417-R-RPT-0027 The Radiation Environment for Electronic Devices-1.pdf

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The Radiation Environment for Electronic Devices on the

GOES-R Series Satellites

Prepared By:

See original signature page

Janet Barth Date

Author/Radiation Effects and Analysis Group/Code 561

NASA/Goddard Space Flight Center

Greenbelt, Maryland

Reviewed By:

Michael Xapsos Date

Radiation Effects and Analysis Group/Code 561

NASA Goddard Space Flight Center

Greenbelt, MD 20771

Approved By:

Alexander Krimchansky Date

GOES R Mission Systems Manager/Code 417

NASA/Goddard Space Flight Center

Greenbelt, Maryland

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TABLE OF CONTENTS

I. Introduction

II. Radiation Environment

III. Description of Radiation Effects

A. TOTAL IONIZING DOSE

B. DISPLACEMENT DAMAGE

C. SINGLE EVENT EFFECTS

D. RELEVANT SPACE RADIATION ENVIRONMENTS

IV. The GOES Mission

V. Total Dose and Degradation Analysis

A. DEGRADATION ENVIRONMENTS

1. The Plasma Environment [3]

2. High Energy Particles – Spacecraft Incident Fluences

3. High Energy Particles – Shielded Fluences

B. TOTAL DOSE ESTIMATES

1. Top Level Ionizing Dose Estimates

2. Dose at Specific Spacecraft Locations

C. DISPLACEMENT DAMAGE ESTIMATES

VI. Single Event Effects Analysis

A. HEAVY ION INDUCED SINGLE EVENT EFFECTS

1. Galactic Cosmic Rays

2. Solar Heavy Ions

B. PROTON INDUCED SINGLE EVENT EFFECTS

1. Trapped Protons

2. Solar Protons

VII. Instrument Interference

VIII. Spacecraft Charging and Discharging

IX. Summary

X. References

Appendices ...................................................................................................................... A-1

Glossary……………...…………………………………………………...………..….A-

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I. Introduction

The purpose of this document is to define the radiation environment for the evaluation of degradation due to total ionizing and non-ionizing dose and of single event effects (SEEs) for the GOES instruments and spacecraft. The analysis took into account the radiation exposure for the nominal 15-year mission (10 years operating & 5 years on-orbit storage) at geosynchronous orbit.

Orbit: Geosynchronous, H = 35790/35790 km, 15-year Mission Requirement including 10 years of operation and up to 5 years of on-orbit storage. On-orbit locations are 75 and 135 degrees W longitude. The storage position is

105 degrees W longitude.

The radiation environment effects that must be considered, and their level of severity for the GOES missions are listed in Table 1.

Table 1: Radiation Effects in Geosynchronous

Effect Level of Severity

Spacecraft Charging/Discharging Severe

Single Event Effects Severe

Total Ionizing Dose Moderate

Displacement Damage Non-shielded (Solar Cells, Sensors) – Severe

Shielded (Optoelectronics, Shielded Sensors) – Moderate

II. Radiation Environment

The natural space radiation environment of concern for damage to spacecraft electronics is classified into two populations, 1) the transient particles which include protons and heavier ions of all of the elements of the periodic table, and 2) the trapped particles which include protons, electrons and heavier ions. The trapped electrons have energies up to 10 MeV and the trapped protons and heavier ions have energies up to 100s of MeV. The transient radiation consists of galactic cosmic ray particles and particles from solar events (coronal mass ejections and flares).

The cosmic rays have low-level fluxes with energies up to TeV. The solar eruptions periodically produce energetic protons, alpha particles, heavy ions, and electrons. The solar protons have energies up to 100’s of MeV and the heavier ions reach the GeV range. All particle fluxes are isotropic and omnidirectional to the first order.

Space also contains low energy plasma of electrons and protons with fluxes up to 1012 cm2/sec. The plasmasphere environment and the low energy (< 0.1 MeV) component of the charged particles are a concern in the near-earth environment. In the outer regions of the magnetosphere and in interplanetary space, the plasma is associated with the solar wind. Because of its low energy, thin layers of material easily stop the plasma so it is not a hazard to most spacecraft electronics. However, it is damaging to surface materials and differentials in the plasma environment can contribute to spacecraft surface charging and discharging problems [1,2].

