.

C.4.2 Telerobotic Control of Robotic Systems with Time 

Delay

Robotics and Mobility (OCT TA 4.3) 

Description 

• Enable astronauts in vehicle, habitat, or EVA to remotely 

operate robots at destinations (natural environment and variable time-delay) to collect samples, deploy instruments, etc.

○ IVA SOA = control of robot arm in structured environment with man-made payloads and zero-delay 

(e.g., ISS crew uses SSRMS to move/position cargo 

modules).

○ EVA SOA = none (no EVA control of external space 

robots exists).

• Enable Earth ground control to remotely operate robots 

in dynamic environments beyond LEO to support crew 

(e.g., reconnaissance, survey, site prep, follow-up, etc. 

during sleep periods) 

○ Ground control SOA = Single command sequence 

per day of slow ground robot in static environment 

without humans (e.g., Mars Exploration Rovers driving few m/d) 

• Enable use of robots deployed by precursor mission, raceahead or crew in mixed ops modes: before—supporting—

after crew, ground control and crew, IVA and EVA

Performance characteristics 

• IVA: Advance SOA to enable telerobotics from inside 

crew vehicle (e.g., approach/orbit NEO)

• Robot functions: detail reconnaissance, sample collection, worksite prep, etc. 

○ Time-delay: 5 s (orbit-to-surface) to 5 min (for raceahead mission architectures)

• EVA: Advance SOA to enable telerobotics from suited 

crew (in-space or on-surface)

○ Robot functions: mobile camera, materials/payload 

transport, etc.

○ Time-delay: up to 10 s 

• Ground control: Advance SOA to enable telerobotics in 

dynamic environments (e.g., tumbling NEO)

○ Robot functions: initial reconnaissance, systematic 

survey, site prep, follow-up, etc. 

○ Time-delay: up to 40 min (Earth to Mars orbit roundtrip)

C.4.3 Autonomous Vehicle Systems Management

Avionics and Software (OCT TA 4.5) 

Description 

• Enables autonomous vehicle management with limited 

crew effort and little to no ground oversight. This autonomous capability is required to ensure safe vehicle operations and monitoring of complex systems, especially at 

increased distances from Earth where communications 

time delays are present. 

Performance characteristics 

• Enable on-board vehicle systems management for mission 

critical functions at destinations with > 3 s time delay

• Enable autonomous nominal operations and FDIR for 

crewed and uncrewed systems
NASA/TM—2012-217670 20

• Reduce on-board crew time to sustain and manage vehicle by factor of 2x at destinations with > 6 s time delay 

(see “Crew Autonomy” description) 

• Reduce Earth-based mission ops “back room engineering” requirements for distant mission support delay (see 

“Mission Control Automation” description) 

C.4.4 Common Avionics

Avionics and Software (OCT TA 4.5) 

Description

• Develop common avionics components such as flight 

computers, sensors, high performance, environmentally 

tolerant, interoperable computing and data busses which 

can be utilized by multiple vehicles. This approach provides support for:

○ Multiple architectures to enable single spares to fulfill multiple electronic functions,

○ Adaptability to system failures,

○ Redundancy by providing adaptable spares, and

○ Multiple interconnection options.

Performance characteristics 

• Exceed 75 percent commonality of avionics components 

across HAT DRM elements for reusability (on-orbit 

spares) and supportability

• Enable up to 1/3 of Planning and Analysis software tools 

(used in MCC “backroom” today) to be run onboard the 

vehicle

• Reduce power use by 30 percent for same processing power

• Reduce avionics weight by 50 percent for same processing power

• Improve reliability of avionics components, thereby improving crew safety and reducing logistics mass

C.4.5 Automated/Autonomous Rendezvous and Docking, 

Proximity Operations and Target Relative

Navigation

AR&D (OCT TA 4.6, 4.2, 4.5)

Description 

• Maturation of subsystem technologies (relative navigation sensors, GN&C flight software, system managers, 

and mechanisms) to accomplish autonomous rendezvous 

and proximity operations for various in-space destinations such as satellite servicing and NEA exploration.

The benefit of this technology development is to improve 

human safety, improve mission performance and flexibility by enabling autonomous rendezvous and proximity

operations interactions with complex or uncontrolled 

planetary bodies. 

Performance characteristics

• System performance driven by the need for autonomous 

operations; high reliability, rapid missionization, rendezvous with non-cooperative targets with unknown geometry, tumbling attitude, and unknown surface features; and 

mass/power constraints. Rendezvous missions include flybys of destinations without landing or docking. Proximity 

operations require loiter at destinations with zero relative 

velocity. Major challenges include the ability to rendezvous and dock in all ranges of lighting, work across near to 

far range, and achieve a docked state in all cases.

C.4.6 Crew Autonomy Beyond LEO

Avionics and Software (OCT TA 4.7, TA 6) 

Description

• Autonomous Crew Operations (planning, commanding, 

fault recovery, maintenance) in Beyond LEO missions. 

Systems and Tools to provide the crew with independence from Earth-based ground operations support. Such 

crew autonomy is essential to accommodate the ground 

communication delays and blackouts at distant locations. 

Performance characteristics

• Enable crew nominal operation of vehicle or habitat at 

destinations with > 6 s time delay to ground

• Enable coordinated ground and crew nominal operations 

at destinations with > 6 s time delay (See “Mission Control Automation” description) 

• Enable crew to detect off nominal situations and put vehicle in safe configuration without ground coordination 

C.4.7 Robots Working Side-by-Side with Suited Crew

Robotics and Mobility (OCT TA 4.7, 4.4) 

Description

• Human mission activities can be performed more effectively if robotically assisted. Coordinated efforts between 

humans and machines/robots can improve the mission 

risk/productivity trade space. 

• The top technical challenges in human-robot interactions 

are multi-sensor feedback, understanding and expressing 

intent between humans and robots, and supervised autonomy of dynamic/contact tasks.

• When robots and humans need to work in close proximity, sensing, planning, and autonomous control system for 

the robots, and overall operational procedures for robots 

and humans, will have to be designed to ensure human 

safety around robots.

• The goal is to enable EVA crew and machine interaction 

without real-time control and support needed from IVA 

or ground control personnel. 

Performance Characteristics 

• Avoid need for IV robot controller Avoid need for IV 

spotter/checker Avoid dependence on Mission Control 

• Create force level safety for proximity operations.

• Create multi-modal human-robot interfaces and autonomy software.
NASA/TM—2012-217670 21

• Create fault tolerant free flyer and EVA positioning 

technology.

• Create asteroid sampling, processing, manipulation.

• Create asteroid grappling and anchoring technology.

C.5 TA 05—Communication and Navigation

Technology advancements to enable higher forward and return link communication data rates, improved navigation precision, minimizing latency, reduced mass, power, volume and 

life-cycle costs.

C.5.1 High Data Rate Forward Link (Flight) 

Communications

Communications (OCT TA 5.2)

Description

• Combine transmitters on the ground across an array of 

antennas to produce uplink data rates 3-4 orders of magnitude higher performance than current DSN capabilities.

• Supports uplinked video, imagery and software uploads. 

Enables spacecraft receiver to receive high data rate with 

reduction avionics size, weight and power burden to Elements. Leverages navigation improvements in orbit determination accuracy and trajectory management from 

improved communication link. 

Performance characteristics

• Enable uplink rates: 25-50 Mbps at 1 AU using X-band

• Size and weight reduction: compared to currently 

achievable receiver: >50 percent

• Leverage navigation improvements in orbit determination accuracy and trajectory management from improved 

communication link

C.5.2 High-Rate, Adaptive, Internetworked Proximity 

Communications

Communications (OCT TA 5.4)

Description

• Enable high data rate communications between multiple inspace elements for situational awareness, enable element 

proximity radios to sense RF conditions and adapt autonomously, enable elements to store, forward, and relay/route 

information to other elements intelligently and when communications is available, enable element radios to be reprogrammed from ground based on in-situ characterization of 

the NEO environment. The benefit of this technology development is to improve situational awareness and communications, improving operational efficiency.

