August 2012 

NASA/CR–2012-217758 

Manned GEO Satellite Servicing Mission 

Environmental Effects Measurements Study 

Douglas G. Murphy and Shaun A. Deacon 

Analytical Mechanics Associates, Inc., Hampton, Virginia 
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August 2012 

NASA/CR–2012-217758 

Manned GEO Satellite Servicing Mission 

Environmental Effects Measurements Study 

Douglas G. Murphy and Shaun A. Deacon 

Analytical Mechanics Associates, Inc., Hampton, Virginia 
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Manned GEO Satellite Servicing Mission

Environmental Effects Measurements Study

National Aeronautics and Space Administration
Manned GEO Satellite Servicing Mission 

Environmental Effects Measurements Study 
MGS Environmental Effects Measurements Study      Page | 3

Contents

Introduction .................................................................................................................................... 5

Study Motivation ............................................................................................................................ 6

Study Process.................................................................................................................................. 8

Prioritized Environment Data Needs.............................................................................................. 9

Existing and Near Term Space Environment Missions ................................................................. 10

Instrument Review........................................................................................................................ 12

Bus Review.................................................................................................................................... 18

Hosted Payload.......................................................................................................................... 18

Robotic Servicing Mission ................................................................................................. 19

Commercial Communications Satellites........................................................................... 21

Secondary Payload .................................................................................................................... 22

Standard Secondary Payload ............................................................................................ 22

The EELV Secondary Payload Adapter (ESPA) Ring........................................................... 24

ESPA Small Satellite Payload............................................................................................. 25

LCROSS Derived ESPA Bus................................................................................................. 26

Primary Payload ........................................................................................................................ 27

Bus and Instrument Trade Space.................................................................................................. 29

Conclusion..................................................................................................................................... 33

Acknowledgements....................................................................................................................... 34

REFERENCES .................................................................................................................................. 35

APPENDIX A:   Instrument Comparisons..................................................................................... 38

APPENDIX B:   Instrument Data .................................................................................................. 41

APPENDIX C:   Acronyms............................................................................................................. 61
MGS Environmental Effects Measurements Study      Page | 4

Figures

Figure 1. GEO environment study within the MGS team organization.......................................... 5

Figure 2.  Map of the AE‐8 MAX integral electron flux >1 MeV. (SPENVIS User Guide) ................ 7

Figure 3.  RBSP orbit and the Van Allen radiation belts. .............................................................. 10

Figure 4.  DSX energetic particle coverage. .................................................................................. 11

Figure 5.  The DSX satellite, stowed.............................................................................................. 11

Figure 6.  Gallery of modern environment instrumentation.  ...................................................... 13

Figure 7.  ESPA configuration diagram from the ESPA User’s Guide............................................ 24

Figure 8.  ESPA ring. ...................................................................................................................... 25

Figure 9.  A spacecraft built using the ESPA Standard Interface Vehicle (SIV). ............................ 25

Figure 10. LCROSS ......................................................................................................................... 26

Figure 11. Notional powered ESPA ring........................................................................................ 26

Figure 12. MGS environment study trade tree............................................................................. 32

Tables

Table 1.  Prioritized GEO Environment Data Needs........................................................................ 9

Table 2.  Instrument Review General Findings............................................................................. 13

Table 3.  Instrument Review Figures of Merit .............................................................................. 15

Table 4.  MEES Option 1 “State of the Art, Maximum Capability” ............................................... 16

Table 5.  MEES Option 2 “Minimum Development, High Capability” .......................................... 17

Table 6.  MEES Option 3 “Low Cost, Reduced Resolution” .......................................................... 17

Table 7.  MEES Option 4 “Size Constrained”................................................................................. 18

Table 8.  GEO Transportation Options Figures of Merit............................................................... 27

Table 9.  Bus Options Evaluation .................................................................................................. 28

Table 10.  Instrument Suite Figures of Merit................................................................................ 29

Table 11.  MEES Compatibility with Bus Options.......................................................................... 30

Table 12.  MEES Configuration Evaluations.................................................................................. 30

   
MGS Environmental Effects Measurements Study      Page | 5

Introduction

This report summarizes a trade study that was conducted to examine options for a flight

experiment designed to collect space environment and effects data in Geosynchronous Earth

Orbit (GEO).   This effort was a part of the joint NASA/DARPA Manned GEO Satellite Servicing

(MGS) study, whose goal was to identify potential architectures, required technologies, and

near‐term demonstration missions needed to develop a capability to service GEO satellites.  

The role of the environment study team within the larger MGS project organization is shown in

Figure 1.

Figure 1. GEO environment study within the MGS team organization.

The aim of this environment study was to provide recommendations to the Human Presence

Requirements Team (“Team 3”) on the best approach for collecting GEO environment data that

will be needed to inform the design of the equipment for a manned GEO mission, including the

manned servicing module and suits for extravehicular activity (EVA).    The study began by

reviewing the existing set of environment and effects measurement priorities identified in

previous work by Team 3 and comparing those priorities against known, near term flight

environment studies that are already being planned by NASA or other agencies.  This evaluation

determined that no currently planned environment mission will fully meet the need.

Next, information was collected on the availability and applicability of potential flight

measurement instruments in order to understand the rough cost, risk and schedule parameters

associated with development of an MGS environment instrument suite.    Four notional

instrument suites were developed and coarsely evaluated using weighted figures‐of‐merit

(FOMs) developed in conjunction with radiation subject matter experts on the MGS team.    
MGS Environmental Effects Measurements Study      Page | 6

A parallel effort examined the available transportation opportunities to GEO, including three

main options: 1) hosted payload; 2) secondary payload; and 3) primary payload.  Each of the

notional instrument suites were also scored in terms of compatibility FOMs with each of the

payload options.    The results of the transportation evaluation, including compatibility of the

four notional instrument suites, identified the hosted payload as the most promising

transportation option.

The study concluded with the development of a trade tree identifying the transportation and

notional MGS environment and effects measurement suite options.    Preliminary conclusions

drawn from the evaluation of each leg in this trade tree are included in this report.  

Study Motivation  

The radiation environment hazards in GEO include energetic electrons, Galactic Cosmic

Radiation (GCR) and Solar Particle Events (SPEs).  GCR and SPE hazards exist in GEO as well as

deep space, while the electron hazard is unique to Earth orbit.  GEO lies in the outer radiation

belt just beyond the peak intensity of energetic electrons trapped in the Earth’s magnetic field

(Figure 2). The high energy electrons in this region can penetrate materials and produce

internal charging. The low energy electrons can quickly produce significant (kilovolts) surface

charges. Damage from charging occurs when the potential difference leads to a discharge

event, which can include high currents and arcing.  Further, high energy electrons pose a direct

harm to human tissue by depositing energy into cells. Because of these effects, electron

radiation exposure is a key contributor to mission duration and EVA limits.  

Key Findings of the MGS Environment Measurements Study

 There is a need to collect environment & effects data at GEO to prepare for a

manned‐servicing mission.

 A suitable science‐quality instrument suite can be developed based on existing

technologies evolved from flight‐heritage components.

 Our desired instrument suite is well suited for a hosted payload option, but can

be made compatible with a small secondary bus of the types delivered using

the EELV Secondary Payload Adapter (ESPA).

 A dedicated launch to GEO would be inefficient unless partners were identified

to fill the spacecraft capacity.