III. Description of Radiation Effects

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Radiation effects that are important to consider for instrument and spacecraft design fall roughly into three categories: degradation from total ionizing dose (TID), degradation from displacement damage, and single event effects (SEEs).

A. Total Ionizing Dose

Total ionizing dose in electronics is a cumulative long term ionizing damage due to protons and electrons. It causes threshold shifts, leakage current and timing skews. The effect usually first appears as parametric degradation of the device and ultimately results in functional failure. It is possible to reduce TID with shielding material that absorbs most electrons and lower energy protons. As shielding is increased, shielding effectiveness decreases because of the difficulty in slowing down the higher energy protons. When a manufacturer advertises a part as “rad-hard”, he is almost always referring to its total ionizing dose characteristics. Rad-hard does not usually imply that the part is hard to non-ionizing dose or single event effects.

B. Displacement Damage

Displacement damage is cumulative long-term non-ionizing damage due to protons, electrons, and neutrons. The particles produce defects in optical materials that result in charge transfer degradation. These defects affect the performance of optocouplers (often a component in power devices), solar cells, CCDs, and linear bipolar devices.

The effectiveness of shielding depends on the location of the device. For example, coverglasses over solar cells reduce electron damage and proton damage by absorbing the low energy particles. Increasing shielding, however, is not usually effective for optoelectronic components because the high-energy protons penetrate the most feasible spacecraft electronic enclosures. For detectors in instruments it is necessary to understand the instrument geometry to determine the vulnerability to the environment.

C. Single Event Effects

Single event effects (SEE) occur as a result of charge being generated along the path of primary or secondary ionizing particles, and then being collected on circuit nodes and disrupting normal circuit response. In most cases, SEE are caused by heavier ions. However, for some devices, protons can induce SEE—in some cases through direct ionization by the proton, or, more commonly, by ionization by secondary particles produced in collisions between the proton and a nucleus in the device material. Some single event effects are non-destructive as in the case of single event upsets (SEUs), single event transients (SETs), multiple bit errors (MBEs), single event hard errors

(SHEs), etc. Single event effects can also be destructive as in the case of single event latchups (SELs), single event gate ruptures (SEGRs), and single event burnouts (SEBs). The severity of the effect can range from noisy data to loss of the mission, depending on the type of effect and the criticality of the system in which it occurs. Shielding is not an effective mitigator for single event effects because they are induced by very penetrating high-energy particles.

The preferred method for dealing with destructive failures is to use SEE-hard parts. When SEE-hard parts are not available, latchup protection circuitry is sometimes used in conjunction with failure mode analysis. However, this approach should be adopted cautiously, as even SEL events that do not result in device destruction have been

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D. Relevant Space Radiation Environments

Total ionizing dose is caused primarily by protons and electrons trapped in the Van Allen belts and solar event protons. As electrons are slowed down, their interactions with orbital electrons of the shielding material produce a secondary photon radiation known as bremsstrahlung. Generally, the dose due to galactic cosmic ray ions and proton secondaries is negligible in the presence of the other sources. For surface degradation, it is also important to include the effects of very low energy particles.

Single event effects can be induced by heavy ions (solar events and galactic cosmic rays) and, in some devices, protons (trapped and solar events) and neutrons. Displacement damage is primarily due to trapped and solar protons and also neutrons that are produced by interactions of primary particles with the atmosphere and spacecraft materials. For lightly shielded applications, such as solar arrays, displacement damage may also occur due to energetic trapped electrons. Spacecraft charging can occur on the surface of the spacecraft due to low energy electrons. Deep dielectric charging occurs when high energy electrons penetrate the spacecraft and collect in dielectric materials.

IV. The GOES Mission

The GOES spacecraft will be launched out to a geosynchronous orbit via a trajectory that has not yet been defined.