Performance characteristics

• Data rate: >20 Mb/s simultaneously between peers

• Employ multiple frequency/modulation/coding/ power 

schemes, including low frequency schemes to enable low 

rate, non-line of sight communication through small 

NEO’s when relay through other elements is not available. (Max range: < 20 km. Max NEO size for penetration: < 50 m)

• Max storage time: <5 min/Element at 20 Mb/s

• Max routing: <20 destinations and/or elements 

• Enable radios to be adapted in frequency of operation, 

modulation and coding to information as it is discovered 

about the NEO environment in near real-time. (Near realtime: < 30 min of each NEO characterization performed 

by in-space elements)

C.5.3 In-Space Timing and Navigation for Autonomy

Communications (OCT TA 5.4)

Description

• Enable elements to perform independent navigation during 

complex in-space maneuvers, enable precision required for 

absolute and relative navigation for in-space elements, enable increased element onboard reference timing generation, 

timekeeping, distribution and inter-element synchronization to eliminate dependence on Earth-based systems. The 

benefit of this technology development is to improve situational awareness and communications, improving operational efficiency. High-precision timekeeping significantly 

reduces accumulated navigation error over long periods of 

time, enabling mission autonomy for long periods of time 

without synchronization events with ground or other (x-ray, 

etc.) synchronization. 

Performance characteristics 

• Complex maneuvers: navigating amongst multiple inspace elements plus 1-3 NEO objects in dynamic motion 

in proximity to elements 

• Absolute position required for navigation: < 0.4 m. Relative position required for navigation: < 0.4 m

• This requires space-qualified clocks that are 10 to 100

times more stable than existing space qualified clock.

(Element timekeeping accuracy required: milliseconds to 

nanoseconds depending on mission)

C.5.4 Quad Function Hybrid RF/Optical Comm, Optical 

Ranging, RF Imaging System

Communications (OCT TA 5.5)

Description

• This technology provides the capability to perform four 

functions with a single system: RF and optical communication, optical ranging and RF imaging. This enables: 

○ Reduced avionics size, weight and power burden to 

Elements through combined RF/Optical capability in 

a single system. 

○ Multiple elements to aggregate communications 

through a single element to solve spectrum and ‘multiple spacecraft located in the same aperture’ issues 

on the Earth side.

○ Reliable high data rate communications between inspace elements and ground regardless of distance 
NASA/TM—2012-217670 22

from Earth and availability of assets on the groundside, to conserve element power whenever possible, 

○ Simplified tracking of terminal by providing simultaneous RF beacon capability with terminal while optical system is operating.

• This is a recommended technology for missions where 

both imaging and long-range, high rate communications 

are required for the mission. 

Performance characteristics 

• Power savings during optical mode: < TBD W. Size and 

weight reduction compared to dual systems: <40 percent

• Optical data rate to 0.5 AU from Earth: >1 Gb/s simultaneous uplink and downlink with ground

• NEO’s/NEA’s at 0.5 AU distance or greater, including 

Mars missions

C.6 TA 06—Human Health, Life Support, and 

Habitation Systems

Improvements in reliability, maintainability, reduced mass 

and volume, advancements in biomedical counter-measures, 

and self-sufficiency with minimal logistics needs are essential 

for long duration spaceflight missions. In addition, advancements in space radiation research are required, including advanced detection and shielding technologies.

C.6.1 Closed-Loop, High Reliability, Life Support 

Systems

Life Support (OCT TA 6.1)

Description 

• Enhance and develop new, flexible Environmental Control and Life Support (ECLS) process technologies and 

systems to reliably increase system closure and reduce 

logistics, enabling autonomous long duration human exploration missions.

• Based on systems analysis and trade studies, targeted 

functions and technologies may include:

○ Close the Atmosphere Revitalization (AR) loop by 

furthering O2 recovery, and reducing logistics. Technologies may include Bosch, methane processing, 

and solid oxide electrolysis as well as advanced trace 

contaminant control and filtration.

○ Further closure of the Water Recovery (WR) loop by 

processing brines. Reduce clothing logistics and enhance crew health by enabling water recovery from 

laundry and hygiene wastewaters, respectively. May 

also include purification of water derived from ISRU 

sources. 

○ Processing of solid waste to recover water, reduce 

volume, and stabilize for long-term storage. Technologies include compaction, drying and mineralization 

of solid wastes, including trash, feces and solid byproducts from AR and WR processes.

○ Opportunities to develop common technologies, processes, and components suitable for multiple vehicle 

and mission applications can enhance the overall sustainability of human space exploration.

• Bring technologies to TRL 6 through progressive levels 

of ground-based integrated testing and ISS flight demonstrations. Perform long duration human in the loop testing to flush out hardware closed-loop issues such as 

contaminant buildup.

• NOTE: “High Reliability Life Support Systems” is a 

subset of this technology item. 

Performance Characteristics 

• Approach 100 percent closure for water and oxygen. Enable vehicle and mission autonomy through high reliability, significantly reduced consumable mass, and reduced 

dependency on logistics. 

Meet new vehicle requirements including operation in 

more extreme cabin environments (reduced pressure 

[8 psia] and elevated O2 [≈32 percent]), reclamation of 

more complex process streams, and planetary protection.

High Reliability Life Support Systems

Life Support (OCT TA 6.1)

Description 

• Development and validation of open and closed-loop 

Environmental Control and Life Support Systems 

(ECLSS), including Atmosphere Revitalization, Water 

Recovery, Waste Management and Crew Accommodations, focused at improving reliability and reducing logistics over the SOA. 

• Base technology selection and development on systems 

analysis and trade studies. Deliver new gap-filling technologies identified by vehicle elements including common adjustable pressure regulator capable of controlling 

a range of cabin, suit loop, and EVA suit pressures, low 

maintenance human waste collector and trash compactor, 

clothing, washer and dryer. 

• Bring technologies to TRL 6 through progressive levels of 

ground-based integrated testing and targeted flight demonstrations for selected process technologies. Perform long 

duration testing to address hardware reliability issues.

• Opportunities to develop common technologies, processes, and components suitable for multiple vehicle and 

mission applications can enhance the overall sustainability of human space exploration

Performance Characteristics 

• Meet or exceed performance over current state of the 

practice (≈90 percent recovery of water from urine and

humidity condensate, and ≈50 percent of O2 from CO2).

• Meet new vehicle element requirements: 

○ More robust and reliable common components (e.g., 

fans, separators, pumps, sensors) to support longer 

(unmanned) loiter and extended mission durations that 

withstand the launch/landing loads environments and 

thermal/dust environments. 
NASA/TM—2012-217670 23

○ Increased vehicle autonomy, including high reliability, reduced logistics and in-flight reparability; 

○ More extreme cabin environmental conditions (reduced 

pressure [8 psia] and elevated O2 [≈32 percent]) 

○ More complex process streams for recycling 

(wastewater from trash, hygiene and laundry).

C.6.2 Deep Space Suit (Block 1)

EVA (OCT TA 6.2) 

Description

• EVA suit with rear entry capability and crew-cabin pressure matching for compatibility with Suit Port; improved 

life support systems for increased life, reliability, and 

flexibility; and improved power-avionics-software to increase crew autonomy and work efficiency.

Performance characteristics

• Suit—rear entry suit, capable of operations at ~8 psid 

(SOA is 4.3 psid)

• DSH needs: Dexterous gloves for IVA contingency 

repairs while the cabin is depressurized. 

○ Experience shows that EVA repair inside a cabin is 

not practical (suits are too bulky), but IVA suited 

repair may be possible, if gloves are flexible enough 

for fine motor skill work. 

