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					   Space Radiation Analysis using free Internet Tool, a
          Case Study using 3-Corner Satellite
                                    J. Hackel
                               SS#: 578-98-8529
                      Department of Aerospace Engineering
                        University of Colorado - Boulder
                            Boulder, Colorado 80303




ABSTRACT: This report details how to use the free internet tools available to
predict the radiation environment for a satellite, specifically for the 3 Corner
Satellite (3CS). The radiation environment is a particular concern for this
mission, since the electronics used are COTS and subsequently non-radiation
hardened. These electronics have a damage threshold of 1000 Rads. To
determine the radiation environment, numerical computer models available from
SPENVIS (SPace ENvironment Information System) were used in the analysis
effort. This computer model is freely available on the web and its results will be
compared with CREME96 model, which is a standard model in industry.
However, the CRÈME96 model is not free for use. A free Internet application
would be a boon to private organizations and universities planning to launch a
satellite in orbit. In this report six possible cases that are explored, stemming
from three possible orbits for this mission and two possible mission lifetimes.




Joe Hackel                             Page 1                               8/8/2011
        ABSTRACT:. ......................................................................................................................................... 1
INTRODUCTION ........................................................................................................................................ 3
    SPACECRAFT OVERVIEW ............................................................................................................................ 3
      Mission Design ..................................................................................................................................... 3
      Mission Goals ....................................................................................................................................... 4
    PREDICTED ORBIT ...................................................................................................................................... 5
RADIATION ENVIRONMENT ................................................................................................................. 5
    INTRODUCTION ........................................................................................................................................... 5
    SOURCES OF PARTICLE RADIATION ............................................................................................................ 6
    EFFECTS OF RADIATION.............................................................................................................................. 6
    SOLAR MAXIMUM ...................................................................................................................................... 7
ANALYSIS .................................................................................................................................................... 7
    GATHERING OF RESULTS ............................................................................................................................ 7
    THE SETUP OF THE SPENVIS WEB PAGE ................................................................................................... 9
    RESULTS FOR 120 DAYS.............................................................................................................................11
    RESULTS FOR 90 DAYS...............................................................................................................................14
    RESULTS OF SPENVIS ..............................................................................................................................15
CONCLUSION ............................................................................................................................................16




Joe Hackel                                                                Page 2                                                                  8/8/2011
Introduction

Spacecraft Overview
This project is a joint effort among Arizona State University (ASU), University of
Colorado at Boulder (CU), and New Mexico State University (NMSU). Aptly named
Three Corner Sat (3CS), our proposed constellation of three identical nanosatellites will
demonstrate stereo imaging, formation flying, and innovative command and data
handling. The Colorado Space Grant Collage has a solid team’s heritage in: space flight1,
conventional satellite design2, and nanosatellite design3. This constellation will be ready
for launch in August 2002.


Mission Design
The 3CS constellation will consist of three satellites (for a picture of one satellite, see
Figure 1 below) flying in a linear follow-formation with slowly increasing separation
distance from each other. The maximum separation distance has been selected is based
on altitude




                                Figure 1: Picture of 3^Sat

and camera field of view (FOV) at 200 km. The satellite will use gravity-gradient (GG)
forces for stabilization with +/- 5 degrees pointing accuracy. The three satellites will be
stacked together (see Figure 2 below) on a MSS plate, and launched from the space
shuttle as an experiment. The space shuttle’s orbit and inclination is not known with

1
  CU’s DATA-CHASER payload via August ’97 Space Shuttle
2
  CU’s Citizen Explorer via Delta, launch pending
3
  ASU’s 4.5kg ASUSat1 via OSP Minotaur, September ’99


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certainty now, but there are 3 likely orbital cases, which will all be analyzed. The stack
will be launched from the plate, and then the stack will be broken up by a Lightband
system (which is being developed currently at AFRL) and each satellite will look like
Figure 1 above. This Lightband system can be seen on Figure 2 as the red/green sections
between satellites. The Lightband system currently is set up to give a mission lifetime of
90 days, although the mission goals state a mission life to 120 days.




