Richard C. Oeftering, Greg L. Kimnach, James Fincannon, Barbara I. Mckissock,

Patricia L. Loyselle, and Edmond Wong

Glenn Research Center, Cleveland, Ohio

Advanced Modular Power Approach to Affordable,

Supportable Space Systems

NASA/TM—2013-217813

June 2013

AIAA–2013–5253
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Richard C. Oeftering, Greg L. Kimnach, James Fincannon, Barbara I. Mckissock,

Patricia L. Loyselle, and Edmond Wong

Glenn Research Center, Cleveland, Ohio

Advanced Modular Power Approach to Affordable,

Supportable Space Systems

NASA/TM—2013-217813

June 2013

AIAA–2013–5253

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Space 2012 Conference and Exposition

sponsored by the American Institute of Aeronautics and Astronautics

Pasadena, California, September 11–13, 2012
Acknowledgments

The authors wish to thank the following for their support, insight and contribution to the paper. Robert Button of Energy Systems 

Branch for his technical guidance and paper review. Advanced Modular Power Systems (AMPS) Project Managers, David Irimies and

Ralph Jansen for project leadership. Study Team Leads, Raymond Beach, and James Soeder of Power Systems Branch for study guidance 

and power system insights. Monica Guzik, Brianne Scheidegger, Ken Burke, and Ian Jakupca for information on Non-Flow Through Fuel 

Cell technology. Bob Scheidegger of Solar Electric Propulsion, and Patrick George of Deep Space Habitat for vehicle power information.

Mike Gernhardt, Joyce Seriale-Grush of Johnson Space Center (JSC) for Multi-Mission Space Exploration Vehicle information 

and insight.

Available from

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7115 Standard Drive

Hanover, MD 21076–1320

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Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.
NASA/TM—2013-217813 1 

Advanced Modular Power Approach to Affordable, 

Supportable Space Systems 

Richard C. Oeftering, Greg L. Kimnach, James Fincannon, Barbara I. Mckissock, 

Patricia L. Loyselle, and Edmond Wong 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Recent studies of missions to the Moon, Mars and Near Earth Asteroids (NEA) indicate that these 

missions often involve several distinct separately launched vehicles that must ultimately be integrated 

together in-flight and operate as one unit. Therefore, it is important to see these vehicles as elements of a 

larger segmented spacecraft rather than separate spacecraft flying in formation. The evolution of large 

multi-vehicle exploration architecture creates the need (and opportunity) to establish a global power 

architecture that is common across all vehicles. The Advanced Exploration Systems (AES) Modular 

Power System (AMPS) project managed by NASA Glenn Research Center (GRC) is aimed at 

establishing the modular power system architecture that will enable power systems to be built from a 

common set of modular building blocks. The project is developing, demonstrating and evaluating key 

modular power technologies that are expected to minimize non-recurring development costs, reduce 

recurring integration costs, as well as, mission operational and support costs. Further, modular power is 

expected to enhance mission flexibility, vehicle reliability, scalability and overall mission supportability. 

The AMPS project not only supports multi-vehicle architectures but should enable multi-mission 

capability as well. The AMPS technology development involves near term demonstrations involving 

developmental prototype vehicles and field demonstrations. These operational demonstrations not only 

serve as a means of evaluating modular technology but also provide feedback to developers that assure 

that they progress toward truly flexible and operationally supportable modular power architecture. 

1.0 Introduction 

Under the Advanced Exploration Systems (AES) Modular Power Systems (AMPS) project lead by 

NASA Glenn Research Center (GRC) the goal is to develop a modular power architecture composed of 

technologies for power generation, energy storage, power distribution and health management that will 

reduce the cost of future space systems. 

The AMPS modular approach is intended to assure that the development and qualification investment 

is preserved in modular building blocks that have space qualified heritage while still providing 

adaptability to serve an array of future applications. This requires that the modular power development 

look across a broad array of vehicles and establish a common architecture that can be scaled to meet 

widely varying power needs. A modular approach can reduce power system integration complexity and 

costs by reducing the number of distinct integration steps and reduce the complexity of external 

equipment. Standardized modules may encapsulate a common set of reusable software, embedded 

diagnostics, health management functions and features that assure interoperability. A modular power 

system can enable the integrator to use a common set of standardized integration procedures, common set 

of tools, and a highly simplified set of test equipment. Encapsulating these capabilities in power modules 

not only streamlines integration but also benefits the supportability of operational spacecraft. 

Future exploration missions emphasize destinations well beyond low Earth orbit (LEO) involving 

long and complex operations with multiple vehicles, (Ref. 1) as shown in Figure 1. Extremely limited and 

costly launch capacity restricts logistics up-mass. Crew consumables dominate logistics mass capacity 

and the capacity for maintenance spares is extremely low. This drives the program to squeeze 
NASA/TM—2013-217813 2 

service life out of every available piece of hardware. For the first time, hardware from vehicles that have 

completed their initial role and would normally be expended, are now considered a potential source of 

spares. The AMPS modular power strategy therefore includes the ability to salvage key elements to be 

reconfigured and repurposed for secondary applications including system spares (Ref. 2). 

The following sections describe the AMPS project’s approach to overcome what have been barriers 

by developing modular power architecture, developing key modular power technologies and 

demonstrating them in user applications. AES Modular Power Systems approach can lead to a new set of 

legacy hardware that users can quickly adapt to new missions, drive down initial development cost, 

streamline prelaunch integration, reduce in-space support cost, and finally, enable the salvaging and 

repurposing of space flight hardware. In doing so, modular power not only preserves the utility and value 

of hardware, but enables the development and operation of multi-mission vehicles. 

2.0 Defining Capabilities of Modular Power 

AMPS capability objectives address the “abilities” that have long been goals of spacecraft designers. 

AMPS defines modular power features that provide commonality, but seeks to define an architecture that 

best supports the overall program by addressing scalability, flexibility, interoperability and supportability. In 

this section we define and discuss these abilities while staying alert to program and development pitfalls. 

OR 

NOTIONAL 

Figure 1.—Notional Multi-Vehicle stack for a near Earth asteroid mission. There are multiple crew vehicles and 

multiple propulsion schemes supporting asteroid missions. One mission scenario has 3 cryogenic propulsion 

stages while another has a mix of solar electric and cryogenic stages. This illustrates that there is an 

extraordinary amount of flight hardware and opportunities to exploit modular power.

Multi-Mission 

Space Exploration 

Vehicle 

Deep Space 

Habitat 

Multi-Purpose 

Crew Vehicle 

Cryo 

Propulsion 

Stage 

Solar Electric Propulsion

Cryo 

Propulsion 

Stages 

Cryo 

Propulsion 

Stage 

Solar Electric Propulsion
NASA/TM—2013-217813 3 

2.1 Commonality and Modularity Pitfalls 

Modularity cannot be developed for commonality sake alone. There are number of pitfalls to avoid. 

Lessons learned in developing a new common standard include: 

2.1.1 Poor Stakeholder Return on Investment 

 A commonality standard limited to one specific project, (thus not truly common) 

 One organization imposes its agenda on others (no mutual benefit) 

 Added development complexity without downstream benefit 

 Fixed or over constrained design that restricts options or suppresses innovation 

 The new standard discards prior investments in established systems 

 Unanticipated upfront cost of new processes, equipment, and staff retraining 

 Incomplete development with end users expected to complete the development 

2.1.2 Technical Penalties Due to Poorly Conceived Commonality Approach 

 Increased mass or complexity with no compensating performance or cost gain 

 Poor match with user needs 

 Poor scalability and inflexible 

 Partial solution, and lacking a complete overall architecture 

 Prone to obsolescence by selecting proprietary solutions without industry buy in 

2.1.3 Program Cost and Schedule Penalties Due to Poor Planning or Execution of Commonality 

 Increased development time 

 Developmental cost savings at the expense of operational and logistics costs 

 Overestimating the benefit and underestimating the development cost 

 Late introduction “the train has left the station” 

The AMPS project proceeds with consideration of these lessons learned. 

2.2 Commonality 

Spacecraft power architecture with high commonality is built upon elements that use a common 

design approach and share common features to a point that hardware is interchangeable. For AMPS this 

common design approach extends to all spacecraft within the space exploration architecture. That is, 

power hardware can be interchangeable among a group of spacecraft. Commonality allows a program to 

pursue a single development effort that is then used by multiple spacecraft. This approach can save cost 

by eliminating the duplication among individual spacecraft development efforts. In some cases new 

standards can be developed to assure that spacecraft are compatible with the common power architecture. 

Commonality should not be imposed on spacecraft developers without first establishing the common 

power system needs and accommodating the variations from spacecraft to spacecraft. Commonality alone 

is insufficient; the approach to commonality must allow for scalability, flexibility, interoperability and 

operational supportability. 

2.3 Scalable—Flexible Power 

The modular power system architecture must accommodate the need for scalable and flexible power. 

Scalable power systems allow for an incremental expansion of capability. The International Space Station 

(ISS) Electric Power System was designed to be scalable at the system level since the ISS could not be 

built in a single mission. It was however important to establish electric power early and thus power 

systems were one of the first elements to arrive. Delivering power as parallel power channels allowed this 
NASA/TM—2013-217813 4 

modest initial power capability to successfully expand over a series of missions. Scalable power 

capability will be needed as exploration missions extend beyond low Earth orbit. 

Future exploration missions are expected to be very long and involve more complex operations 

involving multiple spacecraft. Further, these missions will not have the benefit of robust logistics 

infrastructure that ISS had (Ref. 3). A new generation of multi-mission exploration vehicles will perform 

multiple sorties with diverse objectives and durations. The flexibility to scale the power system up, or 

down, also enables the crew to tailor the capabilities to best match vehicle to the mission range, payload 

and environment. The modular flexibility enables the crew to select the best power generation, energy 

storage and power distribution options. 

Scalability and flexibility is only feasible if it can be achieved without complex operations that require 

complex equipment, tools and extensive crew skills. Features that provide scalability and flexibility must be 

embedded into the design so that it does not need to be provided by additional external equipment. 

2.4 Supportability 

Supportability involves the operations, logistics and equipment required to maintain a system and 

assure a high degree of system availability (Ref. 4). The long term operational support of a space flight 

system has become an important program cost issue in an era of long lived space facilities, such as, the 

ISS. If supportability is not properly addressed, the crew and supporting ground operations may become 

burdened with system maintenance activity that competes with the science and mission objectives. 

In long missions systems wear out or degrade over time and will require maintenance and hardware 

replacement. Upmass and crew time limitations mean that supportability cannot be achieved by simply 

“loading up” on spares and maintenance equipment. Supportability involves using operating techniques 

and technology that minimize the “logistics footprint” in terms of mass and operational crew time. 