• Portable Life Support System (PLSS)

○ Variable set point oxygen regulator provides more 

flexibility for interfacing with multiple vehicles, the 

ability to start an EVA at an 8 psid pressure driven by 

a suit port and then decrease pressure mid-EVA for 

improved mobility, and treat decompression sickness 

in the suit (variable between 0 and 9 psid)

○ On-back regenerable CO2 and humidity control 

(eliminates consumables)

○ Robust water loop that can handle low quality water, 

long duration missions, low pressure operations, and 

bubbles (> 50 EVA life)

• Power-Avionics-Software (PAS)

○ Compatible with high specific energy battery 

(> 235 kW-hr/kg)

○ Radio that is network capable for missions involving 

multiple assets (vehicles and suits) and has data rates 

that support transmitting high definition video 

(> 10 Mbps)

○ EVA display (either helmet mounted or handheld) 

that improves upon the 12 character LCD and laminated flip cards used on ISS

○ EVA information system that increases crew autonomy and work efficiency

C.6.3 Lunar Surface Space Suit (Block 2)

EVA (OCT TA 6.2) 

Description 

• Suit Port-compatible EVA suit for surface destinations 

with small gravity field and hard vacuum atmosphere 

(e.g., Lunar surface)

Performance characteristics

• Assumptions:

○ Block 2 development occurs after Block 1 (deep 

space suit). Block 1 development is successful and 

technologies can be transferred to Block 2 as appropriate

○ Pressurized rover concept of operations with suit port

○ Lunar surface or other mission with small gravity 

field and hard vacuum atmosphere

– For example, a Mars mission with 1/3 g and low 

pressure CO2 atmosphere would require additional development due to environmental constraints

• Technical changes from Block 1 to Block 2

○ Suit: improved lower torso mobility

○ Portable Life Support System (PLSS): upgrade to 

dust tolerant components (quick disconnects, relief 

valves, etc.) 

○ Power-Avionics-Software: upgrade to dust tolerant 

electrical connectors, switches, and controls; increase 

the capabilities of the information system for additional autonomy; take advantage of advances in battery or avionics components as appropriate

C.6.4 Mars Surface Space Suit (Block 3)

EVA (OCT TA 6.2) 

Description 

• Suit Port-compatible EVA suit for surface destinations 

with intermediate gravity field (1/3 g) and low pressure 

atmosphere (Mars)

Performance characteristics

• Assumes Block 3 development occurs after Block 1 

(deep space suit) and Block 2 (surface suit for moons). 

• Technical changes from Block 2 to Block 3

○ All EVA systems components have an increased need 

for decreased mass

○ Suit: additional emphasis on boots, thermal insulation 

for CO2 atmosphere

○ Portable Life Support System (PLSS): Evaluate existing technologies for use in CO2 atmosphere, may 

need to develop a new PLSS schematic

○ Power-Avionics-Software: increase the capabilities of 

the information system for additional autonomy (even 

bigger time delay); take advantage of advances in 

battery or avionics components as appropriate

C.6.5 Long Duration Spaceflight Medical Care

Life Sciences/HRP (OCT TA 6.3)

Description

• Strong evidence from spaceflight and analogs indicate 

that medical conditions of different complexity, severity, 

and emergency will inevitably occur during long-term 

Exploration missions. Long duration missions (>1 yr) increase the risk of serious medical conditions due to lim-
NASA/TM—2012-217670 24

ited options for return to Earth, no resupply, highly limited mass, volume and some communication delays. 

Plans for medical care consider the most likely medical 

conditions, their operational and health consequences 

and the resources needed for treatment. Plans for the 

medical system seek to minimize the probability of mission failure or loss of crew.

• HRP’s Integrated Medical Model (IMM) simulates medical events during space flight missions and estimates the 

impact of these events on crew health and mission success. A three-crew, 386 d, asteroid mission simulation 

with 28, 2-crew EVAs suggests an optimized medical kit 

having a mass of 62 kg and a volume of 0.15 m3

. (These 

figures do not include all of the medical equipment needed for diagnosis). 

• The medical system must monitor and treat crewmembers 

during the mission. The requirements for the medical system are impacted by mission duration; number of EVAs; 

age and gender of the crew; and crew medical expertise

• The return of biological samples is required to assess 

human system response to the mission in order to efficiently mitigate risks in future missions.

• Technologies will be tested on ISS and in flight analog 

environments

Performance characteristics

• Rapidly evolving technologies in this area will be developed to help select and prepare crew and optimize care 

during the mission.

• Platforms that integrate multiple diagnostic and therapeutic smart medical devices, focusing on early detection 

and intervention of high-consequence and remediable 

conditions, with consideration for dual-use technologies. 

Capabilities include: diagnostic imaging, oxygen concentrator, ventilator, laboratory analysis (saliva, blood, 

urine), bone fracture stabilization and healing, medical 

suction, rapid vascular access, dental care, kidney stone 

diagnosis and treatment, IV solution preparation and delivery, medical consumables inventory tracking, and 

medical data management.

C.6.6 Long-Duration Spaceflight Behavioral Health and 

Performance

Life Sciences/HRP (OCT TA 6.3)

Description

• Behavioral health and interpersonal relations among crewmembers are critical to the success of long duration exploration missions in isolated, confined and extreme 

environments. Technologies are required for crew selection 

and composition, training, support, monitoring, and intervention.

Performance characteristics

• The habitable volume must be large enough and laid out 

to execute the necessary tasks and to provide a psychologically acceptable space for the long period of confinement.

• Sensory stimulation (e.g., variable lighting, virtual reality) must be augmented to offset the physically and socially monotonous environment.

• Cognitive performance deficits, stress, fatigue, anxiety, 

depression, behavioral health, task performance, teamwork, and psychosocial performance must be unobtrusively monitored.

• Devices must mitigate the effects of fatigue, circadian 

misalignment, work-overload.

• Communication tools must offset communication delays 

ranging from seconds to minutes. 

C.6.7 Microgravity Biomedical Counter-Measures for 

Long Duration Spaceflight

Life Sciences/HRP (OCT TA 6.3)

Description

• Prolonged exposure to weightlessness deconditions bone, 

muscle, and the cardiovascular system. Other physiological systems (e.g., sensorimotor and immune) are also altered. These changes may cause decrements in both 

health and performance. Countermeasures must mitigate 

these changes with limited resources (mass, power, volume).

• A recently discovered health risk, On-Orbit Intracranial 

Hypertension, would limit missions to 6 months or less. 

20 percent of long duration ISS crewmembers have experienced clinical symptoms; some of these changes were 

temporary and others have been, to date, permanent. 

Performance characteristics

• Assess sensorimotor function within 20 min with a portable hand-held device that also provides rehabilitation.

• Integrate multiple diagnostic and therapeutic smart medical devices, focusing on early detection and intervention 

of high-consequence and remediable conditions, with 

one platform. 

• Non-invasively measure intracranial pressure

• Worst case solution for On-Orbit Intracranial Hypertension: Artificial gravity would be required. 

C.6.8 Microgravity Biomedical Counter-Measures—

Optimized Exercise Equipment

Life Sciences/HRP (OCT TA 6.3)

Description

• Exercise equipment is necessary to address muscle atrophy, cardiovascular atrophy, and bone loss associated 

with long-duration missions in the weightless environment of space.

• Current ISS exercise equipment is too large and heavy to 

be used on a long duration missions (~1 yr duration): the 

latest equipment deployed on ISS occupies 3 International Standard Payload Racks. 
NASA/TM—2012-217670 25

Performance characteristics 

• Provide integrated aerobic and resistive exercises with a 

device no larger than 45- by 25- by 25-cm, with a mass 

of no more than 5.4 kg, requiring no external power, and 

accommodating a range of motion of at least 1 m. 

• Assess the quantity and quality of bone and muscle at 

multiple times over the course of a long-duration space 

mission. 

C.6.9 Deep Space Mission Human Factors and

Habitability

Life Sciences/HRP (OCT TA 6.3, 6.1) 

Description

• Human factors technologies are required in design and 

operations planning to ensure adequate human performance, reduce likelihood of human errors, and increase

mission safety. 

• Technologies are required in the habitable volumes (e.g., 

suit, capsule, habitat, exploration vehicle, lander) to provide an adequate food system, and to meet human environmental standards for air, water, and surface 

contamination. 

Performance characteristics

• Onboard decision support tools assist crew with real-time 

detection and diagnosis of vehicle and habitat operational 

anomalies 

• In-situ capability to assist the crew with contingency 

mission planning and development and execution of contingency operational procedures. 

• Ground-based decision support tools assist crew with 

mission operational anomalies with stale telemetry and 

operationally significant communications delays. 