                        Figure 2: Stack Configuration of 3^Sat


Mission Goals
Stereo Imaging: The primary science objective of the 3CS constellation is to stereo image
small (< 100 meter), highly dynamic (< 1 minute) scenes including deep convective
towers, atmospheric waves, and sand/dust storms. These stereo images will enable the
computation of range to within 100 meters giving accurate data regarding the shape,
thickness and height of the observed phenomena. For stereo imaging, a nominal spacing
of tens of kilometers between the satellites is required. With a controlled deployment to
achieve this initial spacing, the satellites will remain within range for the suggested four-
month lifetime of the mission. Therefore, propulsive capability is not needed for these
satellites.

Formation Flying / Communications: To accomplish the science objectives, a "virtual
formation" is proposed and will be demonstrated as part of our program. The virtual
formation is a cooperative effort between satellites operating as a network where
targeting and data acquisition are accomplished and results transmitted to the ground
segment and to the other satellites via communications links without the need for strict


Joe Hackel                                 Page 4                                    8/8/2011
physical proximity of the satellites. In this mode, the communications links carry the
command and control data necessary to accomplish the mission regardless of the physical
location of the satellites. For the mission to be accomplished, the locations of the
satellites will need to be "in range" and mutually known in order for each to support its
portion of the mission, but physical proximity is not a requirement for the formation
network.


Predicted Orbit
3CS is manifested as a launch experiment on the space shuttle for launch on August 15th,
2002. There are three possible orbits for the payload, as listed in the table below.

                              Altitude        Inclination          Eccentricity
              Orbit 1         350 km             28.5'                0.01
              Orbit 2         400 km             36.0'                0.01
              Orbit 3         400 km             51.7'                0.01
                             Table 1: Orbits for the 3CS Mission

   Orbit 1is the typical space shuttle orbit. It is also the safest orbit for the electronics
    since the low altitude and inclination shield out much of the radiation.
   Orbit 2 is a possible orbit suggested by NASA for the release of instruments.
   Orbit 3 listed is the most likely since it is the orbit of the International Space Station
    (ISS). The space shuttle will have a majority of its missions to the ISS over the next
    few years. The three possible orbits represent a low, middle and high risk to the
    electronics of 3CS.



Radiation Environment

Introduction
The radiation environment of a satellite consists of energetic particles that can impact the
spacecraft and degrade materials, effect computer systems, and even charge the
spacecraft. It is therefore necessary to analyze this environment to determine the number,
type and energies of the particles in 3CS’s possible orbits and how these values will
change over the duration of the mission.

This mission is using non radiation hardened electronics. While this procedure is not the
norm (even with the event of “smaller, better, cheaper, faster”), it is relevant. There are
fewer manufactures willing to preclude radiation-hardened electronics. In the last 5 or so
years, Texas Instruments, Intel, TRW, LSI Logic and AT&T have all exited the radiation
hardened microelectronics market4. This means that spacecraft system designers have to

4
 Taken from “Spacecraft Environments Interactions: Space Radiation and Its Effect on Electronic
Systems” , Howard, Computer Science Corporation


Joe Hackel                                      Page 5                                            8/8/2011
become more aware of the radiation environment and its effect on electronics. In
addition, non radiation-hardened electronics may play a larger role in the future.


Sources of Particle Radiation
Overall, the near-Earth radiation environment can be divided into a trapped radiation
environment and a transient radiation environment. Particle radiation originates from sun
and is transferred to Earth orbit by solar wind or other particle flows in the radial
direction outward from the sun. Particles can then become trapped on Earth's magnetic
field lines and either remain on these field lines or become lost by entering the Earth's
atmosphere.

Transient radiation environments, consisting of events like solar wind, solar flares and
galactic cosmic radiation (GCR). Solar wind typically has low energy particles (electrons
and protons) and is not of much concern for internal components, like electronic boards.
Solar flares are large influence on electronics from the energetic protons and heavy ions.
Some of the particles (those lower in energy) are deflected along magnetic field lines, a
screening effect which is important to satellites. GCR is radiation from beyond the solar
systems. Like particles from the solar flares, these particles can be deflected by the
Earth’s magnetic field. Hence, inclination and altitude is important design
considerations.