Once GOES is at its geosynchronous orbit, its mission requirement is 15 years, with 10 years operational at 135W or

75W longitude and 5 years parking at 105W longitude. Because of its lengthy timescale, the GOES mission will include satellites launched throughout the solar cycle. For this reason, it is prudent to assume that the mission will occur during during worst-case environmental conditions. In terms of TID, displacement damage and spacecraft charging threats, this is the active phase of the solar cycle. In terms of single-event effects, the appropriate galactic cosmic ray (GCR) environment to assume is that for solar minimum, since GCR fluxes are slightly higher for this portion of the solar cycle. During the active phase of the sun, the likelihood that the spacecraft will be exposed to particles from solar events (either solar flare or coronal mass ejections) increases significantly. Based on an average

11-year solar cycle, the GOES mission will encounter 11 years of solar active conditions. Figure 1 shows a projection of the solar cycle during the GOES mission, it is based on the solar activity data of solar cycles 22 and 23.

The geosynchronous radiation environment encountered by the GOES will consist of protons, electrons and heavier ions from solar events, galactic cosmic ray heavy ions, and solar wind plasma consisting of low energy protons, electrons, and heavier ions.

In avionics applications it is necessary to consider neutrons that are produced by interactions of primary particles with the atmosphere.

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Figure 1: Projection of the solar activity during the GOES mission

V. Total Dose and Degradation Analysis

The total ionizing dose accumulation causes performance degradation and failure on memories, power converters, etc. Non-ionizing energy loss in materials (atomic displacement damage) causes degradation of solar cells, optoelectronics, and detectors. The low energy particles also contribute to the erosion of surfaces.

A. Degradation Environments

In geosynchronous orbits low energy particles (< 40 keV) from the solar wind plasma contribute to the degradation of surface materials. The higher energy particles trapped in the Van Allen belts and from high-energy solar events can penetrate solar cell cover glasses and solar array substrate structures and, therefore, are responsible for solar cell degradation. Microelectronics components are also susceptible to high energy particles even though they are usually inside box enclosures.

1. The Plasma Environment [3]

The geosynchronous environment will charge spacecraft exterior surfaces. Since different materials are used and since sunlight can illuminate only one side at a time, there will always be some differential charging as well as absolute charging. The effect of this surface charging on the performance of spacecraft must be evaluated in terms of malfunctions, upsets and failures. Surface charging could disrupt environmental measurements on scientific spacecraft where control of electrostatic fields is required. For this reason, material selection to minimize differential charging is recommended. If the charging analysis indicates differential potentials of less than 500V, there should be no spacecraft discharge problems. However, if predicted potentials on materials exceeds 500V, the NASA

Charging Analyzer Program (NASCAP) can be used. The worst-case geosynchronous plasma environment is given in Table 2 and can be used as input to the NASCAP program.

Table 2: Worst-Case Geosynchronous Plasma Environment

Electron number density, NE, cm-3 1.12

D e c -8

D e c -8

D e c -9

D e c -9

D e c -0

D e c -0

D e c -0

D e c -1

D e c -1

D e c -2

D e c -2

D e c -2

D e c -3

M o n th ly s u n s p o t n u m b e rs

Monthly

Smoothed predicted forecasts

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Electron temperature, TE, eV 1.2x104

Ion number density, NI, cm-3 2.36x10-1

Ion temperature, TI, eV 2.95x104

Surface charging also increases contamination. The contaminants are attracted back to charged surfaces and deposit on them. This changes the surface characteristics. Altered surface optical properties result in higher temperatures.

Changes in secondary and photoelectron yields result in altered charging characteristics. Deposition of the dielectric contaminants can also reduce surface conductivity. If severe discharges were to occur on the surfaces, the materials can be damaged which can change their thermal control performance. Table 3 lists the acceptable and unacceptable surface coatings and materials for spacecraft use.

Table 3: Surface Coatings and Materials

Acceptable Material/Coating Unacceptable Material/Coating

Paint(carbon black) Anodyze

GSFC NS43 paint(yellow) Fiberglass material

Indium tin oxide Paint(white)

Zinc ortho-titanate paint(white) Mylar(uncoated)

Alodyne Teflon(uncoated)

Kapton(uncoated)

Silica cloth

Quartz and glass surfaces(if possible to avoid)

2. High Energy Particles – Spacecraft Incident Fluences

The spacecraft incident proton fluence levels given in this document are most often used for standard solar cell analyses that take into account the coverglass thickness of the cell. There are two possible sources of high-energy particles: trapped electrons encountered in the geosynchronous orbit and protons from solar events that can occur anytime during the 15 years of the mission. The proton fluence levels are also used to determine displacement damage effects, however, most analysis methods require that the surface incident particles be transported through the materials surrounding the sensitive components. The proton fluences behind nominal aluminum shield thicknesses are given in Section V.A.3.