• Reduce food packaging volume (30 percent) and mass 

(34 percent) so that supplies for one crew member for 1 yr

require 440 kg and 1.2 m3 consistent with food shelf-life 

requirements, especially for long duration missions. 

• An EVA suit injury countermeasures garment protects 

against injury caused by hard points in the suit and minimize movement of the crewmember within the volume of 

the suit. The garment protects the arms, legs, and torso.

• The EVA suit supports delivery of nutrition and medication to suited crew

• Microbial and chemical contamination are identified and 

measured in real-time with minimal resupply

C.6.10 In-Flight Environmental Monitoring

Life Support (OCT TA 6.4, 11)

Description 

• Extended duration missions from beyond low Earth orbit 

will require autonomous capabilities for environmental 

monitoring to assess the habitation environment and recycled life support consumables and to enable the crew 

to anticipate, react, and mitigate any risks to continued 

human occupancy. 

Performance characteristics 

• In-flight analysis capabilities are necessary—Returning 

samples to Earth for ground analysis will not be feasible 

for future missions. Environmental habitat problems on 

ISS are solved by sending air and water samples to Earth

for lab analysis, which yields data for diagnosing the 

problems. 

• Rapid detection of hazardous environmental events must 

be monitored and controlled with high accuracy. Chemical (whether predicted or not) hazards are highest in urgency, followed by microbiological threats, based on 

rapidity of impact. 

• Detect contaminants introduced via surface activities 

(dust, etc.) and of importance to planetary protection.

• Air Monitoring is well developed the system size should 

be reduced. Some specific tests for chemicals in water 

and for microorganisms have been flown, but analysis 

needs must be specified and developed. 

C.6.11 Fire Prevention, Detection and Suppression

(Reduced Pressure) 

Life Support (OCT TA 6.4)

Description 

• For longer duration missions, the habitable atmosphere 

will likely be at a lower pressure and higher percent O2

than on STS or ISS increasing the risk of fire. Small crew 

cabins (e.g., MPCV, SEV) preclude use of some of the 

current countermeasures such as the Portable Breathing 

Apparatus. Even with larger cabins (e.g., DSH, Surface 

Elements), immediate evacuation to Earth is not an 

available option and the crew is more dependent on replenishment of a fire protection capability than resupply. 

Performance characteristics 

• The crew is best protected from a fire hazard by an integrated fire protection strategy including:

○ Accurate definition of the risk from flammable materials in low-g. 

– Identify material flammability limits in low-g 

ambient environment. 

– Develop/validate non-flammable materials for 

conditions of use. 

• Early fire detection from structurally integrated distributed sensors. 

• Emergency breathing apparatus with filtering respirator. 

• ECLS-compatible and re-chargeable fire extinguisher. 

• ECLS-compatible emergency air purifier

• Contingency air monitor for relevant chemical markers 

of post-fire cleanup.
NASA/TM—2012-217670 26

C.6.12 Space Radiation Protection—Galactic Cosmic 

Rays (GCR)

Life Sciences/HRP (OCT TA 6.5)

Description

• Current estimates of crew risk from GCR radiation exposure with long duration (~>1 yr) missions beyond LEO 

exceed the NASA acceptable career standards for Risk of 

Exposure Induced Death for fatal cancers. In many cases, 

the risk estimates (Cancer Risk Projection Model currently under review with National Academy of Science) 

greatly exceed the acceptable limit. 

• Research indicates that mortality risk from radiation induced degenerative disease may further exacerbate the 

problem. GCR is difficult to shield against due to its high 

charge and energy, however shielding systems must minimize exposure levels to the maximum extent practical. 

• In addition, there are large associated uncertainties in the 

modeling of the biological damage caused by GCR. These uncertainties limit our ability to accurately evaluate 

risks and the effectiveness of biological and physical mitigation strategies. 

Performance characteristics

• Technological approaches include: risk quantification 

and uncertainty reduction through radiobiology research, 

selection of crew based on individual sensitivity for major risks, new biomedical countermeasures, cost/mass efficient multi-use shield systems, and mission planning 

away from solar minimum. 

C.6.13 Space Radiation Protection—Solar Particle Events 

(SPE)

Life Sciences/HRP (OCT TA 6.5)

Description

• Shielding from solar particle events (SPEs) is much easier than shielding from GCR and is required to mitigate 

the risk of early Acute Radiation Syndromes as well as 

increased risk of late radiation carcinogenesis. Protecting 

humans from SPEs may be a solvable problem in the 

near-term through technology maturation of identified 

shielding solutions, through design and configuration. 

However, mission operational planning has a major 

knowledge gap of forecasting the occurrence and magnitude, as well as all clear periods, of SPEs. NASA’s radiation exposure standards permit a 3 percent risk of 

radiation exposure induced death. This standard limits 

mission durations at solar minimum to 5 to 6 months for 

males and approximately 3 months for females. At solar 

maximum, the recommended limits become 154 d for 

35-yr old females to 300 d for 55-yr old males. 

• Management of the risk of exposure to SPEs requires an 

overall risk model, SPE forecasting for mission planning, 

SPE warnings and alerts to change mission planning, 

shielding options for the crew under different operational 

scenarios, in-mission dosimetry readings, and biological 

countermeasures to mitigate exposures.

Performance characteristics

• Risk projection models. 

• Forecasting/probabilistic models of events and all-clear 

periods. 

• Heliospheric environmental monitoring technology that 

provides accurate alerts for SPEs. 

• Multi-functional SPE shield systems including shelters. 

• Active miniaturized dosimetry.

• Acute biological countermeasures. 

C.6.14 Space Radiation Shielding—SPE

Life Sciences/HRP (OCT TA 6.5)

Description

• Significant risk to crew, digital equipment, and vehicle 

systems associated with Solar Particle Events (SPEs).

• Current NASA goals necessitate the need to develop radiation mitigation schema for the next generation of exploration including materials and their integration into 

vehicle architecture systems.

• Technology development is required to ensure systems 

are capable of providing structural integrity for architectural element design, while also providing sufficient radiation protection (and potentially others, e.g., thermal, 

MMOD, etc.) properties to negate any need for the addition of parasitic shield mass.

Performance characteristics 

• Candidate radiation protection materials currently exist 

but require ground and flight-testing to vet candidates 

and ensure radiation protection viability and multifunctionality. 

• NASA should invest in both ground based testing and 

flight testing to investigate the feasibility of concepts to 

protect astronaut crews from the harmful effects of radiation, both in low Earth orbit and while conducting longterm missions away from Earth.

C.7 TA 07—Human Exploration Destination 

Systems

Technology advancements with ISRU to produce fuel, O2, 

and other resources, improved mobility systems including 

surface, off-surface and Extravehicular Activity and Extravehicular Robotics, advanced habitat systems, and advancements 

in sustainability and supportability technologies.

C.7.1 Lunar ISRU: Oxygen/Water Extraction From 

Regolith

ISRU (OCT TA 7.1)

Description 

• ISRU involves the extraction and processing of local 

resources, both natural and discarded, into useful prod-
NASA/TM—2012-217670 27

ucts and services. In particular the extraction of oxygen, 

water, and other volatiles that can be used for life support, propellants, fuel cell power systems, and radiation 

protection can significantly reduce the mass, cost, and 

risk of short term and sustained human exploration of the 

Moon. Lander reusability and in-space propellant depots 

for Cis-lunar transportation are enabled. The two lunar 

ISRU products and processes that have the biggest impact on human mission architectures are:

○ Oxygen extraction from lunar regolith: This involves 

excavation of loosely consolidated surface regolith, 

regolith transfer and handling (size sorting and mineral beneficiation), and chemical/thermal processing 

to remove oxygen from mineral oxides. The Moon is 

~42 percent oxygen by mass. Operations occur in 

nominal lunar day/night cycle conditions.

○ Water and volatile extraction from lunar polar regolith:

This involves first locating and characterizing lunar 

polar ice/volatile deposits, then excavation (down to 1 

m possible), regolith transfer and handling (possibly 

crushing), heating to evolve water and volatiles, and 

volatile capture and separation. Operations may occur 

at extremely low temperatures (40 to 100 K).