 Due to the availability of compact sensor technologies, our trade space contains

a number of feasible fallback options.
MGS Environmental Effects Measurements Study      Page | 7

Figure 2.  Map of the AE‐8 MAX integral electron flux >1 MeV. (SPENVIS User Guide)

Because of the utility of GEO, this regime has always been of interest to spacecraft designers.

Mitigation strategies for the harms of the GEO environment, provided by NASA and others, are

therefore well developed for unmanned missions (though serious malfunctions of GEO assets

attributed to environmental events still occur) (Reedy 1998, NASA‐HDBK‐4002A). These

strategies, which include shielding of electronic components and Faraday cage construction, do

not easily meet the needs of a manned mission with EVAs. An ergonomic suit, for example,

cannot be made to any desired material thickness.  Further, the current standard model of the

space environment known as “AE‐8” (Vette 1991) has shortcomings for applications in GEO,

and is therefore not ideal for developing requirements for an EVA suit design. The following

caution, for example, appears in the Space Environment Information System (SPENVIS) user

guide: “The AE‐8 model has been criticized also for its deficiency in high energy electrons at

geosynchronous orbit (Baker et al., 1986). In fact, above 2 MeV, the model is just an

extrapolation of unknown validity.” An updated model (AE‐9) is in development, but like AE‐8 is

not specifically aimed at meeting the needs of a manned GEO mission.

Measurement of the effects caused by the environmental hazards is a separate but also high

priority for MGS.  Without correlating the environmental effects (e.g. charging and radiation) in

a like material or simulated tissue, the utility of the models for developing the equipment

requirements is limited.    Such correlating effects measurements are ideally done in situ and

collocated with the environment instruments on the same instrument suite (Koons 2004).  
MGS Environmental Effects Measurements Study      Page | 8

Because the characteristics of the Earth’s magnetosphere vary in time and space, separate

missions in GEO may be experiencing different effects (Mazur 2007). Ground‐based tests may

also not meet the need, since it is difficult to account for the simultaneous effects of a range of

particle species, energies and flux characteristics. While ground testing of instruments will be

conducted as a matter of course to qualify and calibrate the instruments, the addition of tests

to thoroughly simulate the GEO environment may incur a significant cost.  Further, the effects

instrumentation reviewed for this study was generally low cost (compared to the environment

suite) and of miniature design, in some cases being of a “patch” type requiring no more space

than a paper napkin and utilizing only milliwatts of power. For this reason, each of the proposed

notional instrument suites include effects instrumentation to correlate the environment

measurements rather than relying on ground‐based testing alone.

Study Process

The study was divided into two main tasks:    1) trade space definition and 2) trade space

analysis. In task 1, over 70 instruments from more than 30 previous flight missions were

identified and categorized by instrument type, measurement parameter range, orbit regime

and launch date. Descriptions of many of the instruments considered during this study appear

in Appendix B. Several instrument development efforts were also included in this survey. Data

for these instruments was gathered primarily from publicly available literature and web pages.  

Next, a downselect process was used to identify current generation instruments most

applicable to the needs of MGS environment study.   Data collected for this smaller set included

cost, mass, power and volume specifications and a coarse evaluation of the development

required to modify the system for use in GEO (for existing hardware) or to complete

development to a flight‐ready status. For this effort, interviews with the instrument principal

investigators (PI’s) augmented the literature search.    The coarse assessment of the

modifications required to adapt a given instrument was based largely on the current operating

regime of the instrument, as well as input from the PI’s.   Detailed reviews of the instrument

designs were beyond the scope of this task.

Also under task 1, GEO transportation options were reviewed and categorized based on launch

vehicle, bus options, transportation availability (frequency of upcoming opportunities), and

approximate cost.    Both foreign and U.S. launch capabilities and standard bus options were

included in this survey.    Data on bus options, including the proposed NASA Reusable Earth

Synchronous Tele‐Operated Refueler (RESTOR) robotic servicing mission, was gathered both

from public sources and interviews with persons affiliated with the program.

The final part of task 1 was the development of FOMs to assess both the instrument options

and bus options.    FOMs were created to address the specific goal of building a flight

environment suite in the near term, with a target launch date near 2014.
MGS Environmental Effects Measurements Study      Page | 9

In task 2, a preliminary analysis of the defined MGS radiation environment and effects mission

trade space was performed, with resulting initial recommendations presented to the MGS

Team. The recommendations and key findings were presented at the MGS Team Meeting at

Johnson Space Center on September 19, 2011.   

Analysis of the trade options was performed using a series of matrices with each weighted FOM

scored for each trade option.  These matrices are included in Appendix A.  While this exercise

did stratify the instrument options and allowed high value approaches to be identified, the

results should be considered preliminary due to the limited time available during the study to

compare details of the instrument capabilities against the specific shortcomings of the existing

available GEO environment data sets and analytical models.    Such a detailed comparison is

suggested for a future activity.  The results of the preliminary trades do however suggest that a

number of good options are available for achieving the MGS environment mission goals with

existing instrument technologies and transportation options, and that no major technology

gaps would need to be overcome by designers of the flight instrument suite.

Prioritized Environment Data Needs

Table 1 below shows the prioritized data needs identified by radiation subject matter experts

on the MGS team at the start of the trade study.  The list represents the kinds of measurements

that would most benefit the understanding of the space environment specifically for a manned

mission to GEO.    This prioritization remained unchanged after an assessment of upcoming

environment missions was conducted. Note that an implicit desire for all of these

measurements is a continuous, prolonged sample (years) at the GEO orbit, within several

hundred kilometers of the GEO altitude of 35,786 km, zero eccentricity and within plus or

minus five degrees of inclination.

Table 1.  Prioritized GEO Environment Data Needs

Priority Measurement Benefit

1 High energy electron environment

50 keV to 10 MeV

Characterize radiation and internal charging

environments  

2

Low energy electron, ion environment

100 eV to 50 keV

Characterize surface charging environments

3

Radiation effects in tissue   Correlate harm to humans behind candidate suit

materials  

4

Electrostatic discharge effects

(especially during docking)

Understand discharge events at moment of

contact with client vehicle  
MGS Environmental Effects Measurements Study      Page | 10

5

Internal charging effects

(Deep dielectric charging)

Correlate charging to environment using candidate

suit materials  

6 Magnetic field environment Identify magnetosphere boundary and measure

pitch angle particle distributions  

7

Surface charging effects Monitor vehicle surface potential to validate

surface charging models  

The highest priority measurements are energetic electrons, where it is desired to know the

number (flux) and energy of these particles to the greatest resolution possible. Electrons are

divided into “High Energy” and “Low Energy” as measured in electron volts (eV) due to the

different spacecraft effects associated with each. In the notional instrument suite development

activity, the electron environment was divided into three categories, since it was found that

effective measurement of electron energy ranges above 4 MeV usually required a dedicated

instrument utilizing a different sensing technology than that appropriate for sensors looking at

electron energies below that value.   

The item of fourth priority on this list, electrostatic discharge (ESD) effects, is intended to

address both the well‐recognized arcing behavior seen on GEO satellites due to potential build‐

up on dielectric surfaces and the special situation where two satellites come into contact (e.g.  

a capture by a grapple fixture).  ESD due to a contact event between two spacecraft in GEO has

never been measured and constitutes a unique requirement for MGS.  The desire for this type

of data also informs the GEO transportation evaluation, since only the robotic servicing hosted

payload option can meet this need.  