The trapped particle fluxes were estimated with NASA’s AP-8 [4] model for protons and AE-8 [5] model for electrons. The models come in solar minimum and maximum versions. The uncertainty factors defined for the models are a factor of 2 for the AP-8 and 2 to 5 for the AE-8. These uncertainty factors apply to long-term averages expected over a 6-month mission duration. Daily values can fluctuate by two to three orders of magnitude depending on the level of activity on the sun and within the magnetosphere.

The solar proton levels can now be estimated from the new Emission of Solar Proton (ESP) model [6]. Previously, estimates of solar proton levels were obtained from models [7,8] that were largely empirical in nature, making it difficult to add data to the model from more recent solar cycles. The ESP model is based on satellite data from solar cycles 20, 21, and 22. The distribution of the fluences for the events is obtained from maximum entropy theory, and design limits in the worst case models are obtained from extreme value theory. The solar proton predictions are not linear over time; therefore, the levels given in this document may be invalid if extrapolated for longer mission durations.

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-define the spacecraft incident radiation environment

-develop a particle attenuation model using theoretical shielding configurations

(similar to dose-depth curves).

3) Obtain results for each sensitive location:

-divide the structural model into solid angle sectors

-ray trace through the sectors to calculate the material mass distribution

-use the ray trace results to calculate total doses from the particle attenuation model.

Once the basic structural model has been defined, total doses can be obtained for any location in the spacecraft in a short time (in comparison to Monte Carlo methods). The value of dose mitigation measures can be accurately evaluated by adding the changes to the model and recalculating the total dose. For spacecraft with strict weight budgets, the 3-D ray trace method, the total dose design requirement can be defined at a box or instrument level avoiding unnecessary use of expensive or increasingly unavailable radiation hardened parts.

As the design of the GOES evolves, it may become necessary to estimate the doses at specific locations in the spacecraft or instruments. Often the dose requirement can be met by modeling the surrounding electronic box only or by modeling only the instrument.

C. Displacement Damage Estimates

Long-term damage due to atomic displacements (displacement damage) degrades solar cells, optoelectronics, imaging devices such as CCDs, and some bipolar technologies. The displacement damage is caused by exposure of the components to protons and electrons. In the geosynchronous environment, the threat ranges from moderate to severe depending on the amount of shielding surrounding the components. Sensors that are fully exposed to the space environment are especially susceptible.

Displacement damage is evaluated by combining the shielded proton energy spectra given in Section V.A.3 for the material and the results of laboratory irradiation of the devices sensitive to atomic displacement damage. The level of the hazard is highly dependent on the device type and can be process specific. For the GOES mission, it is important to keep in mind that some optoelectronic devices experience enough damage during one large solar proton event to cause the device to fail. It is necessary that the parts list screening for radiation also include a check for devices that are susceptible to displacement damage.

This document does not address solar array degradation due to the space radiation environment. However, the material that is described in the JPL Publication 96-9 titled “GaAs Solar Cell Radiation Handbook”[9], dated July 1, 1996, contains useful material that can be used by the spacecraft contractor as a guide when preparing his solar array degradation model.

VI. Single Event Effects Analysis

A. Heavy Ion Induced Single Event Effects

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Some electronic devices are susceptible to single event effects (SEEs), e.g., single event upsets, single event latch-up, single event burn-out. Because of their ability to penetrate to the sensitive regions of devices and their ability to ionize materials, heavy ions cause SEEs by the direct deposit of charge. The quantity most frequently used to measure an ion’s ability to deposit charge in devices is linear energy transfer (LET). Heavy ion abundances are converted to total LET spectra. Once specific parts are selected for the mission and, if necessary, characterized by laboratory testing, the LET spectra for the heavy ions are integrated with the device characterization to calculate

SEE rates. Heavy ion populations that have sufficient numbers to be a SEE hazard are the galactic cosmic rays and those from solar events.