Effects of Radiation
To understand the effects radiation in space on electronic systems, the sources of particles
were discussed in the last section. The particles impact on the electronic materials in two
ways: electronic interactions and nuclear interactions. In electronic interactions, the
impacting atom deposits a significant amount on energy. In nuclear interactions, the
particles actually disrupt the systems. Electronic interacts are more often. There are 3
effects categories of these interactions: Total ionizing does (TID), displacement damage,
and single event effects.

The term TID implies the dose is deposited to the electronics through ionizing effects
only. The energy deposited from the radiation moves the impacted particle to a higher
energy state. This lead to poorer electron diffusion through the circuit. The increase in
energy can lead to a degradation of the circuit. The effects of TID can be mitigated
through shielding.

Displacement effects can also occur. The extra energy can add a current path that
previously did not exist. This can cause circuit leakage, or simply make a path in the
electronics that did not exist before. Naturally, the effects of displacement can become
cumulative over time.
The final event that can occur because of radiation is the single event effect (SEE).
Overall, SEE is a generic term encompassing a wide range of effects, most notably single
event upsets (SEU) and single event latchtup (SEL). A SEU is when a digital circuit
change logic states (usually the flipping of a bit from 1 to 0). A SEL occurs when the
device freezes into a state where reboot is needed. Unfortunately, SEEs are difficult to


Joe Hackel                                 Page 6                                    8/8/2011
defend against. In this case, the designer simply needs to figure out the frequency of such
an effect and guard against it. In general, the larger the cross section of the piece of
equipment, the higher the chance of an event, according to the LET number calculated.
LET is a paramter that indicates how a particle loses energy as it passes through material.
Most of the times, the number of events per day are calculated and some type of error
protection is implemented.

There are many effects from energetic particles in orbit and, therefore, the following
section contains several different plots that give us a good idea of the numbers of
particles 3CS can expect over time. Also important are the energies of the most prevalent
particles, as this will give us a good idea of what to expect in terms of material
degradation and penetration depth into the shielding on-board the spacecraft.


Solar Maximum
Solar Maximum is important concerning our spacecraft since we will be launching during
the maximum in the solar cycle. During solar maximum, particle fluxes are greatly
increased due to highly increased solar activity. This factor was taken into account when
running the particle flux models and these results are contained in the analysis portion of
this report



Analysis
This section details the expected radiation environments, as total ionizing dose and
proton and LET spectra. The primary concern is the radiation particle levels, and
protecting the radiation from them. There will be little degradation of solar arrays over
this short 120-day mission life. For the analysis, 1000 Rads was chosen as the threshold
limit of the electronics5. However, due to the need for a 2x factor of safety, 500 Rads
was taken as the limit. All three orbits were analyzed for both periods, 90 and 120 days.


Gathering of Results
The results were gathered using the web page http://www.spenvis.oma.be/spenvis/. The
SPENVIS project is funded by the European Space Agency through the General Support
Technologies Programme (GSTP)6. It was used instead of the CREME96 model. The
results were then briefly checked by Kevin Miller at Ball Aerospace, who used the Ball
Aerospace models. This web page may be an alternative to the CRÈME model since
with SPENVIS, one can generate a spacecraft trajectory or a coordinate grid and then
calculate:

5
  ASEN 5519 Class notes, Radiation section, p. 5. Also, from conversation with Kevin Miller at Ball
Aerospace.
6
  ...and was initiated by the ESTEC Space Environment and Effects Analysis Section (TOS-EMA). The
project is developed by the Belgian Institute for Space Aeronomy (BIRA-IASB), with sub-contractors
Space Applications Services (SAS) and the Paul Scherrer Institute (PSI). SPENVIS is registered as ESA
contract n°11711-WO1.