Supportability features are intended to maximize system availability while: 

 Minimizing spares mass 

 Minimizing maintenance equipment mass 

 Minimizing operational time dedicated to maintenance 

 Minimizing crew skill and training 

 Enabling the crew to effectively and independently respond to problems 

The supportability features that support these goals are also essential for effective modular commonality, 

scalability and flexibility. Supportability must be embedded into the modular design. Supportability is 

highly dependent on complete interoperability capabilities. 

2.5 Interoperability 

In general “Interoperability”, can be defined as a capability that allows multiple systems to connect 

and interact with minimum dependency on coordinating or integrating action by external systems or 

users. Ideally the connection and interaction is entirely transparent to the user (Ref. 5). When 

interoperability is extended to lower level modular hardware it enables portability and interchangeability. 

This form of interoperability is often referred to as “Plug-and-Play”. 

For a long life program, interoperability can provide adaptability to long term change. In some cases 

interoperability is achieved by encapsulating hardware behind interface converters that, in effect, provide 

a high degree of “virtual” commonality. This adaptability enables systems to operate with low cost 

“legacy” hardware (Ref. 6). This adaptability tolerates diverse system components, mitigates the impact 

of obsolescence and holds open the path to future upgrades. 
NASA/TM—2013-217813 5 

Interoperability involves a combination of hardware and software that is applied to all systems across 

multiple vehicles and permeates through all levels of the system. Interoperability is dependent on “smart” 

network capable modules connecting to “self-organizing” networks. For data and control the smart 

transducer standards, developed under IEEE 1451, provide the guidance for plug-and-play capability 

(Ref. 7). Modular power may need to establish additional standards for power interoperability. Power 

system interoperability standards may be extracted from the nation’s investment in “smart grid” 

technology. 

3.0 Potential Affordability Aspects of Modular Power 

AMPS anticipates a broad impact on program affordability by addressing non-recurring, recurring 

and long term operational costs. 

3.1 Reducing Non-Recurring Costs 

Reduction of the non-recurring development cost can be achieved by consolidating power system 

development across multiple vehicles. This is in contrast to designing and developing power system 

design unique to each vehicle. One way projects reduce development cost is to exploit existing designs 

with prior flight heritage or so called “legacy” hardware. This is rare for human spacecraft but common 

for satellites or robotic spacecraft. In fact, many NASA space science programs insist on showing flight 

heritage to reduce cost and minimize overall development and operational risk. Often the existing designs 

with heritage poorly match the new mission requirements and thus the new mission must compromise 

objectives or plan on extensive modification and requalification. 

For AMPS the strategy is to develop a set of modular elements at varied scales that can be combined 

to provide the needed capability with a minimum excess. Designers can select from an assortment of 

modular building blocks and build a capability with the fewest number of blocks. As prequalified 

modules, the designers can acquire heritage while building the power architecture that matches their 

mission needs. Rather than simply impose modular and commonality standards, this approach is expected 

to provide spacecraft developers with an early incentive to adopt standardized units because of the 

inherent reduced risk and costs. 

3.2 Reducing Recurring Costs of Integration 

The recurring costs can be reduced by building systems from relatively low cost modular blocks 

where the tooling and fabrication costs are amortized over larger number production units. In comparison 

to large monolithic designs, smaller modules using readily available components are expected to be much 

less expensive, particularly, if production supports a number of vehicles. Recurring costs can be reduced 

by extending the commonality to the set of test and integration equipment that can support multiple 

vehicles. There is further recurring cost reduction if these are supported by standardized test and 

integration procedures. 

3.3 Reducing Operational Life Cycle Costs 

Modular power is expected to reduce operational cost by extending the useful life of spacecraft, 

minimizing the mass of spares, and simplifying maintenance operations. Replaceable modules called 

Orbital Replacement Units (ORU) have been successfully used to support the International Space Station. 

Unlike ISS, however, the crew will not have easy access to Earth launched spare assemblies. All 

spares must be carried at initial launch. The ISS ORUs are relatively large and intended to support rapid 

replacement by humans or robots using simple tools and procedures. The intent was to make efficient use 

of crew time at the expense of mass. Long exploration missions cannot dedicate scarce logistics mass to 

oversized replacement units so repair must be performed at lower levels of assembly. 
NASA/TM—2013-217813 6 

AMPS is working to modularize lower levels of assembly to reduce spares mass. The embedded 

health management diagnostics and prognostics capability will reduce the crew time required to isolate a 

fault. The plug and play interoperability provides self-test features that will assist in reintegrating the new 

module. This plug and play capability in combination with commonality allows multiple vehicles for a 

given mission to be serviced by a common set of spares. Further, modular commonality enables the 

practice of salvaging hardware from other vehicles, when needed, which further reduces spares mass. 

Salvaging hardware can be used in extending the life of flight assets and enable a single vehicle to 

support multiple missions. Reusable spacecraft can potentially save billions of dollars in spacecraft value 

over the life of the program (Ref. 2). 

4.0 Modularizing the Electric Power Systems 

Space power systems are composed of power generation, energy storage and power distribution and 

management. Power generation includes solar arrays, fuel cells, and thermodynamic engines. Energy 

storage includes batteries but may include the evolving flywheel technology. Power distribution and 

management (PMAD) connects the power generation and energy storage to the user loads. It regulates the 

power and handles the delivery of power. It also provides the primary system fault detection, fault 

isolation and rerouting of power. These are typically designed as separate subsystems. The AMPS 

approach to modularity integrates certain PMAD regulator and health management functions into power 

generation and energy storage. Embedding these functions combined with plug-and-play features enables 

them to act as independent self-contained modular subsystems. This makes them more portable so they 

can be moved to different parts of the vehicle or another vehicle entirely. 

There are a number of factors that will affect the effectiveness of the modular design. The ISS is the 

best example of spacecraft power with a high degree of modularity driven by scalability and 

supportability needs. It serves as an important point of departure for modular systems for missions 

beyond low Earth orbit. The ISS design incorporated modularity at various levels of assembly. For future 

missions, modularity and commonality needs to be driven down to lower levels of assembly. 

4.1 Modularity Across Multiple Levels of Assembly 

The AMPS project recognizes that modularity is not a new concept and has been used before but 

limited to higher levels of assembly. To minimize confusion project defines modularity at various levels 

of assembly is illustrated in Figure 2. 

4.1.1 Modular Assembly 

This level of hardware is commonly known as an ORU, which is the preferred level of replacement 

for ISS. It is a convenient scale for handling by humans or by robotics. As shown in Figure 3, ORUs are 

typically encapsulated in harden enclosures with mechanical and electrical interfaces intended to be easy 

to mated or de-mated by manual or robotic means. 

4.1.2 Modular Subassembly 

For power systems this subassembly level would mean circuit card within an ORU and are typically 

replaced by a ground based NASA Depots (Ref. 4). AMPS project seeks to modularize this level so that it 

can be handled as an in-space replaceable unit. There is a significant mass advantage as shown in 

Figure 3(b) (Ref. 8). ISS vintage systems were built around the parallel bus architectures and bus 

backplane drove host chassis size and, in turn, module volume and mass. Subassembly Modules can be 

packaged to be easy to replace with simple tools. This level is where significant commonality is found 

and also where elements are combined to configure the assembly level. 
NASA/TM—2013-217813 7 

(a) (b) 

Figure 3.—ISS Assembly Level ORU. (a) The photos illustrate the common features of the Remote Power 

Controller Module (RPCM) units. The left RPCM is externally mounted and replaced by robotic means. The right 

unit is internally mounted and replaced manually. (NASA Photo), (b) The breakdown of an ORU often reveals 

that the chassis is the largest contributor to module mass and replacement of individual subassemblies is much 

more efficient (Ref. 8).

Modular Vehicle

Modular System

Altair, ISS, Multi-Mission 

Exploration Vehicle

ISS Power Channel

Modular Subsystem ISS PV Module

Modular Assembly 

(Orbital Replacement Unit) ISS DDCU, SSU, BCDU, MBSU, RPCM etc.

Modular Subassembly ISS Option C, SBCU, FRPCM, etc.

EEE Parts

Modular Component ISS Option C, 120/28 Vdc Converter, 120/28 Vdc Switch

Modular Level Example

Lowest Modular level 

for ISS 

Levels for 

AMPS 

Development 

Figure 2.—Levels of Assembly for Modular Power Systems. 
NASA/TM—2013-217813 8 

4.1.3 Modular Component 

AMPS defines a Component Level Module as an encapsulated unit that cannot be further disassembled 

with common tools. This may be a single discrete part or a combination of multiple parts that are 

encapsulated in a manner that makes them a monolithic module. It is the lowest practical level for 

replaceable units that do not require special processing and equipment. The project may simplify the 

replacement process by using space qualified sockets instead of solder joints. The component level serves as 

the common fundamental power building block. A number of different functions can be provided by the 

electrical arrangement of these blocks. These blocks will incorporate sensors and other devices that support 

diagnostics but the diagnostics and health management functions will be handled at the next higher level. 

4.2 Subassembly and Component Level Modularity Features 

Driving modularity down to lower levels of assembly enables the designer to construct higher level 

assemblies with a common set of configurable building blocks. These lower level modules will 

incorporate features normally found in higher level assemblies. 

4.2.4 Physical Encapsulation and Structure 

Similar to high level modules, lower level modules should appear as physically independent encapsulated 

units. Where practical, modules will be designed to be self-contained and structured to be portable. 

4.2.5 Functional and Software Encapsulation 

The modules will have embedded functions and related software. The software may be 

microcontroller code or field programmable gate array firmware. This code may be modified to support 

reconfiguration or upgrades. 

4.2.6 Independent Thermal Paths 

Subassembly level modules may be configured in their own independent enclosure and may also 

provide its own thermal paths for removing heat. 

4.2.7 Standardized Electrical Power Interfaces 

Since many low level modules will be connected in series or parallel groups the module to module 

connections will be standardized.

4.2.8 Smart Network Capable Controllers 

Subassembly modules will include “smart” network capable controllers (as defined in IEEE 1451) 

and will pass data over a network connection rather than a dedicated backplane (Ref. 7). 

4.2.9 Embedded Health Management 

The use of smart network capable controllers accommodates module related health management 

along with control and data functions. Most HM monitor existing control loops but other specialized 

sensors may be provided for added diagnostic and prognostic capability. 