1. Galactic Cosmic Rays

The cosmic ray fluxes for elements hydrogen through uranium were used to calculate daily LET spectra for 100 mils nominal aluminum shielding as given in Table A8 and Figure 9. The range of the cosmic ray abundances is bounded by the extremea of the solar active and inactive phases of the solar cycle with the highest values occurring during the solar inactive phase and the lowest during the solar active phase. With the extended mission goal of 15 years, the highest values should be used for single event effects analyses. The LET fluence values are given for the highest and lowest point of the solar cycle. The CREME96 [10] model was used to obtain the cosmic ray heavy ion abundances. This model has an accuracy of 25-40%.

Integral LET Spectra at 1AU (Z=1-92) for Interplanetary Galactic Cosmic Rays

100 mils Aluminum Shielding, CREME96

1 00E-10

1 00E-09

1 00E-08

1 00E-07

1 00E-06

1 00E-05

1 00E-04

1 00E-03

1 00E-02

1 00E-01

1.00E+00

1.00E+01

1.00E+02

1.00E+03

1.00E+04

1.00E+05

1.00E+06

1 E-03 1.E-02 1 E-01 1.E+00 1.E+01 1.E+02 1.E+03

LET energy (MeV-cm2/mg)

L E

T f lu e n c e

/c m -d a y

Solar minimum

Solar maximum

Figure 9: Integral LET spectra are shown for galactic cosmic ray ions hydrogen through uranium.

2. Solar Heavy Ions

The heavy ions from solar flares and coronal mass ejections can also produce single event effects. The solar event fluxes for the elements hydrogen through uranium were used to calculate daily LET spectra for 100 mils nominal aluminum shielding in units of average LET flux per second. The intensity of the fluxes varies over the duration of an event; therefore, values are averaged over the worst week of the solar cycle, the worst day of the solar cycle, and

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GOES mission. The CREME96 model was also used to calculate the solar heavy ion levels. An uncertainty factor for the solar heavy ion model has not been released.

Integral LET Spectra at 1 AU (Z=1-92) for Interplanetary Solar Particle Events

100 mils Aluminum Shielding, CREME96

1 00E-11

1 00E-10

1 00E-09

1 00E-08

1 00E-07

1 00E-06

1 00E-05

1 00E-04

1 00E-03

1 00E-02

1 00E-01

1.00E+00

1.00E+01

1.00E+02

1.00E+03

1.00E+04

1.00E+05

1.00E+06

1.00E-03 1.00E-02 1 00E-01 1 00E+00 1 00E+01 1.00E+02 1.00E+03

LET Energy (MeV-cm /mg)

L E

T F lu e n c e

/c m -s

Average Over Peak

Average Over Worst Day

Average Over Worst Week

Figure 10: Integral LET spectra are shown for hydrogen through uranium for the October 1989 solar particle event.

B. Proton Induced Single Event Effects

In some devices, single event effects are also induced by protons. Protons from the trapped radiation belts and from solar events do not generate sufficient ionization (LET < 1 MeV-cm2 /mg) to produce the critical charge necessary for SEEs to occur in most electronics. More typically, protons cause Single Event Effects through secondary particles via nuclear interactions, that is, spallation and fractionation products. Because the proton energy is important in the production (and not the LET) of the secondary particles that cause the SEEs, device sensitivity to these particles is typically expressed as a function of proton energy rather than LET.

1. Trapped Protons

At geosynchronous orbit, the trapped protons have energies < 2 MeV. Therefore, they are not a significant factor in producing interference or damage in microelectronics.

2. Solar Protons

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Protons from solar events will also be a single event effects hazard for the GOES spacecraft. These enhanced levels of protons could occur anytime during the 15-year mission but are most likely during the portion of the mission that occurs during the active phase of the solar cycle. As with the solar heavy ion LET spectra, solar proton fluxes are averaged over worst day, worst week, and the peak of the October 1989 solar event. The proton flux averages for a nominal 100 mils of shielding are given in Table A10 and are shown in Figure 11.