Existing and Near Term Space Environment Missions

Understanding the space environment by making

in‐space measurements continues to be a priority

for the science and engineering communities, as

evidenced by the number and variety of

associated flight projects. Decades of data

collected from GEO assets operated by the Los

Alamos National Laboratory (LANL), the GOES

spacecraft operated by the National Atmospheric

and Oceanic Administration (NOAA), and others

has been used to inform the current standard

analytical models, including the soon to be

released AE‐9 and AP‐9 (Onsager 1996, Goka 2007, Reeves 2009).  A flagship NASA mission, the

Figure 3.  RBSP orbit and the Van Allen

radiation belts.
MGS Environmental Effects Measurements Study      Page | 11

Radiation Belt Storm Probe (RBSP), is being prepared for launch in 2012 and will sample the

radiation environment in an elliptical orbit spanning the Van Allen belts (Figure 3).  Results from

RBSP will be incorporated into the updates to AE(P)‐9 in 2015 (Guild 2009). The $150M

Demonstration and Science Experiments (DSX) mission, a joint NASA and Air Force project, will

perform a survey in Medium Earth Orbit (MEO) of a wide spectrum of energetic particles after

launch in 2012 (Schoenberg 2006). Figure 4 shows a map of the particle coverage for each of

the instruments on DSX.  Figure 5 shows a CAD model of DSX in the stowed configuration.   

In 2015, the planned flight of GOES‐R will fly the Magnetospheric Particle Sensor (MPS), the

next generation of NOAA energetic particle spectrometers in GEO.

                    

Figure 4.  DSX energetic particle coverage.    Figure 5.  The DSX satellite, stowed.

               

However, when these and other missions were compared to the MGS need, several gaps were

identified.  First, the desired range and resolution of energetic particle samples does not occur

for a continuous, prolonged period at GEO. The science‐quality instrumentation on RBSP, for

example, will only briefly pass through GEO, whereas the instruments on the GOES satellites

(and similar instruments on satellites operated by the Japan Aerospace Exploration

Agency/JAXA) at GEO have a limited range and resolution (see the Instrument Data in Appendix

B for basic reference data on various instruments surveyed).  An extended duration mission is

desired due to the dynamic nature of the magnetosphere, where influences such as the 11 year

solar cycle play a significant role (Reeves 2009).   

Second, there are few examples of collocated effects sensors at GEO. The DSX spacecraft,

which includes both environment and effects instrumentation, is an example of collocated

effects sensor suite but it operates at MEO.   A hosted payload experiment by the European

Space Agency (ESA) called the Alphasat Environment and Effects Suite (AEEF) was the only

mission identified by the study team that largely addressed this need in GEO. The AEEF is

scheduled for launch in 2013 and will be attached to an Alphasat communications satellite as a

technology demonstration.   
MGS Environmental Effects Measurements Study      Page | 12

Third, existing or planned efforts will not collect the data desired to specifically address the

needs of a manned satellite servicing mission.  Measurement of radiation dose in tissue, for

example, is not a feature of the Radiation‐Sensing Field‐Effect Transistor (RadFET) technology

used by the AEEF dosimeter, which is intended to characterize the effect of radiation on

electronics.    Another gap is in the resolution of highly energetic electrons above 4 MeV, a

region of relatively low flux which represents less of a concern for spacecraft electronics but

will be a hazard for EVA.    Achieving a good signal‐to‐noise ratio is a challenge for detector

design in this regime and has been a factor in the LANL high energy data.   A parallel task at

LANL is being undertaken to develop a process for reducing noise in the data that has already

been collected. The results of this effort, when available, should be used to re‐examine the

priorities in Table 1.   

Examples of discharge monitors in flight are also rare. Data collected by the Spacecraft Charging

at High Altitude (SCATHA) and Combined Release and Radiation Effects Satellite (CRRESS)

spacecraft has been the basis for much of the flight knowledge in this regime (Dyer 1996,

Adamo 1996), but neither addresses the effects likely to be important for the MGS EVA suit

material or the spacecraft‐to‐spacecraft contact event during the target capture.  

Instrument Review

The technologies for collecting environment measurements listed in Table 1 are generally

mature, and modern examples of instruments either with flight heritage or with a short (~1

year) development schedule were identified. Figure 6 shows a collection of instruments of the

type that were traded in this study (see Appendix B for more instrument data).    The largest

instruments in the suite are the particle spectrometers, which range from shoe‐box size to

small microwave ovens.    An exception is the Cosine Multi‐function Particle Spectrometer

(Cosine MPS), a commercially available product measuring roughly 10 cm on a side.  The Cosine

MPS uses a scintillator crystal to stratify and measure electrons, protons and ions over a wide

AEEF Measurements in GEO

• Incident electrons in the overall energy range ~50 keV ‐ 10 MeV

• Incident protons from ~4 MeV to >300 MeV

• Incident heavy ions in the Linear Energy Transfer (LET) range 15‐70 MeV/mg/cm2

• Total ionizing dose measurements utilizing RadFETs, sensitivity of up to 85 mV/rad
MGS Environmental Effects Measurements Study      Page | 13

range of energies (Lampridis 2008, Maddox 2008).  While the Cosine MPS has no flight heritage,

units have been qualified for flight on future ESA missions.   

The smallest instruments include several “patch” style sensors for charge and discharge

monitoring, including the Electronic Discharge Monitor (EDM) and Charge Deposit Sensor (CDS),

both under development at the Aerospace Corporation.  These patches measure several inches

on a side and are several millimeters thick, making them suitable for multiple mounting

locations over the spacecraft.   

Figure 6.  Gallery of modern environment instrumentation.

A review of the literature and interviews with PIs conducted during this study led to the further

general findings listed in Table 2 below.
MGS Environmental Effects Measurements Study      Page | 14

Table 2.  Instrument Review General Findings

Instrument Type General Findings

High energy spectrometer  

(>4 MeV)

 Few heritage instruments in this energy range

 Electron flux is low

 Requires solid‐state or scintillator technique, which may add noise

and resolution challenges

 Science quality instruments typically have high mass (10‐15 kg)

 RBSP instruments are state of the art but have limited field of view,

may be over‐shielded for GEO work  

Medium energy spectrometer    Good population of flight heritage instrumentation

 Electrostatic analyzer technique allows good resolution

 Science quality instruments  typically have high mass

 Multifunction “situation awareness quality” instruments have

coverage in this regime

Low energy spectrometer    Good population of flight heritage instrumentation

 May require multiple ‘sensor heads’ to gain wide field of view

coverage for non‐spinning bus (applies to all spectrometers)

Tissue Equivalent Proportional

Counter (TEPC) Dosimeter  

 Few instruments in this population, ISS focus

 ISS heritage instruments are high Technology Readiness Level (TRL)

 Compact TEPC in work at ARC

ESD Monitor  Few instruments in this population

 SCATHA (1979), CRRES (1990), APEX (1994) inform the current models

 Micro‐sensor from Aerospace Corp. in development  

Internal charging monitor  

(deep dielectric charging)  

 Several instruments in this population  

 Multifunction sensors have coverage in this regime

 Dedicated sensors available from Qinetiq, DPL science

Fluxgate Magnetometer  Widely available  

 Typically boom mounted, but body mount with calibration strategy

will remove some design challenges (e.g. available volume on a

smallsat) and bus constraints (e.g. hosted commercial payload may

not allow boom)

 Engage hosts early to examine spacecraft design issues that may

affect the magnetometer (e.g. solar wing wiring).  