4.2.10 Embedded Supportability 

Supportability functions work alongside control and HM functions. The primary purpose is to assist 

in the repair processes, and provide data that assist logistics and maintenance. Since modules may be 

moved from vehicle to vehicle they need to store their own configuration data and service history in an 

“electronic log book”. This enables the crew and ground support to interrogate the unit for its identity, its 

current configuration, maintenance status, service history and records of any anomalies. This also allows 

the crew to confirm that a module is the appropriate unit for its intended next application. It can also 

indicate how much remaining life is available based on operating experience. 
NASA/TM—2013-217813 9 

4.3 Defining Modular Scaling Increments 

Although the approach is to use modular blocks to build up power systems no one module size can be 

expected to meet the needs of all spacecraft. Using many small blocks means many interconnections and 

drives up the mass of harnessing while reducing overall reliability. Using large blocks means a tendency 

for overcapacity and excessive mass. An assortment of modular elements with varied capacity enables 

designers to mix and match modules to arrive at a system that meets mission needs. 

For the AMPS project the generic term; “scaling increment” is defined here as the incremental 

capacity of a modular building block element. An assortment of several “scaling increments” may be 

needed to provide the flexibility to group elements together to meet the mission needs. Power generation, 

energy storage and power management will each have distinct scaling increments. 

An example of the scaling increment can be found in the assorted sizes found in a standard electronic 

parts kit, E6 series resistors, for example, have a 20 percent tolerance and six distinct increments for 

every decade of value (10, 15, 22, 33, 47, 68) which allows the user to achieve any practical value 

between 10 and 100 with a series combination of just two units. The assortment is further reduced in the 

E3 series to only three values (10, 22, and 47) with 40 percent tolerance. Using the E3 series for power 

implies six increments to span vehicle power range from 1,000 to 100,000 W. This is a simplistic analogy 

and assumes that the design objective is to minimize the number of units. Other considerations, such as 

redundancy, may drive the design toward three or more units. 

The scaling increments are dependent on other design factors, such as, redundancy requirements, 

thermal control requirements, mass of harnessing, and the practical limits of handling by crew or robotics. 

Each power subsystem will need to define its own scaling increment. 

5.0 HAT Vehicle and Design Reference Mission 

The AMPS project uses the vehicles from the Human Architecture Team (HAT) studies and their design 

reference missions (DRM) (Ref. 1). Each vehicle has a mix of common and unique mission power system 

requirements. At this stage the maturity, these requirements also change dramatically base on the specific 

DRM being studied. To assure that Module Power stays relevant to these missions some members of the 

AMPS project team also support the power system architecture development for HAT vehicles. 

Examining the HAT DRMs, shown in Table 1, the vehicles tend to fall into a NEA Vehicle Group 

including Deep Space Habitat (DSH), Solar Electric Propulsion (SEP) vehicle, and Multi-Mission Space 

Exploration Vehicle (MMSEV-NEA). Otherwise they fall into a Lunar Vehicle Group that includes two 

variations of the Lunar Lander and a Multi-Mission Space Exploration Vehicle configured as a surface 

rover. Common to both vehicle groups is the Cryogenic Propulsion Stage (CPS) and thus shown as a 

“Common” vehicle. The study did not include the Space Launch System (SLS) or the Multi-Purpose 

Crew Vehicle (MPCV) because there development is more mature and will not be directly influenced by 

the AMPS technology development. The AMPS project will track these vehicles and develop, as needed, 

a suitable interface that assures interoperability with these legacy vehicles. 

TABLE 1.—AMPS STUDY VEHICLES 

Group Near Earth Asteroids Common Lunar Surface Vehicles 

Vehicle Deep 

Space 

Hab 

Solar 

Electric 

Propulsion 

MultiMission 

Space 

Exploration 

Vehicle 

CryoPropulsion 

Stage 

MultiMission 

Space 

Exploration 

Vehicle 

Lunar Lander: 

Lunar Orbit 

Rendezvous 

Lunar Lander: 

Earth Orbit 

Rendezvous 

DRM NEA 

Transit & 

Proximity 

NEA 

Transit 

NEA 

Proximity 

NEA & 

Lunar 

Surface 

Rover 

Descent 

Module 

Ascent 

Module 

Descent 

Module 

Ascent 

Module 

Note: Near Earth Asteroids are occasionally referred to as Near Earth Objects (NEO)
NASA/TM—2013-217813 10 

5.1 Cryogenic Propulsion Stage 

The Cryogenic Propulsion Stage (CPS) appears in nearly all missions beyond LEO. In many cases 

multiple CPS units are used. CPS had a Block 1 and 2 variant. CPS Block 1 was a simple Cryo Stage 

operating on batteries. CPS Block 2 is configured for missions with long loiter periods (months). The 

substantial solar arrays capacity is needed to drive “zero-boil off” cryo-coolers to preserve propellants. 

CPS Block 2 is also used as a tanker or fuel depot to support NEA Missions. The study focused on CPS 

Block 2 for power requirements. 

5.2 Deep Space Habitat 

The Deep Space Habitat is the most complex system and may involve up to three distinct nonpropulsive vehicle segments. DSH is also expected to acquire system heritage from the ISS. The DSH 

appears in nearly all NEA missions and serves as the habitat and core vehicle for the crew in transit and at 

the destination. It has substantial power needs. For long duration missions the main habitat is augmented 

by a node for attaching multiple vehicles including MMSEV. In addition, a logistics module is added to 

provide crew consumables, as well as, store vehicle spares. The DSH is considered the central hub for 

power system maintenance activity and spares. 

5.3 Solar Electric Propulsion 

Solar Electric Propulsion vehicle is unique since it has a power level exceeding 300 kW but is a 

relatively simple unmanned vehicle. As a low thrust high specific impulse vehicle, it’s primarily used 

well beyond low Earth orbit. In some scenarios, it is delivers to a CPS stage to L1 orbit. This vehicle has 

a split voltage system where 300 V is used as “direct drive” for the electric propulsion modules and 120 V 

is used for vehicle housekeeping. 

5.4 Lunar Landers (Lunar Orbit and Earth Orbit Rendezvous) 

For human missions lunar landers are actually composed of two modules. A Descent Module (DM) 

for landing and an Ascent Module (AM) that returns the crew to orbit. For uncrewed cargo missions the 

vehicle has only a Descent Module. There are two types Lunar Landers based on where the Multi-Purpose 

Crew Vehicle rendezvous with it which, in turn, drives the power configuration. 

 Lunar Orbit Rendezvous (LOR) Lunar Lander is equipped with large solar arrays to power the cryocoolers that preserve propellants during a long loiter period in Lunar Orbit. 

 Earth Orbit Rendezvous (EOR) Lunar Lander is similar to Altair and thus has a very short loiter time. 

This vehicle taps the propellant supply for fuel cell generated power. 

 Ascent Module (AM) The ascent module is very similar on both LOR and EOR Landers. The Descent 

Module powers Ascent Module when mated. After separation the independent Ascent Stage runs 

exclusively on batteries. The HAT studies show them as primary batteries however the long loiter 

times suggest that rechargeable secondary batteries may be more suitable. 

5.5 Multi-Mission Space Exploration Vehicle: Near Earth Asteroid 

The Multi Mission Space Exploration Vehicle (MMSEV) comes in two configurations. For the study 

it is treated as two distinct vehicles that have very different mission profiles and power requirements. The 

MMSEV NEA unit has a modular core habitable designed to support in-space missions. MMSEV NEA is 

a free flying vehicle for close-in exploration of asteroids. It maneuvers via a detachable propulsive 

module. It employs configurable detachable module called a PUP as a means of outfitting the vehicle for 

varied missions. This version of the MMSEV is expected to rely on solar arrays and batteries. 
NASA/TM—2013-217813 11 

5.6 Multi-Mission Space Exploration Vehicle: Planetary Surface Rover 

MMSEV Rover is a surface vehicle the maneuvers on a sophisticated wheeled mobility chassis. This 

version is a direct descendent of the Constellation Lunar Electric Rover (Ref. 9). Like the NEA version, 

the Rover version has a modular core habitat. The modular mobility chassis has six motor driven wheels 

assemblies that have independent steering and suspension height control that are all electrically driven. 

The high power demands of the mobility chassis require a 300 V power system in addition to the 120 V 

core. The MMSEV Rover may use either solar arrays or fuel cells. The rugged terrain and frequent 

shadowing limit the effectiveness of solar arrays. Further, because surface operations offer an opportunity 

to exploit in-situ resources, the unit is expected to use fuel cells when underway. 

MMSEV appears to be most ideally suited for advanced modular power from the operational 

flexibility and configurability standpoint. Because MMSEV is conceived as a highly modular 

reconfigurable vehicle the MMSEV regarded as an important evaluation platform for modular power. 

Thus AMPS is supporting MMSEV ground demonstrations with advanced solar array, battery and fuel 

cell technologies. 

5.7 Level 1 Power Requirements 

Table 2 lists the top level requirements. Crew size and mission duration are key parameters that govern 

the size and performance of vehicle power systems. The maximum, nominal power requirements size the 

capacity of power generation, and power distribution. Energy storage is affected by the availability of solar 

power in insolation (daylight) periods and the duration of eclipse periods where batteries must bridge the 

gap. Fault tolerance levels also affect the series/parallel arrangement of modular power. 

TABLE 2.—LEVEL 1 VEHICLE POWER REQUIREMENTS 

Vehicle Grouping Common

Vehicle or Space Facility Deep‐Space

Habitat

Solar Electric

Propulsion

Multi‐Mission

Space

Exploration

Vehicle

Cryo Propulsion 

Stage Multi‐Mission Space

Exploration Vehicle

Design Reference Mission DSH at NEO SEP NEO MMSEV at

NEO NEO/LEO/HEO Surface Rover Descent Module Ascent Module Descent Module Ascent Module

Level I Requirements

Crew Capacity Maximum 4 N/A 4 N/A 04 4 4 0

Maximium Mission Days 365 676 7 14 TBD 7.5 7.5 21.7 0

Maximum [crew*days] 1640 676 28 N/A TBD DM: 4* 7 days, 6hrs   AM:  4 * 19 hrs 48.8 0

Total Maximum Power [kWe] 20 320 3 25 10 7.9 0 7.1 3.1

Nominal power [kW] 15Kw ‐ Nominal N/A tbd 23 3.5 DM:  4.7 kW    AM:   3.2  kW 4.5 2.9

Fail Safe Power [kW] 10kW for

survival See Total above tbd N/A 1.2 3.1 kW 0 N/A N/A

Insolation [min] 240 (worst case) 0 0 56 0 71.3 min 0 none none

Eclipse [min] 120 (worst case) 44.27 min 0 36 0 46.5 min 0 0 0

NEAVehicles

Lunar Lander

LOR Battery

Lunar Lander

EOR (Altair) Fuel Cell

Lunar Vehicles

6.0 Modular Power Application Scenarios 

To best understand how the modularity features assure flexibility, supportability and reusability we 

will step through operational scenarios. Operational scenarios help expose requirements and drive 

modular power features. Some features can be evaluated in operational field demonstrations. 