1 .0 E - 0 3

1 .0 E - 0 2

1 .0 E - 0 1

1 . 0 E + 0 0

1 . 0 E + 0 1

1 . 0 E + 0 2

1 . 0 E + 0 3

1 . 0 E + 0 4

1 . 0 E + 0 5

1 1 0 1 0 0 1 0 0 0

E n e rg y ( M e V )

P a rt ic le s

/c m

/s

/M e

V

A v e ra g e O v e r P e a k

A v e ra g e O v e r W o rs t D a y

A v e ra g e O v e r W o rs t W e e k

V a lu e s D o N o t I n c lu d e D e s ig n M a r g in

D iff ere n tia l S o lar P r o t o n E v en t F lu xe s a t 1 A U

10 0 m ils A lu m in u m S h ie ld in g , C R E M E 96

Figure 11: Solar proton fluxes for single event effects evaluation.

VII. Instrument Interference

The particle background causes increased noise levels in instruments and other electronics. This can be a concern if low noise levels are required for instrument observations, for example. The particle background of concern in GEO comes from galactic cosmic rays, trapped electrons, and solar particle events. The galactic cosmic ray background is fairly steady and varies slowly with the solar cycle. The trapped electron environment is very volatile at GEO altitudes and the electron fluxes can vary by several orders of magnitude over periods as short as a month. Solar particle events typically last from days to weeks. The particle fluxes during these events vary continuously with time and can reach values that are comparable to trapped electron fluxes. Both long-term and worst-case estimates of these particle fluxes are given in Table 4 for a nominal shielding of 100 mils of aluminum.

Table 4: Long-term and worst-case particle fluxes in GEO behind 100 mils of aluminum shielding

Radiation: Long-term flux (#/cm2/s): Worst-case flux (#/cm2/s):

galactic cosmic rays 2.5 4.6 trapped electrons 6.7 x 104 1.3 x 106 solar particle events --- 2.0 x 105

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No long-term flux is included for solar particle events because the events occur over short periods of time compared to the length of the mission. Such a long-term flux would have little relevance for the viewing and data collection activities on GOES. However, particle interference during solar events is of particular concern because it can impact the observation times of instruments. To present a more realistic picture of the solar particle event flux variations, solar proton flux data from the Space Environment Monitor (SEM) of GOES are presented in Figures

12-14. The elemental composition of a solar particle event consists of about 95% protons, on average. These figures represent the hourly average flux of >50 MeV protons as a function of time for the years 1989, 1990 and

1991. This was a very active period during the solar maximum phase of the last complete solar cycle, #22. Note that the scale for flux in these figures is logarithmic.

0 50 100 150 200 250 300 350

Hourly Average of Solar Proton Flux > 50 MeV for 1989

F lu x

M e V

P a rt ic le s /c m ^2

/s e c

Day of Year

GOES Space Environment Monitor Data

Figure 12: Hourly averages of solar proton fluxes for > 50 MeV protons for 1989.

0 50 100 150 200 250 300 350

Hourly Average of Solar Proton Flux > 50 MeV for 1990

F lu x

M e V

P a rt ic le s /c m ^2

/s e c

Figure 13: Hourly averages of solar proton fluxes for > 50 MeV protons for 1990.

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0 50 100 150 200 250 300 350

Hourly Average of Solar Proton Flux > 50 MeV for 1991 F lu x

M e V

P a rt ic le s /c m ^2

/s e c

Figure 14: Hourly averages of solar proton fluxes for > 50 MeV protons for 1991.

VIII. Spacecraft Charging and Discharging

Surface charging and deep dielectric charging must also be evaluated for the GOES mission. The spacecraft can accumulate high levels of electron build-up on spacecraft surfaces (low energy electrons) and in the dielectrics (high energy electrons). The particle accumulation profiles must be analyzed for possible surface and deep dielectric charging effects. The average electron accumulation profiles on a daily basis are shown in Figure 15. However, it should be remembered that these are long-term, average fluxes predicted by the AE8 Model and that on a given day the electron flux may be substantially higher. Measurements made on the GOES satellites indicate that daily flux levels can exceed those predicted by AE8 by a factor of about 20. When the transfer trajectory for the GOES spacecraft is defined, it also needs to be evaluated for levels of charging environments. The transfer trajectory will likely have a higher charging environment than at geostationary because of possible trajectory passes through more intense regions of the outer zone electron belt. Guidelines for controlling surface charging effects are given in reference [3] and guidelines for controlling deep dielectric charging are given in reference [11].