Surface charging monitor  Several instruments in this population  

 Micro‐sensor from Aerospace Corp. in development

 Multifunction sensors have coverage in this regime

TID Silicon Dosimeter    Flight micro‐sensors commercially available. Lunar Reconnaissance

Orbiter (LRO) heritage. (Mazur 2011)
MGS Environmental Effects Measurements Study      Page | 15

The instruments included in the down‐selected set were compared using a coarse evaluation

against FOMs developed specifically for the MGS application.  Those FOMs are briefly described

in Table 3 below.

Table 3. Instrument Review Figures of Merit

FOM Importance Description/Comments

Spans Desired

Parameter Range High

Charged particle spectrometers with wide energy range

capabilities, for example, are highly rated in this FOM.     

High Measurement

Resolution High

Generally the instrument resolution capability is reduced for larger

parameter ranges. Science quality instruments are highly rated in

both categories but may carry a volume and weight penalty.  

Good Noise

Mitigation Medium

Aimed primarily at particle spectrometers; the low flux, high

energy regimes pose a challenge for mitigating noise, e.g., due to

bremsstrahlung radiation caused by particle collisions.

Latest Generation

Instrument Medium

More recently developed instruments were weighted highly under

the assumption that current technologies were incorporated. A

detailed evaluation of the technology was beyond the scope of the

study.

Low Cost Medium

Costs were estimated from various sources, including NASA

databases and interviews with PIs.

Good FOV Without

Spin Low

Applicable primarily to particle spectrometers; instruments with a

wide field‐of‐view were desirable since there is no guarantee that

the bus will be spinning (a feature relied upon by some sensors to

sweep across a large range of angles).

Available Soon Low

Instruments already flown or qualified for space flight were rated

highly.

Designed for GEO Low

Instruments with flight heritage at or near GEO were rated highly

under the assumption that a similar unit could be built from minor

modifications to the existing design.

Minimal Thermal

Requirements Low

Highly sensitive charged particle instruments may require thermal

stability to function at their specified performance.

Low Mass Low

The MGS instrument suite is not a mass intense payload, but lower

mass may still benefit launch costs in some situations.

Low Power Low

In the trade evaluations, power turned out not to be a

discriminating FOM. All instruments have minimal power needs.
MGS Environmental Effects Measurements Study      Page | 16

Low Volume Low

In general, particularly for the particle spectrometers, smaller

units had less desirable performance in parameter range and

resolution.

Design Life Low

In the trade evaluations, design turned out not to be a

discriminating FOM. All instruments have sufficient design life (> 2

years).

A preliminary comparison of the instruments in the down‐selected set was performed by

assigning numerical weights to the FOMs using the guidance from the “Importance” column in

Table 3.    Conversion of the low/medium/high scale to a numerical value was accomplished  

using engineering judgment for the purposes of the exercise. The tables used for the

comparison are shown in Appendix A.  

For the purposes of examining the trade space at an end‐to‐end level, selected instruments

from the tables shown in Appendix A were grouped into four notional experiment suites that

could serve the MGS need.    The notional experiment is referred to here as the “MGS

Environment and Effects Suite” (MEES).    The four configurations were designed to address

different driving priorities and were grouped into four options as shown in the summary tables

below (Tables 4 through Table 7).  

 Option One: State of the art, maximum capability

 Option Two: Minimum development, high capability

 Option Three: Low cost, reduced resolution

 Option Four: Size constrained

Table 4.  MEES Option 1 “State of the Art, Maximum Capability”

Option 1 utilizes the highest scoring instruments in parameter coverage and resolution, with

each measurement category having a dedicated sensor (no multifunction elements).  
MGS Environmental Effects Measurements Study      Page | 17

Note that the column “equivalent TRL” is used here as a coarse gauge of the viability of the

existing design for the purposes of the MGS application, and may not be equal to the TRL of the

instrument for its original intended purpose. Costs were estimated using available NASA

databases, interviews with the PIs, or comparison with known values from similar projects.

A command and data handling unit (C&DH) was included in each configuration as a placeholder

for the avionics and telemetry system that would be required to support the sensor suite.

Table 5.  MEES Option 2 “Minimum Development, High Capability”

Option 2 reduces development risk while sacrificing some capability. The GOES‐R spectrometers

replace the RBSP elements since the GOES systems are already designed for operation on a

body‐stabilized spacecraft in GEO.

Table 6.  MEES Option 3 “Low Cost, Reduced Resolution”
MGS Environmental Effects Measurements Study      Page | 18

Option 3 makes use of less precise instrumentation and multifunction systems such as the

commercially available “CEASE II” from Amptek. Such systems are typically used for spacecraft

health monitoring.

Table 7.  MEES Option 4 “Size Constrained”

In Option 4 miniature components are used where possible and some instruments are

removed.  

The four options presented here will be revisited after a review of the bus options in the next

section.  

Bus Review

The MGS environment measurement suite configurations identified in Tables 4 through 7 were

developed independently from the transportation options.  In this section of the report, those

transportation options are described and their evaluations summarized such that a trade of the

complete system may be conducted. Three transportation options to GEO were identified: 1)

Hosted Payload, 2) Secondary Mission, and 3) Primary Mission. Each of these options is

discussed below.

(1) Hosted Payload

The Hosted Payload category involves all of the options that rely on utilizing space on another

mission’s spacecraft. The spacecraft primary mission offers the use of excess platform capability

in power, mass, and volume at a cost that is much lower than typical costs for space access. The

cost of these accommodations depends directly on the schedule and resource impacts to the

primary mission. The lower the burden on the primary mission, the less expensive the resource

utilization becomes.

To increase the likelihood of low impact accommodations, interested parties typically come

together 24 to 36 months before the planned launch date. This allows time for both sides to
MGS Environmental Effects Measurements Study      Page | 19

handle design changes necessary to meet the larger shared requirement set. The actual time

required is typically defined by the development schedule of the primary mission, but can also

be impacted by specific needs of the hosted payload. Delivery of the final payload is normally

required at least 12 months before launch, to allow for adequate time for integration and test

with the primary mission (Andraschko 2011, Futron 2010).

For the purposes of this study, all hosted payloads can be divided into two categories: A)

robotic servicing missions and B) geosynchronous communication missions. While the basic

characteristics of these missions are similar, the distinguishing feature for this study is the

possible presence of a “contact event” between two spacecraft. In the robotic servicing

missions, there exists the possibility of an electrostatic discharge event just prior to contact

between two spacecraft.

A) Robotic Servicing Mission

This section describes the more promising of the two hosted options. All missions identified in

this type provide one or more spacecraft contact events, providing the only opportunity to

measure electrostatic discharges between two spacecraft. Without these contact events, the

full set of desired measurements cannot be made.

Further, the investigational nature of robotic servicing missions may allow the hosted payload

to gain concessions on bus design features to accommodate special needs. Features such as

boom mountings and a spin stabilized bus are far more likely to be accommodated on a

demonstration mission than on a communications asset, for example.

Four robotic servicing missions of opportunity reviewed for this study were the NASA RESTOR  

mission, ATK Space Systems ViviSat, the DLR (German Aerospace Agency) Deutsche Orbitale

Servicing Mission (DEOS), and MDA Corp. (MacDonald, Dettwiler and Associates) Space

Infrastructure Services (SIS).