Performance characteristics 

• Pilot plant to produce oxygen from lunar regolith; 250 to 

500 kg/yr; Extraction efficiency >1 percent oxygen by 

weight; Mass Payback (break-even point) is <1 yr compared to bringing oxygen from Earth, considering the 

mass of a complete ISRU system (excavator, plant, power system, and storage system).

• Full Scale plant to produce oxygen from lunar regolith:

1000 to 10,000 kg/yr (depending on crew size and propellant need); Mass Payback (break-even point) is <1 yr

compared to bringing oxygen from Earth, considering the 

mass of a complete ISRU system (excavator, plant, power system, and storage system).

• Water Extraction Plant from polar regolith: TBD. Water 

usage as well as currently unknown polar water/ice 

concentration significantly influence metrics.

C.7.2 Mars ISRU: Oxygen from Atmosphere and Water 

Extraction from Soil

ISRU (OCT TA 7.1)

Description

• ISRU involves the extraction and processing of local resources, both natural and discarded, into useful products 

and services. In particular the production of oxygen, water, and methane that can be used for life support, propellants, fuel cell power systems, and radiation protection 

can significantly reduce the mass, cost, and risk of short 

term and sustained human exploration of Mars. The two 

Mars ISRU products and processes that have the biggest 

impact on robotic sample return and human Mars mission architectures are:

• Oxygen production from Mars atmosphere CO2: This involves the collection and separation of CO2 from the 6 to 

10 torr Mars atmosphere and processing the CO2 to 

extract oxygen. Oxygen can make up >75 percent of propellant mass.

• Oxygen and fuel production from Mars soil water and atmosphere CO2: This involves excavation of Mars soil and 

processing/heating to release water. Water is electrolyzed 

to make oxygen and hydrogen (for processing). This also 

involves collection and separation of CO2 from the 6 to 

10 torr Mars atmosphere and processing with hydrogen to 

make methane (or other hydrocarbon) and water. 

Performance characteristics 

• Atmospheric CO2 processing; 3.5 kg O2/hr and 1 kg 

CH4/hr, 24 hr/d, 300 d. <7 KWe/kg O2 produced.

• Water extraction from soil: 2 kg H2O/hr, 24 hr/d, 300 d. 

~40 kg soil/hr excavation and processing. <15 KWe/kg 

water extracted. 

C.7.3 Anchoring Techniques and EVA Tools for

Micro-G Surface Operations

EVA (OCT TA 7.3)

Description

• Anchoring/mobility for a NEO mission, Exotic Geology 

Sample Acquisition, Real time Geology Sample Analysis

Performance characteristics

• Anchoring techniques for vehicles and EVA systems are 

needed for asteroid missions

○ ISS uses well defined interfaces such as hand rails as 

opposed to unknown rocky surfaces

• The ability to collect geological samples without damaging the sample (minimal heat or stress) or from a location 

with difficult access (bottom of a crater or top of a cliff) 

is needed

• Increased ability to analyze the chemical or physical 

properties of samples collected maximizes the useful data collected and minimizes the need to bring samples 

back to Earth

• All tool development must consider environmental factors and EVA compatibility (safety, mobility limitations)

C.7.4 Suit Port

EVA (OCT TA 7.3)

Description 

• A suit port provides a method of rapidly starting and ending EVAs and provides an increased level of environmental containment of potentially hazardous substances 

that could be encountered during the EVA.

Performance characteristics: 

• Reduce airlock operations time from 4 hr pre- and postEVA to 30 min
NASA/TM—2012-217670 28

• Reduce exposure of habitable volume to dust, particulates, heat transport fluids, propellants, gases such as atmospheric CO2, etc.

• Reduce consumable losses from habitable volume by 

660 kg over two weeks (assumes multiple EVAs/day)

C.7.5 Surface Mobility

Robotics and Mobility (OCT TA 7.3)

Description 

• Surface mobility systems allow for the movement of cargo, instruments and crew on the surface of an object or 

planetary body. Examples include roving, climbing, 

crawling, hopping or burrowing into the surface. Systems 

for moving cargo include prepositioning cargo for future 

human use, or repositioning payloads for re-use. Instruments can be pointed by mobility systems, or pushed into 

contact for data collection, approaching simple manipulation by using the mobility system’s transport mechanisms. Crew mobility aids expand crew range, speed and 

payload capacity while also providing power, habitation 

and environmental shelter. NASA’s experience with 

crew mobility on the lunar surface was limited to unpressurized rovers for short stays. NASA now faces new 

challenges of working on the exteriors of satellites, on 

asteroid surfaces, on planetary surfaces for long durations, or providing access to lunar craters. Complexities 

of dust management and human interaction with NEA 

during extended should also be addressed. 

Performance characteristics 

• Microgravity climbing for satellite or asteroid missions

• Precursor roving in soft/steep soils for lunar crater access

• Ballistic crater explorer, fires projectile into crater for data

• Concurrent design of crew rover and SEV for re use

• Mobile landers for repositioning spacecraft on small bodies 

C.7.6 Mission Control Automation Beyond LEO

Avionics and Software (OCT TA 7.5, 4.7) 

Description

• Support Missions beyond LEO in problem solving activities during remote or long-duration exploration missions, 

where space crew reliance on mission control is critical 

and dependent upon minimum reaction time. Advanced 

decision-support systems are needed in Mission Control 

to reduce operations costs and to maximize mission safety with Earth-based operators. 

Performance characteristics 

• Enable Earth-based nominal operation of vehicle or habitat 

at destinations with > 6 s round-trip time delay to Earth

• Enable hand-offs in Mission Ops between ground and 

crew for operations in transit and at destinations with 

> 6 s round-trip time delay

• Enable Tools to help Flight Controllers resolve off nominal situation after detection and initial response

• Enable highly efficient, small staff Earth-based Mission 

Control for Beyond LEO Crewed Missions

C.7.7 Dust Mitigation

Space Environment (OCT TA 7.5) 

Description

• Technologies are required to address adverse regolith effects in order to reduce life cycle cost and risk, and increase the probability of mission success. Based on 

Apollo lunar surface experience, there is a risk of regolith induced system degradation. The NEO environment 

may include suspended “clouds” of particulates, and is in 

any case an unknown. Particulate mitigation will be accomplished by:

○ Identification of NEO soil contamination issues for 

mechanisms and thermal systems.

○ Investigate specific risk mitigation technologies (e.g., 

seals) applicable to NEO missions. Develop technologies to limit regolith contamination, or mitigate its effects.

○ In a relevant environment, integrate and test mechanical component-level technologies to TRL 6. 

• NEO simulants are required to develop tools for anchoring, sample acquisition, etc., and Mars simulants are 

needed to develop ISRU technology.

• Regolith dust self-cleaning radiators needed for surface 

operations.

• Dust tolerant components or self-cleaning capability is 

needed for Lunar Surface Space Suits (Block 2).

• Active dust removal technology (SPARCLED) can also 

be used to acquire small-sized samples from NEOs or 

dust-sized samples from reduced-gravity bodies.

Performance characteristics

• Mitigation technologies must:

○ Maintain the solar absorptivity of a dust contaminated radiator surface within +20 percent of the pristine 

surface value, 

and 

○ Provide negligible dynamic seal wear to 2 million cycles (approx. 6 month life) or 20 million cycles for a 

5 yr life.

C.8 TA 08—Science Instruments, Observatories 

and Sensors Systems 

Not applicable for this report.

C.9 TA 09—Entry, Descent, and Landing Systems

Human-class capabilities for Mars entry, descent, and landing; technologies advancing low mass high velocity Thermal 

Protection Systems (TPS), atmospheric drag devices, deepthrottling engines, landing gear, advanced sensing, aero-
NASA/TM—2012-217670 29

breaking, aero-capture, etc. Soft precision landing capability is 

also needed, e.g., for Moon and NEA’s.

C.9.1 Entry, Descent and Landing (EDL) Technologies—

Mars Exploration Class Missions

EDL (OCT TA 9.1, 9.2, 9.3, 9.4) 

Description

• EDL systems for Mars exploration-class missions require 

large surface payloads. This technology enables reliable 

and safe delivery of multiple 40 mt payloads to the surface of Mars in order to support human exploration. The 

benefits of focused EDL technology activities include: 

increased mass delivery to a planet surface (or deployment altitude), increased planet surface access (both 

higher elevation and latitudes), increased delivery precision to the planet’s surface, increased robustness of landing system to surface hazards, and enhanced safety and 

probability of mission success for EDL phases of atmospheric flight.