Joe Hackel                                      Page 7                                         8/8/2011
 geomagnetic coordinates
 trapped proton and electron fluxes and solar proton fluences
 radiation doses (ionizing and non-ionizing)
 damage equivalent fluences for Si and GaAs solar panels
 LET spectra and single event upset rates
 trapped proton flux anisotropy
 atmospheric and ionospheric densities and temperatures
 atomic oxygen erosion depths
The below screenshot shows what SPENVIS does in its radiation analysis package.




                          ScreenShot 1: Radiation Package

Magnetic field line tracing is implemented, as well as the generation of world maps and
altitude dependence plots of the magnetic field and the current models of the neutral
atmosphere and the ionosphere. Models for spacecraft charging, both surface charging
and internal charging, are available. A tool to visualize satellite data can be used.


Joe Hackel                               Page 8                                   8/8/2011
Micrometeoroid and space debris models are implemented, and an impact risk analysis
module is currently under development. A tool for sectoring analysis in simple geometry
has been added. While only the Radiation Analysis was explored due to lack of time,
there is hope that the author will get a chance to re-explore the other options, time
permitting.
The below picture shows all the possibilities of SPENVIS.




                          ScreenShot 2: SPENVIS Possibilites

The Setup of the SPENVIS Web Page
Here, I will go through the web page briefly, mentioning my choices. It is hoped that a
future reader can use this information and not have to reinvent the wheel. It is noted that
each option mentioned has other options that can be selected, however, in the interests of
brevity, only the options picked for this case will be mentioned

First, an orbit has to be generated using the orbit generator. Here, you can select the type
of orbit (general), Orbit Epoch (15, Aug, 2002, 09h, 16m, 15s). Orbit: Apogee Perigee;
Eccentricity; Inclination Duration (varied, from three different orbits); Satellite
Orientation (one satellite axis parallel to the velocity vector).




Joe Hackel                                 Page 9                                   8/8/2011
                           ScreenShot 3: Orbit Generator

Next, the trapped proton and electron fluxes are evaluated: Trapped Proton Model (AP-
8) Trapped Electron Model (AE-8). Proton Parameters (AP-8 Max omnidirectional),
Energy Thresholds for Output: (10 MeV to 50 MeV). Electron Parameters (AE-8 Max);
Energy Thresholds for Output: .1MeV to 1 MeV.

After that, the Solar Proton models (JPL-96), with geomagnetic shielding present and
magnetosphere conditions (stormy).

Finally, Shielding and mission parameters: Mission duration (90 and 120 days picked);
Shield depth values (mils). (see Screen shot below)




Joe Hackel                              Page 10                                  8/8/2011
                          ScreenShot 4: Sheilding Parameters

Results for 120 days
From this web page, results were generated for the three orbit chosen. Below is the case
for 120 days, the official lifetime of the mission. On all tables, the shielding necessary to
protect the spacecraft from 500 Rads was determined. Although electronics can
withstand about 1 kRad, we will employ a factor of safety of 2x, just as the NASA
standard do. The 3 orbits are mentioned on Table 1 above, but have been repeated here
for convenience.


                           Altitude      Inclination        Eccentricity
             Orbit 1       350 km           28.5'              0.01
             Orbit 2       400 km           36.0'              0.01
             Orbit 3       400 km           51.7'              0.01
                          Table 2: Orbits for the 3CS Mission




Joe Hackel                                 Page 11                                    8/8/2011
                               Figure 3: Orbit 1, 120 days

As it can be seen, orbit 1 is a safe orbit, with the radiation dose never exceeding 100
Rads. If it turns out that the experiment will be in orbit, there is little to concern for the
mission planners. Notice how the SPENVIS website reports the total ionizing dose due
to: trapped protons, solar protons, bremmstrahlung radiation, trapped electrons, and totals
it.




Joe Hackel                                 Page 12                                    8/8/2011
                              Figure 4: Orbit 2, 120 days

As it can be seen here, the amount of shielding needed to protect against 500 Rads of
radiation is about 10 mils.