RESTOR: This is a NASA GSFC mission that expands on the current robotic servicing

efforts at the International Space Station (ISS) (NASA GSFC 2010). RESTOR seeks to demonstrate

the capability to rendezvous with and tether to a spacecraft already in GEO, then provide

servicing and refueling through the use of a dexterous robotic arm. The mission would provide

at least one contact ESD event observation opportunity for each satellite serviced. Disruptions

of the MGS environment measurements due to movement of the RESTOR spacecraft to client

locations are unlikely to significantly degrade the MGS environment data, since such

transportation events are relatively infrequent and occur at or near GEO.  Because RESTOR is a

NASA mission, it likely offers the most flexibility to accommodate requirements for MGS. Under

the umbrella of an internal NASA mission, payload requirements beyond those of the primary

mission (e.g. spin stabilization mode of operations, deployable boom for the magnetometer
MGS Environmental Effects Measurements Study      Page | 20

instrumentation, and location of instrumentation) can be assessed based on the cost and value

returns to the entire platform, rather than just the cost and risk to the primary mission.  

ViviSat: The ViviSat mission is an ATK program that focuses on extending the life of

existing GEO satellites through the addition of a secondary propulsion system (ViviSat 2011).

Under the current ATK design, each ViviSat platform would be configured to dock with the

apogee kick motor of the customer spacecraft, allowing the ViviSat propulsion system to be

used for maneuvers such as station keeping, reboost, and transfer to a disposal orbit.  

Depending on the customer satellite characteristics, Vivisat may extend the mission lifetime of

the client spacecraft by up to 16 years.   As Vivisat is designed to be an attachment to pre‐

existing spacecraft, features that may be desirable for the MGS measurement suite, such as

spin stabilization and deployable booms, are not likely to be acceptable to the customer

spacecraft. Data on the expected data flow between ViviSat and the customer spacecraft was

not available for this study. Depending on the communications load already existing on the

spacecraft, very little may be available to downlink additional data.

DEOS: The DLR DEOS mission is similar to the NASA RESTOR mission in that it intends to

demonstrate on‐orbit rendezvous and servicing capabilities (DEOS 2011). While the first DEOS

demonstration mission is limited to LEO operations, subsequent missions are intended to

extend the mission capability to GEO. This mission shares many of the advantages listed for the

RESTOR mission, but any payload requirement that introduces additional limits on or risk to the

primary mission will be less of a shared responsibility and more of a burden for the MGS team

to resolve. Additionally, communications capability will be prioritized to the primary mission

and NASA will not have direct control over the spacecraft lifetime or the length of the data

record.

MDA SIS: The SIS mission is MDA’s proposal for establishing a capability to refuel

satellites in GEO (MDA 2011). SIS is distinguished from the other missions by being designed for

reusability. Once the spacecraft has expended its initial supply of fuel, MDA can launch a

replacement fuel canister to resupply the robotic spacecraft. This capability greatly enhances

the expected useful lifetime of the robotic servicer and provides the opportunity for a long

duration measurement set. Similar to the DEOS mission, this platform will have limited

flexibility to offer the MGS payload accommodations, but the communications capabilities of

the spacecraft may be more advanced. The current operations plan calls for ground controlled

servicing and should provide a significant telemetry stream to the payload when the primary

mission is not using it to enable servicing.     
MGS Environmental Effects Measurements Study      Page | 21

B) Commercial Communications Satellites

As shown by the desire to build servicing spacecraft, there exists a large commercial market in

GEO. The hundreds of existing assets in GEO are typically replaced at a rate of 15 to 20 new

launches per year. As the expense of maintaining an orbital fleet has increased, commercial

entities have increased efforts to find ways to offset the upfront costs of placing new assets

into orbit. In recent years, many have begun to explore the option of opening unused capability

on their spacecraft in order to host secondary payloads. This sharing of capability is

advantageous to both sides, as the commercial provider receives a fee for the opportunity and

the hosted payload  gains access to space at a fraction of the traditional cost.  

Several exploratory projects have sought to test the waters with commercial partnerships and

have collected lessons learned. Recently, the U.S. Air Force commissioned Intelsat to document

experiences with hosted payloads and to generate a set of guidelines for parties interested in

such a partnership (Intelsat Web 2011). Futron Corporation and NASA have also examined the

challenges and opportunities associated with secondary hosting on commercial assets.

The Hosted Payload Guide Book (Futron 2010) provides an overview of the entire hosting

process from the point of view of the payload provider, as well as pertinent lower level

considerations during each phase. Additionally, the manual identifies the currently planned

commercial satellites during the 2011‐2016 timeframe. Based on this list, there are

approximately 15 promising launch opportunities with seven of those having a significant

probability to launch on a US launch vehicle.

In 2011, Andraschko, et al. documented a NASA study of five recent government payloads that

were hosted on commercial communications satellites. They reported several key technical

parameters that informed our study. First, results from discussions with the satellite bus

manufacturers indicated that almost all options could provide at least 50 kg of mass and more

power than the radiation instrumentation suite will require. Additionally, telemetry data rates

from even the least capable bus allowed 70 Mbps, far in excess of the needs of the MGS

environment suite. The reference also provided a first order cost estimate trend, derived from

the five identified projects, pricing a 50kg allocation at approximately $15M (FY10). The trend is

provided below.   

Cost = 0.1592*(mass in kg) +6.6343

These factors make the commercially hosted payload an attractive option. In addition to the

cost savings, the high launch rate of commercial satellites to GEO makes missed opportunities

recoverable. If circumstances force the program to be delayed, future missions will typically be

available at a rate of at least two to three per year. Lastly, the nature of the platform provides
MGS Environmental Effects Measurements Study      Page | 22

the hosted payload a stable, long duration observation window with an effectively

unconstrained data rate for returning science data to Earth.

While the commercial hosting option provides one of the most compelling cases found in this

study, it is not without its own set of challenges. Commercial time tables are fairly short and do

not easily accommodate schedule slips. A hosted payload must work with the bus provider very

early on (especially if non‐standard options are being requested) to ensure compatibility with

all involved parties. Options such as a spinning bus and deployable boom mounts are not likely

to be considered by the commercial provider, which places additional engineering and ground

testing burdens on the payload development. Fortunately, techniques have already been

demonstrated to mitigate these issues. Many radiation instruments are being designed with

three‐axis controlled spacecraft in mind, making adjustments to current designs more of an

engineering development rather than a technology effort. Additionally, careful treatment of the

solar panel electrical wiring and other bus electronics can limit the magnetic contamination

enough that ground characterization efforts can remove most of the noise generated by the

spacecraft if the magnetometer cannot be boom mounted.

(2) Secondary Payload

Secondary payloads take advantage of excess capability on a launch vehicle that is already

planning to go to the desired location. Sometimes this happens through planned cooperation of

multiple missions, but often it is coordination through the vehicle provider or the sponsoring

agency that brings these opportunities to life. The Ariane V launch vehicle routinely lofts

multiple spacecraft to GEO Transfer Orbit (GTO), but our research did not identify any US

launches that achieved the same (Aziz 2000, Space Launch 2011). The highly configurable

nature of the US fleet and the lack of a developed multiple payload interface on the current

EELVs may contribute to this situation.

Three primary options were identified for this category of launch opportunities: A) standard

secondary payloads; B) extremely small satellites capable of attaching to the standard ESPA ring

adapter used on the current US EELVs, and C) a Lunar Crater Observation and Sensing Satellite

(LCROSS) derived bus that uses an ESPA ring as the primary bus structure. Each of these options  

is discussed below.  