Performance characteristics

• Aeroassist, Aerocapture, and Entry—AAES are defined 

as the intra-atmospheric technologies that decelerate a 

spacecraft from hyperbolic arrival through the hypersonic phase of entry. Options include deployable, inflatable, 

and mid-L/D vehicles, which need to be actively guided 

to limit loads and achieve accurate landings. 

• Descent—These technology advancements primarily focus on providing greater deceleration in the supersonic 

and subsonic regimes in a manner that does not reduce 

landing accuracy or result in transient unsteadiness or 

loss of performance in the transonic regime. For humanclass missions, inflatable and retropropulsion technologies are options.

• Landing—The key areas of technology development are 

the systems to sense the surface and avoid hazards, descent propulsion motors and plume-surface interaction 

mitigation, touchdown systems, high-G survivable systems, and small-body guidance. Landed payloads include: Large Robotic Landers (100 to 1500 kg) and 

Human Class (1500 to 45000 kg) 

• Vehicle Systems—EDL systems are by their nature an 

integrated framework of technologies that necessitate 

system level validation for robust maturation.

• Modeling and simulation along with atmospheric and 

surface characterization activities are essential for advancing these technologies.

C.9.2 EDL Technologies—Earth Return

EDL (OCT TA 9.1, 9.2, 9.3)

Description

• Earth Return entry, descent and landing systems for Human exploration architecture missions include highvelocity (8 to 14 km/s) Earth entries from beyond LEO—

from HEO, NEAs, libration points, the Moon, and Mars.

This technology enables reliable and safe return of crew 

and/or logistics, and may have reusability requirements.

The benefits of focused Earth return technology activities 

include: human safety during return from missions beyond LEO, lower-mass return capsules, increased landing system robustness, enhanced safety and probability 

of mission success, architecture flexibility and element 

reusability, and for robotic missions, sample return reliability and planetary protection. Technology developments must begin immediately in order to enable early 

exploration architectures. Extensive ground testing and 

flight tests in Earth’s atmosphere will be necessary to 

meet reliability requirements.

Performance characteristics 

• Aeroassist, Aerocapture, and Entry- AAES are defined as 

the intra-atmospheric technologies that decelerate a spacecraft from hyperbolic arrival through the hypersonic phase 

of entry. Ablative materials are an enabling technology 

needed for high velocity entries (>8 km/s, up to 16 km/s 

for robotic comet sample return missions), possible 

aerocaptures for reusability or skip entries for downrange 

capability, and possible dual heat pulse entries. Keys are 

low-cost, high-reliability manufacturing and subsystem/system performance modeling and validation.

• Descent—At Earth, these are usually parachutes; systems 

for this flight regime could have increased requirements 

due to higher entry velocities. For sample return capsules, 

inherently stable vehicles without parachutes are preferred 

to meet the reliability requirements for minimal mass.

• Landing—The key area of technology development is 

the impact attenuation system; some large-system progress has been made through Orion (sample return capsules will likely have different requirements) 

• Vehicle Systems—EDL systems are by their nature an 

integrated framework of technologies that necessitate 

system level validation for robust maturation.

• Modeling and simulation are essential for quantifying the 

reliability of these systems.

C.10 TA 10—Nanotechnology 

Not applicable for this report.

C.11 TA 11—Modeling, Simulation, Information 

Technology and Processing 

Advancements in technologies associated with flight and

ground computing, integrated s/w and h/w modeling systems, 

simulation and information processing.

C.11.1 Advanced Software Development/Tools

Avionics and Software (OCT TA 11.2)

Description

• Reliable software engineering tools and technologies to 

ensure system reliability and reduce software costs (and 

hence system and mission costs). 
NASA/TM—2012-217670 30

Performance Characteristics

• Increase software design productivity and reduce lifecycle software DDT&E and maintenance costs, greatly 

lowering cost/SLOC (source line of code)

○ Qualification of model-based software development 

methods 

○ Dynamic certification/recertification of software developed through model-based and other highly automated methods

○ Software system infrastructure to leverage multi-core 

avionics

○ Reusable software platforms suitable for human-rated 

spaceflight

• Ensure on-board software reliability for long-duration 

human missions with light-time delay 

• Enable verification of advanced software-based functions 

for: crew autonomy, autonomous systems, vehicle systems health management, and situational awareness capabilities

C.12 TA 12—Materials, Structures, Mechanical 

Systems and Manufacturing

Technology advancements for lightweight structures providing radiation protection, multifunctional structural design and 

innovative manufacturing. In addition, new technologies associated with reducing design, manufacturing, certification and 

life-cycle costs.

C.12.1 Structures and Materials for Inflatable Modules

Structures/Materials (OCT TA 12.1, 12.2, 7.4.2) 

Description 

• The primary advantage of inflatable/expandable structures is the readily collapsible walls that reduce stowage 

volume for the launch package, but provide extra volume 

for living space when expanded. The resulting mass-tovolume ratio for expandable structures can be lower than 

that for conventional hard shell structures.

• The objective is to develop expandable structures technology for application as pressurized elements such as 

crew habitats, logistics add-ons, and airlocks. The goal is 

to develop expandable technology for increased deployed-habitable volume for minimal packing volume, 

with improved confidence in structural and thermal performance in the space environment.

Performance characteristics 

• Long-term creep performance characterization of the 

structural shell of the inflatable module 

• Inflatable Structure Restraint Layer damage tolerance 

(predictive modeling validated with testing). 

• Multi-layer insulation performance degradation prediction after folding/deployment (predictive modeling validated with testing). 

• Bladder material selection. 

• Bladder-to-metal interface seal. 

• Predictive modeling of deployment dynamics. 

C.12.2 Lightweight and Efficient Structures and 

Materials

Structures/Materials (OCT TA 12.1, 12.2)

Description

• Efficient Structures and Materials that demonstrate significant weight and cost savings for aerospace applications to 

provide a total systems based efficiency. This includes 

multifunctional, lightweight and robust (i.e., inspectable, 

repairable, damage tolerant, etc.) structures and materials 

specifically tailored for mission applications. 

• Emerging Innovations in Manufacturing Technology that 

offer significant improvement over SOA, critical to 

achieving the destination, performance, and affordability 

objectives for exploration

• Design and Certification Methods to ensure timely introduction of advanced, multifunctional structures and materials into future reliable space systems 

○ Damage models for reliability (certification and sustainment) 

○ Optimized analysis and test for verification and validation

○ Streamlined Design-Analysis-Certification processes

○ Rapid material properties development

Performance Characteristics

• Lightweight structures and materials optimization to realize structural system dry mass savings (minimum of 20 

to 25 percent) and operational cost savings. 

• Multifunctional structures that offer improvements in radiation protection, MMOD shielding, thermal management, structural health management, and system damping 

benefits over conventional structures. Includes composite 

and metallic materials.

C.12.3 Mechanisms for Long Duration, Deep Space

Missions 

Mechanical Systems (OCT TA 12.3) 

Description

• Recent high impact, infant mortality and pre-mature 

hardware failures aboard the ISS (e.g., SARJ, Urine Processor bearings, Ammonia cooling pump, Canada Arm 

LEE, etc.) accentuate the need for tribological and mechanical component innovations to enable future HSF 

missions. Reliable, long-life, mission critical systems 

such as cooling pumps, circulators and components for 

Zero Boil-Off systems, control moment gyros, robotic 

manipulation hardware, docking/hatch devices and pointing mechanisms must be more resilient and capable than 

current COTS technology allows. New lubricants, bearing and gear materials and designs are needed to ensure 

mission success. 
NASA/TM—2012-217670 31

• Emerging lightweight superelastic materials (Nitinol alloys), advanced lubricants (ionic fluids), and novel 

mechanism designs (low sliding high contact ratio gears) 

are poised to help avoid mission ending/crippling mechanism failures but must be matured. Such innovations 

will enable silent, ultra-reliable spacecraft systems such 

as cabin blower motors and fans, thermal management 

pumps, etc. Innovative power transfer technologies 

(magnetic gears) can significantly reduce cabin noise 

levels enhancing astronaut health and operational efficiency over long duration missions.