Joe Hackel                               Page 13                                  8/8/2011
                               Figure 5: Orbit 3, 120 days
Figure 5 shows the worst case possible for this mission: the worst case orbit and the
longest interval of time. This higher rad dosage comes from the higher latitude orbit
where the solar protons at the higher latitude where the solar protons access the
atmosphere along field lines. The shielding level should be around 20 mils.


Results for 90 days
In addition, the case was run for 90 days. The untested Lightband system, as currently
designed, will push the satellites apart “too hard”, giving it a mission life time of about 90
days instead of the planned 120. Planning for this case, the following graphs show the
radiation seen behind mils of shielding.




        Figure 6: Orbit 1, 90 days                            Figure 7: Orbit 2, 90 days
Again, Orbit 1 is safe from radiation, regardless of the time scale. Also, we see the
advantage of the 30 less days: only 6 mils of shielding are needed. Below is Orbit 3.




Joe Hackel                                 Page 14                                    8/8/2011
                                     Figure 8: Orbit 3, 90 days
The less amount of time does not produce much of a saving here: the satellites need 16
mils. This is the result of the exponential decay-like curve of the radiation shielding
graphs.

Also calculated on the web page was the peak flux and solar proton fluence over the 120-
day mission lifetime. This was done as solar proton flux over the life of the satellite is
very important in determining the more serious effects a spacecraft will encounter, the
following plot is of the solar proton fluence integrated over the mission. Notice it
calculates it for both integral and differential flux versus energy. In addition, the LET
spectra were calculated over an orbit. The LET spectra can help in calculating the SEU
rate. Again, the LET spectra are calculated using both the differential and integral
method.




Figure 9: Orbit Averaged Spacecraft Shielded Proton Spectra      Figure 10: Orbit Averaged LET Spectra




Results of SPENVIS

                                              90 Days         120 Days
                                 Orbit 1       0 mils           0 mils
                                 Orbit 2       6 mils          10 mils
                                 Orbit 3      16 mils          20 mils
                              Table 3: Results of Needed Shielding

Notice that Orbit 3 needs the highest amount of shielding to protect against radiation.
This higher rad dosage comes from the higher latitude orbit where the solar protons at the
higher latitude access the atmosphere along field lines. From the LET spectra it is clear
that there will be a SEE rate of greater than one per day.




Joe Hackel                                       Page 15                                        8/8/2011
In comparison, the CREME96 model shows the total dose for the space station orbital
parameters (351 km @ 51 deg) behind 26 mils of aluminum, for a 2 month time period, is
484 Rad(Si). The total dose at 400 km is going to be increased to about 500-525 rad(Si).
Assuming a 2X design margin, the parts will see about 1 kRad(Si) total dose behind 26
mils of aluminum. [Ed Note: Figures unable to be gotten because of Kevin Miller being
on vacation. Hopefully, figures will be included soon]


Conclusion
The analysis of the orbital debris and radiation environments demonstrate that there is a
definite need for shielding, preferably 20 mils. This value is less than the amount
suggested by the CREME96 model. While it is known that the CREME96 model is built
conservatively, the SPENVIS model is probably not. Still, this amount should maintain
the reliability of the electronics over the mission lifetime, regardless of the orbit.
However, the amount of shielding is not great and benefits from the limited amount of
time of the mission. This is true although solar max is approaching and 3CS will have to
operate during a period of the highest solar flux. Some properties of 3CS and its mission
such as the small cross-sectional area and short design mission lifetime help mitigate the
high risks (even in Orbit 3). There will be Single Event Events and some very minor
material/solar array degradation since these can not be avoided, but measures will be
taken to lessen the severity of their affects. All of the calculations assume a 4 shield,
which is a worst-case assumption-- an actual electronics box with 20 mils aluminum will
probably provide more than 20 mils shielding. The vast majority of electronic parts (with
a few notable exceptions) can withstand 1 kRad (Si) total dose just fine, and so there is
margin (2X margin + shielding margin).




Joe Hackel                               Page 16                                  8/8/2011

				
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