6.1 In-Flight Reconfiguration Scenario 

In the following discussion we focus on the MMSEV in a NEA mission scenario. This vehicle is 

unique in terms of its level of flexibility and its development heritage. This vehicle is a result of years of 

lunar rover field demonstrations. It evolved as a multipurpose vehicle with a high degree of 

configurability and modularity. This vehicle level modularity is now being extended to operations beyond 

cis-lunar space to Near Earth Asteroid. 
NASA/TM—2013-217813 12 

As shown in Figure 4, for NEA applications, the MMSEV’s wheeled rover mobility chassis is 

replaced by a propulsive module that uses Reaction Control System (RCS) thrusters for mobility. A 

modular feature that is carried over from the rover concept is the PUP (illustrated in Fig. 5). The PUP is a 

detachable vehicle level module that configures the vehicle for a specific mission or sortie. The PUP is 

particularly important because it serves as a platform that can exploit the flexibility of modular power. 

The Modular Power can provide combinations of Solar Arrays, Batteries and Fuel Cells that are 

appropriately scaled for the mission. Further, the modular PMAD features of AMPS Power will provide 

any power conversions required for adapting the power hardware to the vehicle. The integration of the 

hardware may be a “live” process where the hardware is internally powered and aware of the integration 

process. Power transmission may be suppressed until the AMPS system can validate connections and 

hardware readiness. To minimize crew operations the configuration of power hardware is handled via the 

plug and play network interoperability features. Data links to modules prior to connections may be 

provided by a wireless means such as Bluetooth or other protocols covered by IEEE 1451. 

To assure a smooth and safe integration process embedded Supportability and Health Management 

features assess the status and readiness of individual elements before, during, and after they are integrated. 

Once verified that hardware is safe and properly configured, AMPS will enable power delivery. 

MMSEV: Core 

Modular 

PUP

Solar Array/Battery

Solar Array/Flywheel 

Fuel Cell Modules 

PUP Power 

Configuration 

Options 

Figure 5.—MMSEV portable utility pallet power configurations. For specific missions, the PUP power subsystem 

may be configured by using interchangeable modular power building blocks. 

MMSEV Surface Rover version MMSEV Near Earth Asteroid Version 

Mobility Chassis RCS Sled 

Figure 4.—Multi-mission space exploration vehicle variants: MMSEV provides the common platform with the 

flexibility to configure the vehicle for specific missions. The Rover Mobility Chassis is interchangeable with the 

Reaction Control System (RCS) sled. 
NASA/TM—2013-217813 13 

A further benefit of the AMPS approach is that the PUP’s power equipment need not be composed of 

dedicated hardware. The flexibility and supportability features also allow the hardware to be salvaged 

from other vehicles in the mission. For example a solar array may be “borrowed” from the CPS or SEP 

vehicles. The Solar Electric Propulsion stage, in particular, is capable of sparing a solar array to support 

MMSEV. Practical application of this approach is highly dependent on vehicle-to-vehicle commonality, 

rapid access to power hardware, the system scaling increment, and the ability of the donor vehicle to 

adapt to the reduced capacity temporarily or permanently. 

NASA field demonstrations with the MMSEV and PUP can be used to validate the modular power 

approach. Demonstrations will be used evaluate the effectiveness of modular features and their impact on 

operational flexibility. 

6.2 In-Flight Repair Scenario 

This example scenario takes place on the Deep Space Habitat during a NEA mission. The DSH is a 

substantial vehicle that houses the crew for most of the mission. It has many design features that it 

inherited from the ISS habitable modules. In this case DSH also includes a logistics module that aside 

from providing mission consumables it provides spares storage and a modest work area. 

Because all the vehicles in the mission stack have common systems the inventory serves as a common 

source for spares for all the mission vehicles, including spare power modules. To allow the flight 

hardware to be replaced by the limited dexterity of EVA encumbered astronauts or by remotely controlled 

robots, systems are accessed at the assembly level as ORUs. To avoid the mass penalty of sparing full 

sized ORU assemblies most spares are handled as subassembly and component level modules, as shown 

in Figure 6. Therefore, ORU units are retrieved and repairs are performed in the logistics module by 

replacing faulty sub assemblies or components. Once repaired, ORU is returned to service. If the 

subassembly is repaired by a replacement component then it becomes a spare subassembly. 

 

Sub Assembly 

Spares 

Component 

Spares 

MMSEV 

Repair & 

Spare 

Return to 

Service 

Replacement 

Unit 

Replacement 

Unit Deep Space Habitat 

Logistics Module 

Figure 6.—In-flight modular repair scenario. A faulty Assembly Level ORU is removed and a faulty subassembly 

module is replaced allowing the ORU to be immediately returned to service. The subassembly fault is traced to 

a component module that is replaced and the repaired subassembly is placed in spares. In this scenario, 

modular assemblies and sub assemblies are preserved and only a component level module is consumed. The 

system is restored with minimum impact on logistics. 

. 
NASA/TM—2013-217813 14 

Examining the ISS logistics and maintenance practice reveals that repairs are not typically performed 

onboard the ISS below the ORU level (Ref. 4). Lower level repair on conventional electronic and electrical 

hardware implies a significant increase in tools, diagnostics equipment, process equipment such as 

soldering, coating removal and application equipment. Electronics repair studies indicate that there are 

substantial mass and volume penalties for this equipment (Ref. 7). Repair of non-modular component 

replacement requires substantial practiced skill and manual dexterity that is not typical of flight crews 

training. Grouping of loose components into “simple to replace” modules minimizes tools and process 

complexity. 

A supportability study (Ref. 4) examined repairs of ORUs for space shuttle and space station at 

NASA depots and found that the actual removal and replacement of faulty hardware only made up 

20 percent of the overall process. The remaining 80 percent involved problem diagnostics, hardware 

de-integration, post replacement re-integration and functional test and check outs. Much of the work was 

a consequence of the violation of system integrity that occurs during the repair process. 

The AMPS approach is to modularize the lower level subassemblies and components in a manner that 

makes them easy to remove and replace with minimum disruption of integrity. Further diagnostics and 

built in functional tests will be embedded at these lower levels to quickly isolate the root cause and 

expedite the repair process. The Supportability and Health Management features also aid in reducing 

repair process complexity. HM monitors internal sensors and control loops. By tracking long term 

changes in a control feedback loop it can indicate that a particular component is degrading. Reporting this 

through the HM channels alerts the crew to a fault and indentifies the specific replacement needed. This 

not only reduces diagnostics effort but also allows the crew to carry the specific spare to the site for an 

in-situ low level replacement. Supportability features can assist by rerouting power and isolating the unit 

for safe in-situ repair. Supportability Plug-and Play features also allow the module to be tested 

automatically before reconnecting to the system and automatically updating the system to indicate the 

presence of a new hardware. This type of modular repair operation scenario can be demonstrated and 

evaluated in concert with vehicle field demonstrations. 

Component 

Spares 

Sub Assembly 

Spares 

Assembly Spares 

Salvaged 

ORUs

Salvaged 

ORU 

Hardware replaced 

as needed 

Salvaged Hardware 

Reconfigured 

Reallocated to MMSEV 

Salvaged Hardware 

evaluated and 

placed in Spares 

Hardware Salvaged 

from Expended CPS

Deep Space Habitat 

MMSEV Logistics Module 

Figure 7.—Modular hardware salvage scenario. Salvaged hardware from expended vehicles may be 

reallocated directly to another vehicle or to DSH spares inventory. All salvaged hardware will need 

to be evaluated and reworked if needed before they are reused. Embedded Health management 

and supportability features will expedite this process. This approach also provides DSH with a 

source of assembly level hardware. 
NASA/TM—2013-217813 15 

6.3 In-Flight Hardware Salvage Scenario 

As noted earlier, modular power offers some opportunities to exploit other vehicles for system 

hardware. This can also apply to spares inventory. Even though spare sizes are minimized, DSH may still 

struggle to provide spares for all vehicles. In this scenario we consider another resource for flight 

hardware (Refs. 2 and 10) as illustrated in Figure 7. 

Most vehicles are still functional after they have served their initial mission role and are thus a rich 

source of hardware. Commonality in power architectures, common interfaces, along with, mechanical and 

electrical encapsulation makes modular power hardware portable. Smart Plug and Play interfaces 

simplifies de-integration and re-integration process (Ref. 6). Using appropriate scaling increments assures 

that the modules are a manageable size for EVA and robotic handling. 

There are a number of considerations for scavenging or salvaging hardware. Although hardware can 

be expected to be in good condition it may still experience degradation which will impact all the 

spacecraft in the stack. The crew will need to determine the relative condition of the power systems across 

all spacecraft. The crew will want to salvage the best for the remainder of the mission and that may mean 

removing hardware from the DSH or MPCV and replacing it with hardware from the other expended 

vehicles like CPS or SEP. This salvaging and redistribution operation ranges from small regulator 

modules to battery, solar array and fuel cell subsystems. 

To sort out which units will be selected in a redistribution of hardware, the embedded supportability 

and health management features will, once again, assist the crew. Supportability and HM features can 

assess individual modules, report prior anomalies, indicate accumulated service life, and predict 

remaining life. 

7.0 Modular Power Management and Distribution 

7.1 The Approach to Modularity in PMAD Systems 

Current practice in Power Management and Distribution design is to develop mission specific 

solutions. Each mission has unique loads requirements, power quality specifications, operational 

requirements, etc. As shown in Figure 8, PMAD elements (e.g., load and bus regulators, energy storage 

subsystem interfaces, and protection) are designed to optimally meet these unique system and subsystem 

requirements and cannot be used in any other vehicles. With future exploration dependent on multivehicle missions there is an opportunity to consolidate power system Design, Development, Test and 

Evaluation costs. To do this however, means that power systems need to share a common approach and 

architecture. These vehicles vary widely in their power needs; therefore, imposing a common power 

Figure 8.—Elements of a power management and distribution system. 
NASA/TM—2013-217813 16 

quality specification is not sufficient to achieve true commonality. However, by organizing low level 

elements into modular building blocks, a power architecture that is scalable with substantial commonality 

can be built. PMAD system is composed of two types of electronic devices; power electronics and control 

electronics. The power electronics (includes electromechanical devices) conduct and direct the current 

and tend to scale in relation to the power loads. Control electronics manages the power electronics with 

low power mixed signal (digital and analog) devices that scale somewhat independently of power loads. 