A) Standard Secondary Payload

In order to assess the feasibility of this option, a first order spacecraft sizing effort was

undertaken using the LaRC SmallSat model (Ferebee 1997). For the spacecraft mass estimates,

the radiation suite was assumed to be a static payload of 100 kg requiring 150W of power. It

was assumed that the payload had no significant unique thermal requirements and that a

system capable of operating reliably for more than five years was necessary. Additionally, the

flight system had to be capable of (1) performing the orbital transfer from GTO to GEO and
MGS Environmental Effects Measurements Study      Page | 23

station keeping, and (2) providing 10 mbps capacity in the communications system. Combined,

these requirements translated into a flight system that was on the order of 1000 kg (including

contingency). A second pass through the requirements set reduced the flight hardware mass to

approximately 700 kg. This reduction was achieved through significant relaxation of the station

keeping requirements and specific orbital planning to reduce station keeping propellant usage.

The reduction in station keeping propellant provided substantial additional mass savings in the

propellant necessary to perform the GTO to GEO transfer burns.

After completing initial sizing of the spacecraft bus, a first order cost estimate was developed

using the Small Satellite Cost Model (SSCM) (Aerospace Web 2010). This model is maintained by

The Aerospace Corporation and uses parametric cost estimating relationships (CERs) to produce

a first order cost of the spacecraft. The spacecraft bus is estimated to cost on the order of

$50M (FY11).

Since a significant mass driver in the preliminary sizing effort was the propellant required, a

subsequent effort examined electric propulsion options. However, this option could not be

addressed effectively at the same level of fidelity and required some gross assumptions to be

made. In the end, all of the mass saved from the conversion (and then some) went back into

the system to provide significantly increased power and to address growth in the other flight

system components. Based on the resulting mass, the electric propulsion system did not

provide a benefit sufficient to offset the additional system complexity.

To complement the first order sizing of an electric propulsion bus, a literature survey was

conducted to see if any similar systems existed or were in development. The research found the

CX2 spacecraft bus, a small (approximately 1200 kg) GEO‐class bus that uses electric propulsion

for both the GTO to GEO transfer and GEO operations (Dutch Space 2009). The CX2 is designed

to fit inside the Ariane V and utilizes some of the typically unused volume and mass allocations

of that launch vehicle. It has a nominal payload mass of 100 kg and is designed as a

communications platform. This allowed a straight forward comparison to the previous sizing

attempt. While the CX2 is slightly heavier and more capable, the comparison between the

chemical propulsion and the CX2 remains the same as the in‐house sizing attempt. However,

the CX2 platform does have several advantages in that it is already mostly designed, has a

planned demonstration near the end of 2012, and is specifically configured to work within the

Ariane V launch constraints. This combination offers significant potential cost savings, and the

expected cost of the CX2 system, including launch and insurance costs, is expected to be on the

order of $100M.

Use of the Ariane V or other foreign launch vehicles is limited by NASA to circumstances in

which the launch is 100% contributed (no U.S. government investment).    Such a limitation is

likely to dramatically reduce the opportunities for pursuing this option.
MGS Environmental Effects Measurements Study      Page | 24

Without a multiple payload adapter for GTO payload delivery, the use of the U.S. EELV fleet is

also unlikely to yield standard secondary payload opportunities in the near future.  If such an

adapter became available (for example, the Atlas V has a design that has completed PDR at this

time), it would provide some additional options that are described here. The mass of any

additional hardware necessary to support the launch of multiple full mission‐class payloads was

not included, and evaluations relative to both the base loft capability of the 5m shroud option

and the incremental lift capability were assessed. For the Atlas V, the step size in loft capability

was found to be on the order of 1000 kg per increment. Based on our minimum 700 kg mass

requirement, it is not likely that our additional mass can be accommodated within the

configuration selected for the primary mission with enough margin to make the primary

mission comfortable. The addition of a single increment in capability could accommodate the

secondary payload, but the secondary payload would most likely have to pay both a share of

the base launch vehicle costs as well as the entire cost of adding the additional capability. The

Delta IV has some larger increments and the heavy variety would likely have enough excess

capability to absorb the mass of the small secondary payload.  

B) The EELV Secondary Payload Adapter (ESPA) Ring

The ESPA ring is a secondary payload adapter designed to accommodate several small payloads

in the excess capacity of the EELV fleet.  It mounts below the primary spacecraft and acts as a

structural support between the EELV cone adapter and the primary payload fairing. An image is

shown below in Figure 7.

Figure 7.  ESPA configuration diagram from the ESPA User’s Guide.
MGS Environmental Effects Measurements Study      Page | 25

The ring is capable of supporting up to six small secondary payloads and can be equipped with a

“soft ride” system to limit the potential shock and vibration environment the payloads endure.

Additional development efforts are under way to provide smaller and larger versions of the

ESPA. The smaller version is designed to provide the ESPA capability to the smaller class launch

vehicles (e.g. Delta II class) while the large version is designed to increase the payload mass

limit on each port for the EELVs. A novel use of the entire ring as a bus structure for a single

secondary satellite was successfully implemented on LCROSS (launched in June, 2009) and is

planned for DSX (scheduled for launch in October, 2012).  Options for using the ESPA ring both

as a small satellite deployment mechanism and as a primary structure were examined for this

study. Note that the ESPA Payload Planners Guide recommends submitting a spacecraft

questionnaire to the Space Test Program (STP) three years prior to the desired launch

opportunity (DoD STP 2001).

C) ESPA Small Satellite Payload

The ESPA payload satellites are small spacecraft that fall between “cubesats” and traditional

spacecraft in terms of size. Satellites of this class are designed to anchor to the ESPA ring

utilizing pre‐determined bolt holes and specified mass and volume constraints. A picture of the

size and attachment location is shown below in Figure 8. Full specifications on the satellite

constraints can be found in the ESPA User’s Guide (DoD STP 2001).   

Figure 8. ESPA ring (payload volumes in inches:

35.50” x 28.00” x 24.00”).

Figure 9.  A spacecraft built using the ESPA

Standard Interface Vehicle (SIV).

As shown by the dimensions in Figure 8, this option is extremely volume limited. Additionally,

each payload has a not to exceed (NTE) mass of 180 kg, which can be further limited by the

launch authority if needed to maintain center of gravity (cg) balancing. Using the STP’s Standard

Interface Vehicle (SIV) as a reference point, only 60 kg of the mass is available for payload. An

image of the SIV is shown above in Figure 9 to illustrate the relative size.

While this option is limited in what instrumentation it can accommodate, it does offer a

significant advantage: the STP office will supply the access to space at no cost to the payload.
MGS Environmental Effects Measurements Study      Page | 26

To qualify, the requesting party must submit a proposal to the STP review board and be

accepted into their candidate pool (Marlow 2011). The STP expects to provide at least one

opportunity per year for these flights, with a fraction of those going to GEO.  

D) Lunar Crater Observation and Sensing Satellite (LCROSS) Derived ESPA Bus

This option is an enhancement  to the standard ESPA class mentioned above. This vehicle uses

the full ring as the basic bus structure and builds control modules that can be attached to the

ring. This configuration allows the craft to use the full mass and volume allocation available to

the ESPA secondary payload.

The first mission to show the viability of this concept was LCROSS (shown in Figure 10). LCROSS

was a competitively selected mission borne out of the excess capability available once LRO was

moved from a Delta II to an Atlas V (Lo 2008). The DSX vehicle design used a similar approach,

proving the concept had general applicability. Based on these successes, versions of this

spacecraft bus have been made available on the NASA Rapid Spacecraft Development Office

(RSDO) catalog offerings. A notional version of a powered ESPA ring is shown in Figure 11.