Performance Characteristics

• Mission critical systems (e.g., cooling pumps, circulators, control moment gyros): 

○ Current SOA: <10yr, sustain 6 g loads (designs must 

be 2X mission life and 2X Shuttle launch load)

○ Goal: >10 yr at + or –50 °C from operating temperature sustaining 10 g loads (2X mission life, 2X launch 

load of 5g’s)

• Bearing and Gear Materials to handle higher loads:

○ Current SOA: steel 

○ Goal: 15 percent weight reduction with comparable 

capability (superelastic materials)

C.12.4 Low Temperature Mechanisms

Mechanical Systems (OCT TA 12.3) 

Description 

• Future deep space missions will demand of the mechanical 

system technologies both safety and reliability over long 

durations, and in extremely challenging environments such 

as cryogenic temperatures and ultra-high vacuum. Long 

life, cryogenic actuators are a key technology challenge, 

and enabling for outer planet and deep space probe missions. Long-life-by-design, modular (for ease of integration) actuators consisting of motors, gearboxes, 

position/speed sensors, and motor controller electronics 

will need to be capable of operating in dusty NEO or lunar 

environments at temperatures between 400 and 40 K, for 

years, in order to meet those reliability demands.

Performance characteristics 

• Current SOA calls for heating to keep liquid lubricated 

actuators above –55 to –70 °C, with control electronics 

housed separately in a “warm electronics box” above –

55 °C. Cryogenic compatible actuator components (lubricants, bearings, gears, position sensor) and control electronics operational to –230 °C allow integration of the 

motor controller with the actuator, greatly enhancing reliability, modularity and scalability. Cryo-compatible actuators/electronics would eliminate the hardware and 

wiring for heating (with ~30 percent power savings), and 

reduce by two orders of magnitude the interconnect cables, resulting in up to 50 percent reduction in mass of 

the electronics and electronic housings.

C.13 TA 13—Ground and Launch Systems

Processing

Technologies to optimize the life-cycle operational costs, increase reliability and mission availability, improve mission 

safety, reduce mission risk, and reduce environmental impacts 

(i.e., green technologies).

C.13.1 Low Loss Cryogenic Ground Systems Storage and 

Transfer

Ground Ops/Systems (OCT TA 13.1) 

Description 

• Reducing the cost of ground operations is important to 

reduce the life cycle costs of exploration. 

• Thermal inefficiencies in cryogenic storage and transfer 

systems increase vehicle safety risk, drive costly commodity boil off losses, and limit the distance over which 

these commodities can be transferred. 

• Lower life cycle costs, increase supply capacity/launch 

availability, reduce propellant losses with zero boil off, 

zero loss chilldown, propellant conditioning/densification 

systems; reduce large volume consumption of helium 

(nonrenewable, finite resource)

• Examples include approaches for insulation systems materials/structures, active refrigeration (cryocoolers), highefficiency fluid conditioning, recovery, purification, and 

reliquefaction systems, in-situ measurements to reduce/eliminate He over-purging 

Performance characteristics 

• Percent fluid loss, SOA: 50 percent of propellants lost 

per launch/test

○ Active refrigeration systems target: zero boil-off storage in tank, reliquefaction of vapors lost in transfer 

line chilldown, recovery of losses. Goal: eliminate 

boiloff, recover 75 percent of losses. 

○ Insulation systems target is 40 percent reduction in 

convective/radiative heat transfer (storage tank materials) and 25 percent reduction in conductive heat 

transfer (load supporting insulation for supports and 

penetrations). Goal: < 10 percent losses 

• Reduce overall He usage by at least 50 percent compared 

to current Shuttle purge operations

C.13.2 Corrosion Detection and Control for Ground 

Systems

Ground Ops/Systems (OCT TA 13.2) 

Description 

• Preclude the severe degradation of structures from corrosion to lower maintenance/inspection costs, reduce corrosion-related damage/structural failures, mitigate 

environmental impacts, and prevent liberation of material 

from flame trench during launch
NASA/TM—2012-217670 32

• Examples include approaches for environmentally 

friendly corrosion protective coatings and corrosion resistant flame trench refractory materials

Performance characteristics 

• Percent of occurrences of hidden corrosion identified, 

SOA: Visual detection of rust on surfaces and/or structural failures, no detection of hidden corrosion. Goal: 

Identify 90 percent of occurrences of hidden corrosion 

with damage responsive coatings

• Number of launch exposures the refractory material can 

withstand before damage occurs and repairs are necessary, SOA: 0 launch exposures without damage. Goal: 10 

launch exposures 

C.13.3 Fault Detection, Isolation, and Recovery for 

Ground Systems

Ground Ops/Systems (OCT TA 13.3) 

Description 

• Currently, ground and launch systems design strategies 

and operations concepts depend heavily on the use of redundant systems and elimination of single points of failure to help ensure system reliability and availability

• Lack of insight into the system’s health and status requires a high level of redundancy

• Minimize the dependence on humans to isolate failures 

and diagnose problems, identify fault propagation paths 

and speed the identification and isolation of suspected or 

failed components, decrease the time to maintain, test, 

and repair systems

• Examples include approaches for functional fault modeling of the ground systems; intelligent devices that can 

self-detect and identify faults, failures or anomalous reporting; fault isolation and diagnostics to reduce troubleshooting time; automated recovery of select failures; and 

prognostics of subsystem/component performance leading to condition based maintenance

Performance characteristics 

• Preventing launch scrub costs due to technical failures. 

Goal: reduce 50 percent of technical issues using detection, isolation, problem resolution

• Calibration and maintenance of transducers. Goal: extending calibration cycles (through the use of prognostics 

and anomaly detection) ≥ 1 yr 

C.13.4 Wiring Fault Detection and Repair for Ground 

Systems

Ground Ops/Systems (OCT TA 13.3) 

Description 

• SOA wiring constructions and manufacturing have remained essentially unchanged for decades

• Wiring insulation tends to crack and fray as it ages, is susceptible to maintenance-related damage during ground 

processing

• Wiring failures have caused significant delays/costs for 

Shuttle and catastrophic system failures in other spacecraft/aircraft

• Reduce labor and ground processing costs for investigation, troubleshooting, repair of wiring failures, invasiveness, unnecessary removal of structures and equipment, 

occurrences of wire insulation damage and potential system failures 

• Reduce exploration program risk (loss of crew due to system failures) and life cycle costs (for logistics re-supply)

• Examples include approaches for in-situ monitoring/detection systems to locate wiring failures/intermittent faults and predict remaining useful life;

reconfigurable components, rerouting and self-healing 

systems to isolate and repair a wire once damaged

Performance characteristics

• Detect potential or actual wiring problems, SOA: limited 

detection capability for intermittent failures cannot be

used on ‘live” wires. Goal: Conductive polymer with 

< 50 ohms resistance; thermally stable above 260 °C; no 

delamination upon bending; near 1000 S/cm conductivity; and increases wire weight by < 10 percent with added 

detection capability. Goal: detects conductor damage that 

is intermittent on “live” wire networks.

C.14 TA 14—Thermal Management Systems

Technology advancement for cryogenic systems performance and efficiency, effective thermal control systems for 

heat acquisition/transport/rejection, and increased robustness 

and reduced maintenance for thermal protection systems.

C.14.1 In-Space Cryogenic Propellant Storage—Zero Boil 

Off (ZBO) LO2 and Reduced/Zero Boil Off (RBO)

LH2

Cryogenic Fluid Systems (OCT TA 14.1) 

Description

• Thermal control technologies to extend the in-space and 

planetary storage of cryogenic propellants include both 

passive and active thermal control technologies. The 

former reduce heat input to a propellant tank through a 

low thermal conductivity support structure and advanced 

insulation, and the latter intercept the remaining heat 

with a refrigeration technique such as employing a 

cryocooler integrated with a broad area cooling (BAC) 

shield attached to the propellant tank surface or embedded within the tank insulation. These technologies can 

significantly reduce propellant launch mass, required onorbit margins and the complexity of vehicle operations.