A modular building block approach must account for differing scaling factors. 

7.2 PMAD Module “Scaling Increment” 

As discussed in Section 4.0, the scaling increment is the predefined module capacity associated with a 

building block. Several scaling increments may be needed to provide the flexibility to group elements in 

parallel and/or series combinations to meet the wide range of mission needs. 

While different control techniques are required for different applications, regulation functions can all 

be addressed with common power converter designs. That is, solar array regulation, bus regulation, 

battery charge/discharge control, and load power regulation can all be done with common DC-DC 

converters (referred to as Flexible Power Modules). These need only be paralleled in sufficient numbers 

to meet the operational requirements. 

A preliminary analysis of the HAT vehicle electrical systems requirements resulted in the 

identification of four modular converter ratings: 500 W, 750 W, 1 kW, and 2.5 kW that serve as the 

primary scaling increment for converters. With multiple increments in converter size one can mix and 

match the needs with minimum surplus capacity and the fewest number of parallel converters. 

7.3 Modular Impact on System Redundancy 

The secondary effect of the modular approach is to provide redundant paths for power delivery. On 

ISS, redundancy of PMAD systems is accomplished at the channel level, and requires that either the 

distribution system be re-configured (e.g., channel cross-ties closed) or low-criticality loads be shed. This 

approach is suitable for ISS with eight power channels and a total system capacity much larger than any 

individual load. These redundant channels provide a full power capacity with a substantial mass penalty 

which may not be acceptable for smaller spacecraft. 

An advantage of the modular approach is the ability to continue operation but at a reduced capacity 

after a failure. The need for full continuous power depends on the phase of flight and the criticality of the 

system supported. For launch and landings, in particular, full power is essential. Other phases of flight 

could tolerate the degradation, particularly if it is a temporary condition and recoverable by a maintenance 

action by the crew. This is consistent with the desire to have “graceful degradation” where the crew has 

time to detect and respond to a failure. 

By designing backplanes and associated controllers capable of accepting converters of multiple power 

ratings, inter-changeability results. This inter-changeability allows for the potential scavenging of power 

converters from a non-critical system to a higher criticality system. 

An assumption made for the modular PMAD systems approach was that the converters would process 

all of the power in the same way that the baseline systems typically do. However, mass savings can be 

attained by using advanced topologies. The Series Connected Boost Regulator (SCBU), for example, is 

one such topology which implements a DC-DC analogously to an autotransformer. By using this 

connection the power converter need only boost a fraction of the total load power, resulting in significant 

mass savings (Ref. 11). 
NASA/TM—2013-217813 17 

7.4 Modular Power Distribution Units 

The regulators are paralleled to build-up an assembly level Power Distribution Unit (PDUs) to 

accommodate specific requirements. The following describes the modular elements used to define a 

Power Distribution Unit. 

7.4.1 Control and Protection 

System protection and switching functions can be accomplished with either solid-state or a hybrid 

combination of electro-mechanical and solid-state switchgear. These are packaged as Remote Power 

Controllers (RPCs), which consist of a family of solid-state switchgear of different ratings. RPCs also 

offer benefits such as active current-limiting, soft-starting, paralleling, etc. Modular RPCs are 

subassembly level modules connected to a chassis backplane. They are a mix of capabilities in terms of 

channel count and channel capacity. Based on the HAT vehicle analysis, RPC composition ranges from 

multi-channel 5 A modules to single channel 250 A modules. 

7.4.2 Modular Chassis

For all functions, excluding battery controllers, four modular chassis were identified to accommodate 

the typical PMAD functions of regulation, switching, and protection. All chassis types accommodate RPC 

cards, housekeeping, communication, data acquisition cards and an application specific control card. 

Chassis Types: 

 Type 1 Chassis accommodates five 500 W converters and nine RPC cards. 

 Type 2 Chassis accommodates five 2.5 kW and 9 RPC cards. 

 Type 3 Chassis accommodates switchgear-only. 

 Type 4 Chassis accommodates one 2.5 kW (or five 500 W) converter, and it has fewer slots for RPC 

cards. 

Note: An advantage of the modular design approach is that a single 2.5 kW converter can be used in the 

Type 1 chassis if scavenging is necessary. Similarly, a single, or up to five, 500 W converters can be 

placed in any single 2.5 kW “slot” of the Type 2 chassis. It is possible to fit up to (25) 500 W converters 

in a Type 2 chassis. 

7.4.3 Bidirectional Energy Storage Controllers 

Modular batteries may incorporate bi-directional converters and controllers that are integrated with 

the battery chassis. Battery regulators have the additional requirement to provide bi-directional power 

flow which separates them from other regulators. Once again a unique power need is met by using an 

appropriate selection of scaling increments and using them as modular building blocks. For battery 

charge/discharge regulation the modular scaling increments were determined to be 750 and 1000 W. 

7.5 Modular PMAD Tech Development and Demonstration 

A goal of AMPS PMAD is to develop a Flexible Power Module (FPM), illustrated in Figure 9 which 

is one of the power converters to be used as building block for any type of electrical regulator in a PMAD 

system. The design of the FPM targets four typical space power applications: solar array regulators, bus 

voltage regulators, battery charge regulators/discharge regulators, and load converters. The modules will 

be interchangeable (plug-and-play) and automatically configure themselves for the application. These 

FPMs will communicate with each other to offer power sharing and phase staggering to reduce overall 

EMI. Individual modules will be enabled or disabled based on demand to increase overall system 

efficiency and compensate for failed modules. 
NASA/TM—2013-217813 18 

The FPM is being developed along three parallel paths. Path #1 uses COTS power converters to 

implement a solar array regulator to which custom controls are added to demonstrate a majority of the 

features of the FPM. Path #2 is a bus regulator using ISS power converters that allow for additional 

features to be demonstrated, such as phase staggering. Path #3 involves development of a bi-directional 

power stage to be used in all applications. Intelligent controls developed in the first two paths will be 

integrated with the power stage from the third path to demonstrate the flexibility and adaptability of the 

modular PMAD approach. 

Currently, demonstrations of Path #1 and #2 have been successfully tested to prove basic 

functionality of these concepts. Path #1 was successfully tested in an end-to-end power system simulator, 

proving the power sharing, failure compensation, and increased efficiency concepts. Path #2 added 

pseudo-masterless control, serial communication, and phase staggering among the power converter 

modules. The bi-directional power stage of Path #3 is under development and will be combined with the 

controls proven in the first two demonstrations to yield the final prototype module. This module will be 

tested as a battery charge/discharge regulator, bus regulator, and solar array regulator. 

8.0 Modular Solar Array Technology 

8.1 Key Solar Cell Performance Parameters 

The key parameters that characterize solar array performance are specific power (W/kg) and areal 

power (W/m2

). Various mission-related environmental effects and degradations affect both the solar array 

power and these parameters. 

Another key performance characteristic is the acceleration capability. Normally, large solar arrays are 

designed for low accelerations because they are heavy and the added structure needed to maintain their 

integrity for higher accelerations add a prohibitive amount of mass. For the range of vehicles in this study, 

the option to fire thrusters after the solar arrays are deployed necessitates either stiffer solar arrays or 

retractable ones. The latter option is typically unacceptable due to risk involved in redeployment after 

performing a thruster maneuver (such as trans-lunar insertion). During the study, it was assumed that all 

vehicles had to endure 1 g acceleration except the MMSEV Surface Rover and Lunar Descent Module 

(which had a 2.7 g requirement) and the very large SEP vehicle (0.2 g requirement). 

Figure 9.—Flexible power module. The FPM unit serves as a common interchangeable building block that is 

employed throughout the system for a variety of PMAD functions. 
NASA/TM—2013-217813 19 

Voltage level was also a key attribute and 120 V was assumed for all vehicles except the SEP which 

assumed 300 V. These high voltages enable the reduction of electrical harness mass. 

8.2 Solar Cell Technology Options 

Two approaches were assumed for the solar cells. (1) Because of mass and size limitations, the SEP 

vehicle was assumed to use IMM (inverted metamorphic multi-junction) cells with an assumed beginning 

of life (BOL) efficiency of 34 percent. These cells are one tenth the thickness of traditional cells. These 

cells are assumed to be available in quantity for the large SEP vehicle for the medium to far term. (2) 

Thinned gallium arsenide triple junction solar cells with a BOL efficiency of 30 percent are assumed for 

use in the solar arrays of the remaining vehicles primarily because they will be available in large 

quantities in the near to medium term at lower cost than IMM cells. 

8.3 Modular Solar Array Concepts 

8.3.1 Folding Rigid Panels 

A wing composed of multiple, modular panels could be utilized, although these are typically are used 

for low acceleration (1 g) and/or low required spacecraft power levels or when the structural mass 

needed to withstand accelerations can be accommodated. The following Figure 10(a) illustrates this type 

of solar array on ISS. 

In the AMPS modularity study, due to the higher power levels and accelerations, folded rigid panels 

were not considered. The MMSEV Rover baseline was originally rigid panel solar arrays; however, a 

modular approach was able to replace the rigid panels with an adequately designed roll out solar array. 

Figure 11.—A Roll Out Solar Array (ROSA) shown deployed on left and stowed on right (Ref. 19). 

Figure 10.—Solar arrays examples. (a) Folding rigid panels used on ISS. (b) UltraFlex solar array wing deployed 

in ground tests. (NASA photo) 

(a) (b) 
NASA/TM—2013-217813 20 

8.3.2 UltraFlex, MegaFlex 

This type of solar array is best seen in examples such as Mars Phoenix (Ref. 12) and the CEV Orion 

(Refs. 13 to 15) spacecraft. Figure 10(b) shows an Ultraflex wing on the ground. Regardless of diameter 

(up to ~12 m or ~35 kW), the Ultraflex solar array (Ref. 16) is typically composed of 10 “gores” (i.e., 

triangular segments) to approximate a circle. Solar cells are attached throughout the gores into “strings” 

of series connected cells. Although these strings may potentially be able to be assembled modularly, this 

study assumed the gore was the minimum modular level. Larger Ultraflex wing designs (i.e., MegaFlex 

(ATK Corporation)) have been proposed for much higher power levels in one deployable structure 

(~30 m diameter, ~180 kW). The acceleration capability of the UltraFlex wings can be as high as 2.7 g 

(CEV Orion), although more likely 0.2 g for MegaFlex sized versions. 