Figure 10. LCROSS

Figure 11. Notional powered ESPA ring

(image by Vector Design)

The biggest challenge to this design for the purposes of GEO spacecraft is the propulsive needs

of the mission and the volume available for propellant tanks. None of the current RSDO

offerings for this bus option can support a mission to GEO. However, Northrop Grumman has

been working on a design capable of doing the orbital transfer and relevant station keeping

while maintaining enough resiliency to ensure several years on orbit (Drucker 2011). Based on

the LCROSS bus development costs, we estimate the LCROSS‐derived ESPA bus cost to be in the

$50‐60M dollar range.
MGS Environmental Effects Measurements Study      Page | 27

3) Primary Payload

This option is the path that most missions follow. In this option, a custom, dedicated bus is built

and the mission procures its own launch services. Due to the prevalence of the commercial

market, there are many standard options for both spacecraft buses and launch services to GEO.

Based on the general availability and the unique customization this option allows, no significant

effort was spent to identify specific available options. However, typical mass and power

capabilities were examined to provide the springboard for the identification of pathways

forward.

Based on the sizing done for the “Secondary Payload” option, all standard GEO platforms

provide mass, power, and volume capacity well beyond anything our suite could effectively

utilize. As the least capable launch vehicles have loft capability well above our secondary

payload mass, there exists very little reason to not utilize a platform similar in size to the

average communications spacecraft.

The use of an average communications spacecraft size bus makes a significant amount of mass,

volume, and power available for other uses. Under this scenario, the mission could host other

payloads and experiments. NASA could use this platform to develop a competitive selection of

secondary experiments and greatly increase the science returns of the mission.

Should NASA decide not to fund extra payloads, it is possible that the excess capability could be

utilized to forgo the deployment of the spacecraft at GTO and instead have the launch vehicle

(LV) do the insertion burn. This is not commonly done by US vehicles but does sometimes

happen for foreign launches. The data and analysis required to determine the viability of this

option is beyond the scope of this task, though the authors believe that this option is unlikely to

represent substantial benefits over the other options presented.       

The following FOMS for the GEO transportation options were developed to enable a coarse,

quantitative trade and are summarized in Table 8.

Table 8.  GEO Transportation Options Figures of Merit

FOM Importance Description/Comments

Cost High

Addresses the cost of  delivery of the payload to orbit and relevant

operations costs. Its weighting is derived from the idea that the

higher the data per dollar ratio the mission achieves, the more likely

it is to be developed.   

Launch Availability High

Addresses the expected rate of flight of each option. The weighting

is based on the desire to avoid “one shot” opportunities which may

or may not be compatible with the necessary development

schedule.  
MGS Environmental Effects Measurements Study      Page | 28

External Schedule

Dependency Medium

Addresses the influence of the platform option on the payload’s

development schedule. The weighting is based on the significant

impact such dependencies can have on the payload’s development

plans.

Operational Life Low

Addresses the expected length of the data record. The weighting is

derived from the short time available to influence new EVA suit

design.  

Mounting & Pointing Low

Addresses the ability of the platform to meet the measurement field

of view needs. The weighting is based on the ability of instrument

design to mitigate issues of this type.  

Telemetry Bandwidth Medium

Addresses the downlink capabilities of the platform.  

Power Low

Addresses the platform’s capability to meet the base power

requirements of the payload. The weighting is based on the

potential impacts of power shortfalls.   

Docking Event for

ESD High

Assessment of whether or not the option will provide contact

events between two spacecraft. The weighting is due to the inability

to complete the full measurement set without this event.  

Scores for each option based on weighted FOMs are shown below in Table 9.  The hosted

robotic servicing mission, with a FOM score of 395, was determined to be the best opportunity

for GEO transportation.

Table 9.  Bus Options Evaluation

The ranking of all options considered based on score is shown below. Note that in the ranking

we further distinguish a government robotic servicing mission from a commercial opportunity,

since the former may represent a greater opportunity to  affect the bus design.
MGS Environmental Effects Measurements Study      Page | 29

                Ranked Bus Options

1. Hosted, Government Robotic Servicing Mission

2. Hosted, Commercial Robotic Servicing Mission

3. Hosted, Other GEO Sat

4. Secondary ESPA Payload

5. Secondary ESPASat

6. Secondary, Unspecified Launch Vehicle

7. Primary Payload

Bus and Instrument Trade Space

The final part of the MGS measurements trade study examined paths through the complete

trade space by evaluating the instrument suites using weighted “Instrument Suite” FOMs,

including a metric called “Bus Flexibility.” This FOM was used to track any drawbacks that a

given MEES configuration would incur on a given bus.    Only the volume‐constrained

configuration, for example, is likely to fit into the available volume on an ESPA smallsat payload.

A total of five FOMs were used to rank the instrument suite configurations as shown in Table

10. Table 11 shows a coarse numerical evaluation used to quantify the Bus Flexibility metric,

and Table 12 shows the numerical evaluation of the four MEES configurations.

Table 10.  Instrument Suite Figures of Merit

FOM Importance Description/Comments

High Data Value High

Suites that include the highest capability instruments with good

coverage of the parameter range and high resolution score well in

this FOM.   

Low Cost High

Includes the estimate for modifying the instrument for GEO

applications, if required.  

Bus Flexibility Medium

The ESPA payload transportation option will not accommodate all of

the MEES options. A separate quantitative compatibility evaluation

was done to create a compatibility score.

Short Development

Time Low

A general assessment of the availability of the instruments in the

configuration within a short time frame.  

Low Development &

Operational Risk Low

A general assessment of the overall readiness of the configuration

for application to MGS.  
MGS Environmental Effects Measurements Study      Page | 30

Table 11.  MEES Compatibility with Bus Options

Note: Cell colorization is for enhanced readability only.

The compatibility comparison shown in Table 11 indicates that the notional MEES experiment

can be well accommodated on most bus options (shown in the green areas).  Space constraints

may be a factor for the ESPAsat (using the ESPA ring as a primary structure) and will be almost

certainly a limiting factor for an ESPA payload option (using the ESPA Standard Interface

Vehicle).  

Table 12.  MEES Configuration Evaluations

Table 12 compares the MEES configurations using weighted FOMs. The results equally favor the

maximum capability approach (“Max Capability” = 230) or a volume optimized approach

(“Small” = 230).  Clearly, the high weighting of “High Data Value” is significantly affecting the

outcome, and revised weightings and finer resolution in the FOM ratings will produce different

results.    One interpretation of these preliminary results is that investment in an MGS

instrument suite should be “all‐or‐nothing” in the sense that the cost and development savings

for a less than ideal system also greatly undermines the utility.  Having a viable “small volume”

option, however, is also valuable due to the ability to capitalize on a greater variety of
MGS Environmental Effects Measurements Study      Page | 31

transportation opportunities, especially since GEO is not a common destination for secondary

payloads.  Because of the FOM weightings used in this table, a key contributor to the high score

of the small volume configuration was the Cosine MPS energetic particle spectrometer.   This

instrument advertises good resolution over a wide energy range and is flight qualified, but does

not have flight heritage (Maddox 2008).    Further evaluation of the capabilities of the Cosine

MPS by independent instrumentation experts is suggested as future work for the MGS team.

Figure 12 depicts the entire trade tree of the options discussed in this study.  In general we see

that while all bus options are viable, the Hosted options are recommended.  Becoming a hosted

payload on a robotic servicing mission is of particular value due to the opportunities to measure

contact ESD events; a U.S. government mission is recommended due to potentially increased

ability to influence the system design requirements. The “Other GEO Sat” branch includes

commercial communications satellites, where frequent launch opportunities are found.  MEES

Option 1 is recommended due to high science return, and Option 4 is recommended for cost or

size constraints that may be encountered once the actual vehicle is identified and the program

budget determined.