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GOES Daily Integral Electron Levels

(Values Do Not Include Design Margins)

1.E+05

1.E+06

1.E+07

1.E+08

1.E+09

1.E+10

1.E+11

1.E+12

1.E+13

0.01 0.10 1.00 10.00

Energy (> MeV)

E le c tr o n s

/c m /D a y

135W

105W

75W

Figure1

5: Daily integral electron levels for GOES.

IX. Summary

A top-level radiation environment specification was presented for the GOES mission. Although the environment is considered “moderate”, the environment poses challenges to mission designers because of its highly variable nature caused by activity on the sun.

Spacecraft and instrument designers must be made aware that some newer technologies and commercial-off-the-shelf (COTS) devices are very soft to radiation effects. COTS devices that lose functionality at 5 krads of dose are not uncommon. Also, one extremely large solar proton event can cause enough displacement damage degradation in some optocoupler devices to cause failure. Increasingly, single event effects require careful part selection and mitigation schemes. With its full exposure to galactic cosmic heavy ions and particles from solar events, GOES must have a carefully planned radiation engineering program.

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X. References

[1] A. Holmes-Siedle and L. Adams, Handbook of Radiation Effects, p. 16, Oxford University Press, Oxford, 1993.

[2] A. R. Frederickson, “Upsets Related to Spacecraft Charging,” IEEE Trans. on Nucl. Science, Vol. 43, No. 2, pp. 426-441, April 1996.

[3] C.K. Purvis, H.B. Garrett, A.C. Wittlesey and N.J. Stevens, “Design Guidelines for Assessing and

Controlling Spacecraft Charging Effects”, NASA Technical Paper 2361, Sept. 1984.

[4] D. M. Sawyer and J. I. Vette, “AP-8 Trapped Proton Environment,” NSSDC/WDC-A-R&S 76-06, NASA/Goddard Space Flight Center, Greenbelt, MD, December 1991.

[5] J. I. Vette, “The AE-8 Trapped Electron Model Environment,” NSSDC/WDC-A-R&S 91-24, NASA/Goddard Space Flight Center, Greenbelt, MD, November 1991.

[6] M. A. Xapsos, J. L. Barth, E. G. Stassinopoulos, G. P. Summers, E.A. Burke, G. B. Gee, “Model for

Prediction of Solar Proton Events”, to be published in Proceedings of the 1999 Space Environment and

Effects Workshop, Farnborough, UK.

[7] E. G. Stassinopoulos, “SOLPRO: A Computer Code to Calculate Probabilistic Energetic Solar Flare

Protons,” NSSDC 74-11, NASA/Goddard Space Flight Center, Greenbelt, MD, April 1975.

[8] J. Feynman, T. P. Armstrong, L. Dao-Gibner, and S. Silverman, “New Interplanetary Proton Fluence

Model,” J. Spacecraft, Vol. 27 No. 24, pp 403-410, July-August 1990.

[9] Anspaugh, B.E., “GaAs Solar Cell Radiation Handbook”, JPL Publication 96-9, July 1, 1996.

[10] A. J. Tylka, J. H. Adams, Jr., P. R. Boberg, W. F. Dietrich, E.O. Flueckiger, E.L. Petersen, M.A. Shea, D.F.

Smart, and E.C. Smith, “CREME96: A Revision of the Cosmic Ray Effects on Micro-Electronics Code: to be published in IEEE Trans. On Nuc. Sci., December 1997.

[11] “Avoiding Problems Caused by Spacecraft On-Orbit Internal Charging Effects”, NASA Technical

Handbook NASA-HDBK-4002, Feb. 1999.