8.4 Solar Array Scaling Increments 

Modularity based on solar arrays composed of numerous mass producible cell module building blocks 

was assumed. Cell modules are typically a series connection of solar cells (i.e., a string) to provide the 

voltage required by the power system. In this study, the cell module is rectangular and includes the coverglasses for radiation shielding, coatings, substrate structure (if any), and mechanical/electrical interfaces. 

These cell modules are assembled into wings to obtain the required power levels with each vehicle’s wing 

sharing the maximum number of parts across the range of vehicles. Only ROSA-like wings shown in 

Figure 11 were considered due to the straightforward approach in modularizing at the string (cell module) 

level. 

For ROSA-like wings, the strategy is to design the roll out tubes to have common diameters, but 

obtain more power by increasing the length (i.e., adding more solar cell modules). It was desirable to 

maintain the wing width (to maintain similar parts, increase commonality of structural analysis and 

testing). Comparing the required power levels and acceleration levels, it was determined that two kinds of 

ROSA-like wings were needed. The SEP vehicle wing width was 6.5 m but it requires two wing lengths 

(10.5 and 17.2 m), to accommodate electric propulsion plume impingement, with a 4 in. roll out tube 

diameter. For all the other vehicles, the width of the wing was 3.6 m with the wing lengths varying from 

4.5 to 14.4 m and with an 8 in. roll out tube diameter. The reason the SEP roll out tube diameter is less 

than that of the other vehicle wings is because of the lower acceleration level (1/10 to 1/5 lower). The 

reason the range of wings can utilize the same 8 in. roll out tube diameter is because the vehicles with 

high acceleration requirements (2.7 g) have lower power levels than the ones with 1.0 g acceleration 

requirements (i.e., since power level is tied to the length of the wing, then a lower power level means the 

wing length, mass, and moment arm is lower, increasing its stiffness using the same diameter tube). 

The cell modules are similarly uniquely designed between vehicles, namely the high voltage (300 V), 

IMM solar cells of the SEP vehicle have their own cell module design (0.77 by 0.23 m in size and 

0.29 kg in mass), while all the other vehicles at the lower voltage (120 V) using thinned triple junction 

cells have their own common cell module design (0.77 by 0.23 m and 0.22 kg). Cell modules for all the 

120 V vehicles are designed to handle the worse case common environments to minimize the unique 

designs and enable maximum modularity and economies of scale and common testing/qualification. 

8.5 Modular Solar Array Logistics Benefits 

Standardizing the number of ROSA module widths with their associated roll out tubes (diameter, 

material, and thickness) has the potential of reducing development and testing costs for the wing 

structures. Selection of optimal cell module sizes enables automated testing and inspection of modules. 

Modules that fail can be swapped out with ease. The SEP vehicle due to its size and number of wings (16) 

can be considered modular on that standpoint alone. SEP as part of a multi-vehicle mission, can serve as a 

source for spares for other vehicles. Once SEP’s mission phase is complete, the SEP arrays are suitable 

for salvage and reuse. The multiple deployment capability of the ROSA array simplifies the salvage 

process and its compact roll-up form is suitable for stowing and redeployment. 
NASA/TM—2013-217813 21 

Figure 12.—ROSA array deployment demonstration. MMSEV 

and ROSA are shown integrated on to the RCS Sled to 

demonstrate array deployment and its behavior during RCS 

propelled maneuvers. (Credit: Rafael Jimenez NASA JSC). 

8.6 Modular Solar Array Development and Demonstration 

In 2012 AMPS worked in concert with the MMSEV project at NASA Johnson Space Center (JSC) to 

demonstrate the multiple deployment capabilities of the DSS ROSA array. Figure 12, shows a threedimensional (3D) model of ROSA demonstrator integrated with a MMSEV vehicle mockup in its near 

Earth asteroid configuration. MMSEV and ROSA are supported on the Reaction Control System (RCS) 

Sled. The RCS Sled is an air bearing suspension system that enables the vehicle to maneuver using cold 

gas jets, in simulated low gravity operations. 

9.0 Modular Battery Technology 

The battery is an ideal example of a system composed of simple modular blocks. Battery capabilities 

are defined by the characteristics of the fundamental cell unit. The selection of the cell chemistry and 

geometry translate into battery characteristics and performance. 

9.1 Performance Parameters 

Major battery design drivers include the Amp-hour (Ah) capacity, peak and average currents, and 

voltage that the battery is required to deliver, as well as the cycle life, operational life, and redundancy 

requirements on the battery. Ah capacity, current, and voltage define the size of the battery. Fixed cell 

size establishes the Ah, current capacities and voltage increments. Ah capacity and the current scale up 

with the number of cells in parallel. Voltage scales by the number of cells in series. Life requirements 

define the extra margin required to be able to perform the required functions at the end of life (Ref. 20). 

Redundancy requirements determine the number of extra strings or batteries required to meet reliability, 

loss-of-crew, or loss-of-mission requirements. Overall, the primary attribute used to choose between 

design options for a given set of requirements is usually mass. 
NASA/TM—2013-217813 22 

9.2 Battery Chemistry and Geometry Options 

Current spacecraft batteries are mainly designed using lithium-ion battery technology for any mission 

that requires recharging. Other rechargeable battery chemistries are available, but their energy per unit 

mass is not as high at the current time. Lithium-ion technology has a variety of choices available in 

“space-qualified” cell designs. Lithium-ion cells are available in prismatic, cylindrical, and pouch 

formats. Pouch and prismatic cells require compression to achieve their optimum performance, while 

cylindrical cells do not. Pouch cells also have more specific handling requirements to avoid damage to the 

internal components. As shown in Figure 13(b), cylindrical cell are stable pressure vessels and suited for 

space with little additional structure. 

9.3 Modular Battery Concept 

Lithium-ion cells are available commercially in a large range of Ah capacities. The Ah capacity and 

current requirement of the battery can be met by using a large number of small Ah capacity cells or by 

using a small number of large Ah capacity cells. For safety reasons, overheating of any cell in a lithiumion battery should be avoided. Overheating can be caused by overcharging, short circuit, cell reversal, or 

over-temperature of any cell. These conditions are normally avoided by controlling the charge/discharge 

of the individual cells and/or extremely close matching of the cells in the battery to ensure similar 

performance during operation. Matching and control of fewer numbers of large capacity cells is easier 

than that of large numbers of smaller capacity cells. 

AMPS modular batteries will be closely coupled with charge/discharge functions provided by PMAD. 

Modular batteries will also incorporate Health Management largely through extensions of the charge/ 

discharge controls. Not only will HM monitor battery health it will keep the system apprised of cell level 

changes and in some instances provide the capability to isolate a faulty cell while keeping the remaining 

cells safe and in operations. PMAD control can use power converters to bridge the lost cell and boost 

output voltage making the fault transparent to the user. 

(b) Cylindrical Cells (a) Prismatic Cells 

Figure 13.—Common prismatic and cylindrical cell geometry. (a) Prismatic cells tend to pack more efficiently in 

terms of volume but need additional structure to constrain the internal pressure in high vacuum space 

applications. [NASA Photo] (b) Cylindrical cell geometry provides a suitable pressure vessel with minimum 

additional packaging. [Photo Permission: Saft Specialty Batteries] (Ref. 24).
NASA/TM—2013-217813 23 

9.4 Battery Scaling Increment 

The battery beginning-of-life Ah capacity requirements (including redundancy and the extra margin 

required to meet end-of-life requirements) of various missions considered by the AMPS program sorted into 

two major types. Several missions had Ah requirements ranging from 23 to 76 Ah; others had Ah 

requirements ranging from 126 to 614 Ah. Peak discharge current requirements ranged from C/32 to 1.26 C. 

All missions had a requirement to support a 120 V bus. This range of requirements was met by designing 

two module types using lithium-ion cells designed to handle the appropriate current: a 27 Ah module and a 

150 Ah module, both of which operate at approximately 120 V shown in Figure 14. These two sizes of 

modules could be used in integral numbers to meet the requirements of all of the missions under 

consideration. Three types of cells available from current battery suppliers were considered in conceptual 

designs: larger capacity cylindrical cells produced by SAFT, prismatic cells produced by Yardney, and 

commercially available small capacity cylindrical 18650 cells (Ref. 21). After comparison of the mass and 

the complexity of control of the designs, larger capacity cylindrical cells were chosen for the conceptual 

design. The 27 Ah module uses 33 SAFT VES 100 cells in series (Ref. 22). The 150 Ah module consists of 

33 “virtual cells” in series, where the “virtual cell” consists of 3 SAFT VES 180 cells in parallel (Ref. 23). 

9.5 Modular Battery Logistics Benefits 

A common modular battery also reduces recurring logistics cost. If every spacecraft in a given 

mission uses the same types of modular battery then they could share a common source of flight spares. 

This reduces the overall spares inventory and improves supportability. Modular batteries would be 

primary targets for hardware salvaging operations. For example, modular batteries could be salvaged 

from an expended CPS or SEP vehicle and reused as spares or reallocated to another system. 

10.0 Modular Fuel Cell Technology 

10.1 Key Fuel Cell Performance Parameters 

The factors that drive the design of a fuel cell power plant are primarily mission driven. The peak and 

nominal power required defines the size of the fuel cell stack and balance of plant. This requirement 

defines not only the number of cells and/or stacks to be included, but also the size of the cells or stacks. 

27 Ah Modular Cell String 

150 Ah Modular Cell String 

Figure 14.—Conceptual drawings of 27 and 150 Ah. Two modular concepts at scaling increments of 27 and 

150 Ah were defined and used in cost modeling. [Credit: Kathleen Sukel, Vantage Partners, LLC]. 
NASA/TM—2013-217813 24 

Typically, a power-plant is characterized by the specific power (W/kg) and power density (W/l) of the 

power-plant as a whole. The voltage to be delivered to the vehicle bus by the fuel cell power-plant also 

impacts the design of the fuel cell stack. Higher voltages require a larger number of cells and/or stacks 

and how those cells or stacks will be arranged, i.e., series/parallel arrangements. 

Total system energy, in Watt-hours (Wh), is driven by mission power level and mission duration. In 

long duration missions, the system mass is dominated by the mass of the reactants and storage tanks. The 

fuel cell hardware may represent a relatively small fraction of the mass. 

Specific power (kW/kg), system efficiency, power density, desired peak to nominal power delivery 

ratio are typical parameters imposed by the mission. Fuel cells, however, impose their own requirements 

on the vehicle including, reactant pressure and flow rate, reactant purity, and heat loads handled by the 

vehicle thermal system. Future missions beyond Earth orbit will have additional requirements such as fuel 

flexibility. 