The Secondary branch is rated as a feasible fallback option rather than a recommended option

due largely to the need to invest in a spacecraft bus.    Further, the capability to deliver GEO

payloads using the ESPA has been proposed but not demonstrated, and is unlikely to become as

common as primary launches to GEO.    The secondary payload approach is also identified as

viable but not recommended due to the lack of launch opportunities, since this approach is not

in use for U.S. vehicles and foreign launches may not be purchased by NASA.  Use of a Primary

launch is likely to be viable only if a collection of other experiments can be assembled to justify

the cost.
MGS Environmental Effects Measurements Study      Page | 32

Figure 12. MGS environment study trade tree.
MGS Environmental Effects Measurements Study      Page | 33

Conclusion

The desire to conduct a manned mission in Geosynchronous Earth orbit requires a thorough

understanding of the space environment in order to design vehicles and suits that will provide

sufficient protection for astronauts and equipment.    Predictive models used to develop

requirements for such designs have not historically been exercised for manned GEO missions,

and do not at present provide sufficient confidence for MGS needs.  While efforts to refine the

models are on‐going, the challenges of predicting space weather prescribe the need for

sustained, long duration flight data in the regime of interest.

The data needed for the success of MGS can be gathered by developing an instrument suite to

collect both environment and effects measurements. These data can then be used to refine the

predictive models and inform the design of mitigation strategies for radiation harms and

damaging discharge currents.    The instruments comprising this suite are of sufficient

technology readiness to enable near term development, and have relatively low mass, volume,

power and telemetry resource needs.  A compact system capable of being operated as a hosted

payload on a GEO asset is recommended. This study presented representative configurations of

known instruments of the type that could collect the desired data and evaluated these notional

instrument suites using weighted FOMs informed by subject matter experts in radiation and

spacecraft effects.    Suggestions for further study include revisiting the data needs after the

completion of the ongoing LANL spectrometer data analysis task and conducting a survey of

instrument providers in the form of a Request for Information to collect more detailed

instrument design and cost data.  

   
MGS Environmental Effects Measurements Study      Page | 34

Acknowledgements

Invaluable guidance on space weather, radiation, and environmental effects on spacecraft, as

well as key input for developing the instrument trades, was provided by Dr. Steve Blattnig  of

the NASA Langley Research Center and Dr. Joseph Minow of NASA’s Marshall Space Flight

Center.  Leadership provided by Dana Weigel (NASA Johnson Space Center) enabled the authors

to focus the study on the key needs of the MGS team.    The following individuals graciously

spent time in responding to queries from the authors:

Mark Andraschko    NASA Langley Research Center

Janet Barth     NASA Goddard Space Flight Center

J. Bernard Blake    Aerospace Corporation

Thomas Conroy     University of Regina

David Cornelius     NASA Langley Research Center

William Crain     Aerospace Corporation

Eamonn Daly     European Space Agency

Eric Drucker     Northrop Grumman Corporation

Theodore  Fritz     Boston University

Gary Galica     Assurance Technology Corporation

Henry Garrett        Jet Propulsion Laboratory

Peter Harvey     University of California, Berkeley

Gary Hendel       Aerospace Corporation

Insoo Jun     Jet Propulsion Laboratory

Bernard Kelm     Naval Research Laboratory

Craig Ketzing     University of Iowa

Wousik Kim     Jet Propulsion Laboratory

Erik Maddox     Cosine Inc.

Joseph Mazur     Aerospace Corporation

Gary Mullen     Assurance Technology Corporation

Mark de Payrebrune    DPL Science Inc.

Benjamin Reed     NASA Goddard Space Flight Center

Geoffrey Reeves    Los Alamos National Laboratory

Edward Semones    NASA Johnson Space Center

Harlan Spence     University of New Hampshire

Brook Sullivan     Booz Allen Hamilton

JacquelineTownsend    NASA Goddard Space Flight Center

Julie Williams‐Byrd    NASA Langley Research Center

Michael Xapsos     NASA Goddard Space Flight Center

  

   
MGS Environmental Effects Measurements Study      Page | 35

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APPENDIX A:   Instrument Comparisons
MGS Environmental Effects Measurements Study      Page | 39
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MGS Environmental Effects Measurements Study      Page | 41

APPENDIX B:   Instrument Data

Note:  Slides contain partial information and, in some cases, estimated information and are provided for

reference only.  Please refer to the instrument PI for complete data.
MGS Environmental Effects Measurements Study      Page | 42
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MGS Environmental Effects Measurements Study      Page | 61

APPENDIX C:   Acronyms

APEX       Advanced Photovoltaic and Electronics eXperiments

AEEF       Alphasat Environment and Effects Suite

ARC       Ames Research Center

CDS       Charge Deposit Sensor

CER         Cost Estimating Relationship

CG       Center of Gravity

CRRES       Combined Release and Radiation Effects Satellite

DLR       Deutsches Zentrum fur Luft‐und Raumfahrt  

     (German Aerospace Agency)

DEOS       Deutsche Orbitale Servicing Mission

DPAF       Dual Payload Attachment Fitting

DSX       Demonstration and Science Experiments Mission

EDM       Electronic Discharge Monitor

EELV       Evolved Expendable Launch Vehicle

ESA       European Space Agency

ESD       Electrostatic Discharge

ESPA       EELV Secondary Payload Adapter

eV       Electron Volt

EVA       Extravehicular Activity

FOM        Figure of Merit

FOV        Field of View

FY       Fiscal Year

GCR       Galactic Cosmic Radiation

GEO       Geosynchronous Earth Orbit

GSFC        Goddard Space Flight Center

GTO       GEO Transfer Orbit
MGS Environmental Effects Measurements Study      Page | 62

ISS       International Space Station

LANL       Los Alamos National Laboratory

LaRC        Langley Research Center

LCROSS    Lunar Crater Observation and Sensing Satellite

LEO       Low Earth Orbit

LET       Linear Energy Transfer

LRO       Lunar Reconnaissance Orbiter

LV         Launch Vehicle

Mbps       Megabits per second

MEES         MGS Environment and Effects Suite

MEO       Medium Earth Orbit

MGS       Manned GEO Satellite Servicing

MPS       Magnetospheric Particle Sensor

MPS       Multifunction Particle Spectrometer

NOAA       National Atmospheric and Oceanic Administration

NTE       Not to Exceed

PDR       Preliminary Design Review

PI       Principle Investigator

RBSP       Radiation Belt Storm Probe Mission

RESTOR     Reusable Earth Synchronous Tele‐Operated Refueler

RSDO       Rapid Spacecraft Development Office

SCATHA     Spacecraft Charging at High Altitude

SIS       Space Infrastructure Systems

SIV       Standard Interface Vehicle

SPE       Solar Particle Event

SSCM       Small Satellite Cost Model

SPENVIS     Space Environment Information System

STP       Space Test Program Office
MGS Environmental Effects Measurements Study      Page | 63

TEPC       Tissue Equivalent Proportional Counter

TID       Total Ionizing Dose

TRL       Technology Readiness Level

USAF       United States Air Force
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Manned GEO Satellite Servicing Mission Environmental Effects 

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A trade study was conducted to evaluate options for collecting space environment data in geosynchronous earth orbit to 

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Dosimeter; EVA; GEO; Magnetometer; Radiation; Satellite servicing; Space environment; Spacecraft charging; Spectrometer

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