 
NASA/TM—2012-217670 33

○ LO2 ZBO in-space storage: uses passive thermal control technology and the integration with an active 

thermal control system such as a 90 K cryocooler 

with cooling tubes attached to the walls of the LO2

storage tank.

○ LH2 RBO in-space storage: uses passive thermal control technology integrated with an active thermal control system such as a 90 K cryocooler integrated with 

the tank support structure and a BAC shield embedded in the storage tank MLI. 

○ LH2 ZBO in-space storage: uses passive thermal control technology such as a cryocooler with two-stages 

of cooling. The first stage of the cryocooler operating 

at 90 K and integrated with the tank support structure 

and a BAC shield embedded in the storage tank MLI 

and the second stage operating at 20 K and integrated 

with cooling tubes attached to the walls of the LH2

storage tank.

○ These technologies are directly applicable to LO2/ 

LH2 and LO2/CH4 propellant systems.

In-Space Performance Characteristics

• LO2 Storage: Less than 8.0 W of active storage system 

power per watt of heat removal at 90 K; ZBO for >400 d 

• H2 Storage: Less than 120 W of active storage system 

power per watt of heat removal at 20 K; ZBO for >400 d 

• Cryocooler mass must be less than mass of propellant 

saved.

C.14.2 Thermal Control

Thermal Systems (OCT TA 14.2) 

Description

• All future vehicles (both crewed and uncrewed) will 

require thermal control systems (TCS)

• Improve thermal control system performance and reliability to reduce mass transportation requirements and enable performance over a wide range of mission 

requirements.

• Thermal control in day/night with dust mitigation on radiators is critical for continuous ops and survival.

• Technologies that will be required include:

○ TCS fluids and variable heat rejection radiators enabling single-loop TCS architecture

○ Low mass/volume heat exchangers and coldplates

○ Advanced Supplemental Heat Rejection Devices including evaporative heat sinks and fusible heat sinks

○ Solid state devices (thermal electrics) and thermal 

sensors/health monitoring

○ Operations in Lunar Perennially Shadowed Regions 

at Cryogenic Temperatures (40 K)

Performance Characteristics

• Capable of maintaining system setpoint for large turndown ratio requirements (12 to 1 kW)

○ Exacerbated by low load in cold environment (~0 K) 

and high load in hot environment (~220 K)

• Capable of efficient operation in rapidly changing thermal 

environments and/or transient heat rejection requirements

• Reduces component and system mass 

C.14.3 Robust Ablative Heat Shield (Beyond Lunar 

Return Conditions)—Thermal Protection System

Structures and Materials (OCT TA 14.3)

Description

• A robust, scalable heat shield TPS architecture is 

required that can be used for multiple missions. Ablative 

TPS solution for primary MPCV heat shield protection 

for beyond Lunar return conditions. Improve human 

safety by detecting critical issues with MPCV TPS or 

structure prior to entry. 

Performance characteristics

• Ablative TPS Solution for primary CTV heat shield capable of withstanding ~2500 W/cm2 under 0.8 atm pressure

• Peak heat rate dominated (~90 percent) by shock layer 

radiation

• Technology needs to enter DDT&E cycle including TPS 

development, aerothermal and shock layer radiation 

modeling validation, reliability/margin quantification 

methodology, integrated system health monitoring, and 

hyperthermal ground test capability to approximate convective-radiative environment.

C.14.4 Robust Ablative Heat Shield (Lunar Return 

Conditions)—Thermal Protection System

Structures and Materials (OCT TA 14.3)

Description

• A robust, scalable heat shield TPS architecture is required that can be used for multiple missions. Ablative 

TPS solution for primary MPCV heat shield protection.

Improve human safety by detecting critical issues with 

MPCV TPS or structure prior to entry. 

Performance Characteristic

• Capable of withstanding ~1000 W/cm2 (about 33 percent

radiation) and ~1 atm pressure
NASA/TM—2012-217670 34

References

1. NASA Exploration Systems Mission Directorate presentation, “Human Space Exploration Summary,” January 

2011. http://www.nasa.gov/pdf/525162main_HEFT_Final_ 

Brief_508_20110309.pdf [accessed 27 April 2012].

2. Olson, J. “The Way Forward—New Pathways for Human 

Spaceflight,” April 2011. http://www.nasa.gov/pdf/ 

552846main_New_Pathways_Human_Spaceflight_Olson.p

df [accessed 27 April 2012].

3. Culbert, C. “Human Space Flight Architecture Team 

(HAT) Overview,” November 2011. http://www.nasa.gov/ 

pdf/603232main_Culbert-HAT%20Overview%20for% 

20GER%20Workshop.pdf [accessed 27 April 2012].

4. Olson, J.; Culbert, C.; Laurini, K.C. “NASA’s Human 

Space Exploration Plans and Architecture,” 62nd International Astronautical Congress, Cape Town, SA. October 

3-7 2011. IAC-11.E3.2.1 http://www.iafastro.net/ 

iac/archive/browse/IAC-11/E3/2/11969/ [accessed 27 

April 2012].

5. Mueller, R.P., Connolly, J.C., Whitley, R.J., “NASA 

Human Spaceflight Architecture Team: Lunar Surface 

Exploration Strategies,” Global Space Exploration Conference, Washington DC, May 2012. GLEX2012.02.P.17.x12620

6. Hoffman, Stephen J., Baker, John D., and Voels, Stephen 

A., “Mars as a Destination in a Capability- Driven 

Framework”, ASCE Earth and Space 2012, Pasadena, 

CA, April 2012

7. Friedensen, Victoria, and Mazanek, Daniel D., “NASA’s 

Plans for Human Exploration of Near-Earth Asteroids”, 

ASCE Earth and Space 2012, Pasadena, CA, April 2012

8. Bobskill, Marianne R., and Lupisella, Mark L., “The 

Role of Cis-Lunar Space in Future Global Space Exploration,” Global Space Exploration Conference, Washington 

DC, May 2012. GLEX-2012.05.5.4x12270

9. NASA Office of the Chief Technologist, “Draft Space 

Technology Roadmap Overview” http://www.nasa.gov/ 

pdf/501317main_STR-Overview-Final_rev3.pdf [accessed 

27 April 2012].

10. Moore, C., “Advanced Exploration Systems,” presentation to the NASA Advisory Council, March 2012. 

http://www.nasa.gov/pdf/630144main_2-

Moore_NAC%20Briefing%20AES_508.pdf [accessed 27 

April 2012].

11. Lange, C. et al., “Assessment of Technology Developments for the ISECG Global Exploration Roadmap,” 

May 2012. GLEX-2012.09.3.1x12269

12. National Research Council, “NASA Space Technology 

Roadmaps and Priorities: Restoring NASA’s Technological Edge and Paving the Way for a New Era in Space,” 

January 2012. http://www.nap.edu/catalog. 

php?record_id=13354 [accessed 27 April 2012].

13. NASA, “Human Exploration of Mars Design Reference 

Mission 5.0,” July 2009. NASA/SP—2009-566 

http://www.nasa.gov/pdf/373665main_NASA-SP-2009-

566.pdf [accessed 27 April 2012]

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As the National Aeronautics and Space Administration (NASA) prepares to extend human presence throughout the solar system, technical capabilities must be developed to 

enable long duration flights to destinations such as near Earth asteroids, Mars, and extended stays on the Moon. As part of the NASA Human Spaceflight Architecture Team, a 

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vehicle developers, and technologists was used to identify a minimum but sufficient set of technologies, noting that needs are created by specific mission architecture 

requirements, yet specific designs are enabled by technologies. Further consideration was given to the re-use of underlying technologies to cover multiple missions to 

effectively use scarce resources. This suite of critical technologies is expected to provide the needed base capability to enable a variety of possible destinations and missions. 

This paper describes the methodology used to provide an architecture-driven technology development assessment (“technology pull”), including technology advancement 

needs identified by trade studies encompassing a spectrum of flight elements and destination design reference missions.

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Technology assessment 

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