The flexibility of a fuel cell technology depends on its ability to utilize available reactants. This 

includes reactants scavenged from propulsion systems or reactants extracted from in-situ sources. Where 

some designs require pure hydrogen/oxygen, others provide flexibility to operate on relatively impure 

reactants or hydrocarbon fuels. 

Reliability and redundancy requirements also affect the design of the power-plant as a whole. Higher 

required reliability not only impacts how many redundant components are included within the powerplant design but also how those redundant components are handled, i.e., actively operating at all times 

within the power-plant or unpowered but available on a stand-by basis to take over a key function within 

the power-plant. 

Vehicles and their Design Reference Mission requirements influenced the selection of fuel cell 

technology for the Modular Fuel Cell. The HAT study vehicles were predominately solar array + battery 

architectures and did not include fuel cells. The AMPS project determined that this approach did not 

account for opportunities to exploit in-situ resources found at surface destinations such as the Moon and 

Mars. Therefore, the Lunar Lander Earth Orbit Rendezvous and the MMSEV Rover were deemed as 

likely users of Modular Fuel Cell technology. This narrows the mission set and elevates the importance of 

flexibility, in terms of, reactant types and reactant sources. 

10.2 Basic Fuel Cell Concepts 

Fuel Cells are electrochemical devices, similar to batteries, which convert chemical energy to 

electricity. They use a fuel, such as hydrogen and oxygen to produce electrical power. The fuel cell 

reaction also creates water and heat as by-products. 

The electrochemical reaction that produces power takes place within the cell stack but all the 

supporting processes are provided by the remaining “balance of plant” (BOP). The balance of plant is 

responsible for any preconditioning of reactants, moving reactants into the stack, removing by-products 

(water) and removing waste heat. The balance of plant interfaces with the vehicle to access stored 

reactants, transfer waste heat, vent reactants, and move product water out of the power-plant to be used by 

the crew, vented overboard, or reused in a closed loop regenerative system. 

Like batteries, fuel cells can be primary (non-rechargeable) or regenerative (rechargeable). Primary 

fuel cells, as can be seen in Figure 15, are composed of a fuel cell stack which produces the electrical 

power and a balance of plant which delivers the reactants to the fuel cell stack and removes the product 

water and waste heat. 

The voltage and current produced by a fuel cell is determined by the design’s cell stack series and/or 

parallel arrangement and the size of the individual cells. Unlike batteries, the fuel cell will continue to 

provide power as long as fuel and oxidant continue to be fed into the fuel cell stack. Unlike solar arrays, 

the compact and rugged fuel cells are suitable for underwater vehicles, aircraft, and automotive 

applications. 
NASA/TM—2013-217813 25 

Solar arrays and fuel cells are often seen as alternative power sources. Solar is seen as renewable 

while fuel cells are limited by reactant supplies. A hybrid system composed of both fuel cells and solar 

arrays can exploit the benefits of both. In a Regenerative fuel cell system, as shown in Figure 16, the 

reactants and water are part of a closed loop system where the water is reconverted back to reactants by a 

solar powered electrolyzer. This is similar to rechargeable batteries except that the regeneration can occur 

concurrently while the fuel cell continues to produce power without diminishing as long as reactant 

production stays ahead of consumption. 

Figure 15.—Primary fuel cell block diagram. 

Figure 16.—Regenerative fuel cell system block diagram. 
NASA/TM—2013-217813 26 

10.3 Fuel Cell Chemistry and Balance of Plant Options 

Only three types of fuel cell chemistries have actively been investigated for space applications, 

specifically, Alkaline, Proton Exchange Membrane (PEM) and Solid Oxide Fuel Cells (SOFC). Alkaline 

fuel cells have been the workhorse power source from Apollo to Shuttle Orbiter. Concerns with usable 

life, cost and sensitivity to contaminants, have driven the investigation into other options. 

PEM fuel cells operate at relatively low temperatures (80 C) and can bootstrap themselves to 

operational status as needed. PEM fuel cells typically operate using hydrogen as a fuel (although they 

have been shown to operate on methanol/air) and are generally intolerant of most reactant impurities. 

Three basic types of supporting Balance of Plant (BOP) designs have been under investigation for PEM 

fuel cell systems, active, passive and non-flow thru. 

SOFC operate at higher temperatures (600 C) and need to be pre-heated to that temperature before the 

fuel cell reaction can begin. Like PEM and alkaline fuel cells SOFC’s can operate directly on hydrogen and 

air or oxygen, but also operate with impure hydrogen (as would be expected from ISRU operations), or 

reformate from hydrocarbon fuels and methane. SOFC’s are also very tolerant of most impurities while the 

waste heat from SOFC’s is high enough to support cogeneration (via Stirling, turbines, etc.) for additional 

power and potentially increasing the overall system efficiency to greater than 70 percent. 

10.4 Modular Fuel Cell Concept and Scaling Increment 

Fuel cell power-plants have traditionally been designed incorporating one fuel cell stack with one 

balance of plant. The BOP delivers preconditioned reactants to the stack, removes water and waste heat. 

The BOP provides the interfaces with the vehicle reactant stores, thermal control, vent lines, and external 

water processing. The total power-plant is normally scaled to meet the entire power requirement of the 

vehicle. System redundancy is provided by additional fuel cell power-plants operating in parallel as 

shown in Figure 17(a) to (c). 

Figure 17.—Potential approaches to fuel cell system architecture. 
NASA/TM—2013-217813 27 

Figure 18.—Fuel cell stack and balance of plant mass with increasing power. 

Sized to meet the application power requirements the fuel cell stack scales linearly by adding cells 

and/or rescaling the basic cell size. An alternative method is to expand the power-plant by adding 

additional stacks. The power-plant has a variable output and thus, BOP fluid lines and control 

components are sized to cover a range of flow rates. 

Components are sized for maximum flow with margin. Beyond that the lines and controls jump to the 

next commercially available size step. As a result the BOP would follow a continuous increase in power 

with a stepwise increase in component size as shown in Figure 18. Therefore a given balance of plant 

design can typically accommodate a range of power levels about some nominal value before stepping up 

to the next size increment. 

The proposed modular fuel cell handles redundancy for the BOP and the modular cell stack with 

distinctly different approaches. Further, the sizing of the BOP and the modular cell stack differ. At first 

look, it may seem advantageous to size one power-plant for the largest vehicle power demand needed, 

however this may levy a large mass penalty upon the vehicle. For example, a balance of plant sized to 

deliver sufficient reactants with matched thermal rejection to handle 12 kW may be sufficient to handle 

the potential growth of vehicle power requirements. For higher power requirements, instead of adding 

more 12 kW Fuel Cell subsystem, a BOP with extra capacity can be built and separate modular fuel cell 

stacks added. For this same example, the attached modular fuel cell stacks are arranged as three 4 kW 

units to handle the anticipated power requirements. Additional 4 kW units can be added incrementally. 

For redundancy considerations, the three primary 4 kW stacks operate in parallel and each supports 

33 percent of the 12 kW total output. If one unit fails the remaining two units can be ramped up from 

33 percent total load to 50 percent of total load each. This is an increase of roughly 1.5 times and well 

within a fuel cell stacks range. The power thermal control interface will need to provide the additional 

heat rejection margin. 

However, if two of the three fail the remaining single stack would need to boost output by 3 times to 

carry the full load. This is much more difficult because there would be an attending decrease in efficiency 

and increase in waste heat. It is more reasonable to push the single stack output up between 1.5 to 2 times 

nominal and let the overall vehicle power drop to a level between 50 to 66 percent. 
NASA/TM—2013-217813 28 

This could be a temporary situation that can be handled by a so called Hot (Standby) Spare. This 

would be a 4th unit that sits in a standby condition and is started by the BOP rerouting fluids from the 

faulty unit to the hot spare unit. This assures that even after 2 faults the system still provides full capacity 

without overstressing the hardware. The hot spare unit is another 33 percent more mass plus the 

additional isolation valves, fluid lines, and controls. 

Alternatively, a modular cell stack “cold spare” approach can be used if the vehicle can tolerate the 

reduced capacity temporarily. In this case, spare module can be salvaged from another vehicle that has 

been expended but still contains functioning fuel cell hardware. This alternative eliminates the mass 

penalty of a hot spare cell stack. 

10.5 Modular Fuel Cell Technology Development and Demonstration 

A spacecraft fuel cell is literally an electrochemical power plant that often runs on the fuel/oxidizer 

reactants that it taps from the propulsion system. Their innate complexity makes them a challenge to 

integrate into a modular package. Modularization of a fuel cell power-plant begins with packaging the 

modular components to simplify the interfaces. In the case of the Non-Flow-Through (NFT) Fuel Cell 

technology, the process of separating the water from the oxygen is done within the fuel cell stack rather 

than within the balance of plant, potentially simplifying the interface between stack and balance of plant. 

NASA Glenn Research Center has been working with Infinity Fuel Cell and Hydrogen, Inc. to develop 

non-flow-through (NFT) proton-exchange-membrane (PEM) fuel cell power systems. The goal of the 

AMPS fuel cell demonstrations so far has to evaluate the technology for robotic and human rover vehicles. 

10.5.1 NFT Fuel Cells SCARAB Demo

In the first Fuel Cell demonstration within the AMPS program, the NFT fuel cell technology was 

being integrated onto surface system demonstration vehicles as shown in Figure 19. A 16-cell, 132 W 

developmental NFT fuel cell power-plant was used to augment the power delivered to the Carnegie 

Mellon University’s SCARAB rover (Ref. 25). This modest developmental test was intended to 

demonstrate a NFT fuel cell system on a test vehicle as a precursor to larger demonstrations. The 

successful tests performed in late 2011 helped to identify integration issues and evaluate features that in 

turn can improve the NFT Fuel design for future vehicle operations. 

10.5.2 3-kW NFT Fuel Cells MMSEV Demonstration 

The goal of the NFT Fuel Cell technology development is to develop and produce units of 

progressively higher capability. The next steps are 1 and 3 kW units. The 1 kW NFT unit is currently 

undergoing laboratory evaluation testing. The new design integrates parts of the fuel cell “balance of 

plant” hardware onto the stack interface plates, as can be seen in the power-plant mockup in Figure 20. A 

3 kW NFT Fuel Cell unit is being developed that is expected to match the power the core MMSEV 

vehicle loads, (currently, estimated at approximately 3 kW). A demonstration of this unit onboard the 

MMSEV prototype is planned in FY 13. The 3 kW NFT Fuel Cell developed by Infinity Fuel Cell and 

Hydrogen, Inc. is composed of 144 cells with a nominal stack voltage of 120 Vdc to produce a nominal 

stack power level of 3 kW and a peak level of 6 kW. 
NASA/TM—2013-217813 29 

Figure 19.—Carnegie-Mellon University’s SCARAB Rover with 

NFT fuel cell technology installed. [NASA Photo]. 

Figure 20.—3 kW non-flow through fuel cell power-plant mock-up. 

This mockup unit illustrates the integration of the balance of plant 

onto the fuel cell stack itself. This approach could enable a more 

compact PEM fuel cell power-plants. [NASA Photo]. 

11.0 Health Management and Supportability Technology 

The effectiveness of a Modular Power System depends on the successful development and application 

of Health Management and Supportability technologies. Normally, crew and ground operations, logistics, 

and specialized equipment are required to maintain a system and assure a high degree of system 

availability. Health Management and Supportability are tightly coupled to provide embedded solution to 

system health and maintenance. 

11.1 Supportability Dependency on Health Management 

The long term operational support of space flight systems has become an important program cost 

issue in an era of long lived space facilities such as the ISS (Ref. 4). Until recently, ISS had the benefit of 

a robust logistics infrastructure provided by the Space Shuttle. The Space Shuttle allowed the program to 

move Orbital Replacement Units between ISS and ground based depot on a regular basis. The ORUs were 

replaced by the crew manually or by robotic means and the detailed “root cause” diagnostic and repairs 

were off loaded to the depots. Note that without isolating a problem’s root cause, we cannot improve the 

hardware reliability. 

NFT Fuel 

Cell Stack 
NASA/TM—2013-217813 30 

Without the benefits of a logistics infrastructure, long missions beyond low Earth orbit have a much 

greater need to carry spares and related repair and test equipment to perform repairs in-situ. This collides 

with the extremely scarce upmass capacity. 

As discussed in Section 4.0 there is substantial mass savings achieved by sparing at levels below the 

complete ORU (Ref. 8). There is, however, a greater dependency on external repair, diagnostic and test 

equipment particularly, when searching for the root cause of a problem. Not only is there an uncertain 

mass impact, but there can be a substantial growth in the amount of crew time dedicated to maintenance. 

Based on the Lunar Surface Systems Supportability Study (Ref. 4) the time required to perform a repair 

grows 5 to 10 fold depending on the level of replacement. As noted in Section 6.0, roughly 80 percent of 

the repair process time involves diagnostic, integration and post repair functional tests. Proper modular 

encapsulation that simplifies interfaces to minimize violation will be a major step toward supportability. 

However, this must be matched with greater insight into root causes with embedded diagnostic and 

prognostic capabilities. 

A negative consequence of integrating functions into modular blocks is the tendency to physically 

hide details of the internal functions. This makes it very difficult to diagnose the root cause of problems. 

Measurement ports and test points are added but they increase complexity and the number of intrusive 

interfaces. There is thus greater demand for a built in test (BIT) to evaluate system functions. Simple 

functional BIT tests, however, do not indicate the root cause of a problem they only indicate what is not 

working. Embedded health management and supportability becomes enabling technologies for increasing 

system supportability while minimizing external equipment and operational complexity. 

Embedded Health Management (HM) exploits sensors and measurements that are already in the 

system usually as part of internal control loops. HM typically looks for slowly progressive changes in 

control loop signals that indicate the onset of a fault. This early warning allows the crew to prepare for, 

and where possible, preempt the fault with corrective actions. This may include changing operations to 

minimize degradation or switching hardware to channels with lighter loads. The early warning also lets 

the crew investigate behaviors that further identify the root cause. Further HM will monitor and analyze 

sensor signals for inconsistencies between with related sensors to determine if sensor rather than the 

hardware is faulty. 

 

Figure 21.—AMPS health management. Health management spans the 

range of hardware from vehicle level down to component level. 
NASA/TM—2013-217813 31 

These features also apply to scenarios where hardware is salvaged and reallocated. They are useful in 

assessing the health of the salvaged unit and the remaining operational life. These features can 

dramatically improve the effectiveness of modular power as a supportable system. 

11.2 Health Management and Supportability Performance Parameters 

Health Management and Supportability will be evaluated by its ability to: 

 Detect the onset of a fault sufficiently in advanced to allow the crew to preempt the fault. 

 Predict the time remaining before failure 

 Pinpoint the fault to individual components. 

 Determine the root cause in sufficient detail to prevent a recurrence. 

 Identify, disqualify and compensate for faulty sensors. 

 Assess the remaining life of any element 

 Embed an “electronic log book” to store unique information such as, module ID, current configuration, 

operating hours, estimated remaining life 

11.3 Health Management and Supportability Concept 

In order for a modular power system to deliver maximum benefit, health management functions will 

need to be integrated with the overall system control functions in the form of a hierarchical and 

distributed architecture shown in Figure 21. Many health management functions can be delegated to 

lower levels of assembly closer to the potential fault sources. This architecture makes use of intelligent 

sensors and components with embedded processing capability that makes it possible for health 

management to function at lower levels and subsequently reported up the hierarchy. In addition, 

distributed and hierarchical monitoring, diagnostic and prognostic approaches will allow the intelligent 

sensors and components to incorporate local knowledge and history to enable local health assessment. 

The locally assessed health knowledge then drives higher-level health management models that 

incorporate high-level knowledge (e.g., analytical and/or empirical models) that allows the integration of 

lower-level knowledge through interdependency relationships and allow consistency checks of lowerlevel data. As an example, intelligent sensors can actively determine their own condition and provide an 

indication of the quality of the sensed data that they are providing. This additional health assessment 

information can then be included in the evaluation and decision-making by higher functions in the 

hierarchy. Another benefit of distributed intelligent components is the potential for collaborative 

processing of complex algorithms in parallel. Modular and distributed HM enables increased flexibility, 

portability and reusability of the modular components. 

11.4 Health Management Development Framework 

The distributed architecture for health management and system level-control will be achieved through 

the development of a software framework that enables the flexible organization and structuring of all 

required health management, supportability and control functions. This framework will provide a scheme 

for encapsulation of the knowledge, behavior and interactions associated with the system. The framework 

is intended to enable developers to create health management and supportability software objects that will 

run on systems composed of distributed heterogeneous processors (e.g., embedded processors, 

microcontrollers, smart sensors and smart components). It will accommodate a range of processing 

capabilities including processors that need, so called, “light-weight” implementations due to 

computational limitations. In addition, the framework will allow for the definition of standard interfaces 

enabling communication of inter-process data and commands between distributed elements. To assure 

interoperability, the framework will be based on open system industry standards for data interchange. 
NASA/TM—2013-217813 32 

To minimize growth in complexity, the framework will emphasize code re-usability at all levels of the 

architecture. It provides the flexibility to implement and execute health management, control and 

supportability functions throughout the hierarchy from vehicle system-level control, decision-making and 

fault management all the way down to low-level monitoring, diagnostic, prognostic and remediation 

response processes. Plug-in libraries enable the user to select and use a wide variety of control, decisionmaking, and HM algorithms, various system models, built in test (BIT), and self-contained analysis routines. 

The framework will accommodate advanced diagnostic and prognostic algorithms based on analytical 

system models, empirical data acquired from hardware characterization tests, or hybrid combination of 

analytical and empirical techniques. Additionally, the framework will accommodate advanced control 

algorithms such as “agent-based” controls and power/load flow optimization. Verification of the 

distributed architecture that results from this framework will be accomplished through implementation 

and demonstration on the AMPS project’s PMAD development breadboards. 

12.0 Conclusion 

For the Advanced Exploration Systems (AES) Modular Power Systems (AMPS) project at GRC the 

goal is to develop a modular power architecture composed of technologies for power generation, energy 

storage, power distribution and health management that will reduce the cost of future space systems. 

Examining the several new vehicles required for future missions, there is a clear opportunity to 

consolidate the power system development into common power architecture and reduce development 

cost. In addition, the vehicle recurring and life cycle cost is reduced because the common set of spares, 

common equipment and common procedures reduces integration complexity and logistics. Modular 

power is believed to enable spacecraft to be reusable, serve multiple missions, and contribute hardware to 

a growing space infrastructure. 

Commonality is necessary but not sufficient for defining modular power systems. To assure 

flexibility and scalability a modular approach must incorporate standardized smart interfaces that provide 

interoperability that allow the modules to be reused and reconfigured to meet changing needs. To manage 

this flexibility and keep hardware organized, each module will need embedded supportability and health 

management features that preserve the module identity, their current configuration, operating history, and 

cumulative operating life. Supportability and Health Management will assist the crew in predicting faults, 

diagnosing and isolating problems, and quickly restoring full capability with minimum complexity. 

In the next few years modular power system technologies along with health management and 

supportability technologies will continue to evolve. Demonstrations are expected to provide validation of 

the AMPS modular power concept as an essential element in the development of multi-mission 

exploration vehicles. 

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PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-06-2013 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Advanced Modular Power Approach to Affordable, Supportable Space Systems 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Oeftering, Richard, C.; Kimnach, Greg, L.; Fincannon, James; Mckissock, Barbara, I.; 

Loyselle, Patricia, L.; Wong, Edmond 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 756414.02.02 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-18549 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2013-217813 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 12 and 20 

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

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Recent studies of missions to the Moon, Mars and Near Earth Asteroids (NEA) indicate that these missions often involve several distinct 

separately launched vehicles that must ultimately be integrated together in-flight and operate as one unit. Therefore, it is important to see 

these vehicles as elements of a larger segmented spacecraft rather than separate spacecraft flying in formation. The evolution of large multivehicle exploration architecture creates the need (and opportunity) to establish a global power architecture that is common across all 

vehicles. The Advanced Exploration Systems (AES) Modular Power System (AMPS) project managed by NASA Glenn Research Center 

(GRC) is aimed at establishing the modular power system architecture that will enable power systems to be built from a common set of 

modular building blocks. The project is developing, demonstrating and evaluating key modular power technologies that are expected to 

minimize non-recurring development costs, reduce recurring integration costs, as well as, mission operational and support costs. Further, 

modular power is expected to enhance mission flexibility, vehicle reliability, scalability and overall mission supportability. The AMPS 

project not only supports multi-vehicle architectures but should enable multi-mission capability as well. The AMPS technology 

development involves near term demonstrations involving developmental prototype vehicles and field demonstrations. These operational 

demonstrations not only serve as a means of evaluating modular technology but also provide feedback to developers that assure that they 

progress toward truly flexible and operationally supportable modular power architecture.

15. SUBJECT TERMS

Logistics; Flight operations; Space architecture; Multi-mission spacecraft; Space power; Solar array; Batteries; Fuel cells 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

40 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18