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-1

To verify the correct version of this document, please contact the GOES R Series Configuration

Management Office

Appendices

Table A1

Spacecraft Incident Integral Solar Proton Fluences (#/cm2) for 15 Year Mission (11 Solar Active Years)

Values Do Not Include Design Margins

Energy

Confidence

>MeV 90% 95%

1 1.42E+12 1.60E+12

3 6.39E+11 7.85E+11

5 4.19E+11 5.43E+11

7 3.08E+11 4.14E+11

10 2.15E+11 3.01E+11

15 1.34E+11 1.98E+11

20 9.23E+10 1.42E+11

25 6.72E+10 1.06E+11

30 5.09E+10 8.21E+10

35 3.96E+10 6.52E+10

40 3.15E+10 5.28E+10

45 2.56E+10 4.33E+10

50 2.10E+10 3.62E+10

55 1.76E+10 3.04E+10

60 1.49E+10 2.59E+10

70 1.08E+10 1.91E+10

80 8.15E+09 1.46E+10

90 6.28E+09 1.13E+10

100 4.94E+09 8.98E+09

125 2.97E+09 5.42E+09

150 1.93E+09 3.51E+09

175 1.32E+09 2.39E+09

200 9.43E+08 1.72E+09

225 6.96E+08 1.27E+09

250 5.25E+08 9.57E+08

275 4.06E+08 7.38E+08

300 3.19E+08 5.80E+08

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-2

Table A2

Spacecraft Incident Integral Electron Fluences (#/cm2) for GOES Mission

(>MeV)

10yrs at

135W +

5yrs at 105W

10 yrs at 75W

5yrs at 105W

10 yrs at

135W

5yrs at 105W

10 yrs at 75W

0.04 2.05E+16 1.80E+16 1.41E+16 6.33E+15 1.17E+16

0.07 1.68E+16 1.44E+16 1.16E+16 5.12E+15 9.31E+15

0.10 1.38E+16 1.16E+16 9.60E+15 4.15E+15 7.43E+15

0.20 6.72E+15 5.49E+15 4.73E+15 1.99E+15 3.50E+15

0.30 3.59E+15 2.86E+15 2.54E+15 1.05E+15 1.81E+15

0.40 2.09E+15 1.63E+15 1.49E+15 6.02E+14 1.03E+15

0.50 1.21E+15 9.29E+14 8.67E+14 3.46E+14 5.83E+14

0.60 8.19E+14 5.95E+14 5.93E+14 2.26E+14 3.69E+14

0.70 5.54E+14 3.81E+14 4.06E+14 1.48E+14 2.33E+14

0.80 3.84E+14 2.58E+14 2.83E+14 1.01E+14 1.57E+14

0.90 2.73E+14 1.87E+14 2.00E+14 7.27E+13 1.15E+14

1.00 1.93E+14 1.36E+14 1.41E+14 5.23E+13 8.33E+13

1.25 9.58E+13 6.55E+13 7.03E+13 2.55E+13 4.00E+13

1.50 4.74E+13 3.15E+13 3.50E+13 1.24E+13 1.91E+13

1.75 2.69E+13 1.72E+13 2.00E+13 6.87E+12 1.03E+13

2.00 1.53E+13 9.40E+12 1.15E+13 3.81E+12 5.59E+12

2.25 7.73E+12 4.62E+12 5.84E+12 1.89E+12 2.73E+12

2.50 3.93E+12 2.27E+12 2.99E+12 9.41E+11 1.33E+12

2.75 2.34E+12 1.40E+12 1.77E+12 5.73E+11 8.24E+11

3.00 1.41E+12 8.59E+11 1.06E+12 3.49E+11 5.10E+11

3.25 9.11E+11 5.54E+11 6.86E+11 2.25E+11 3.29E+11

3.50 5.88E+11 3.56E+11 4.43E+11 1.45E+11 2.11E+11

3.75 3.22E+11 2.09E+11 2.39E+11 8.31E+10 1.26E+11

4.00 1.77E+11 1.23E+11 1.29E+11 4.75E+10 7.54E+10

4.50 4.67E+10 3.31E+10 3.40E+10 1.27E+10 2.04E+10

5.00 1.56E+09 8.40E+08 1.20E+09 3.59E+08 4.81E+08

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-3

Table A3

Integral Solar Proton Fluences (#/cm2) Behind Solid Sphere Aluminum Shields

15 Year Mission (11 Active Solar Years) – 90% Confidence Level

Degraded

Shielded Solar Proton Fluences

50 mils Al

(1.27 mm)

100 mils Al

(2.54 mm)

200 mils Al

(5.08 mm)

350 mils Al

(8.89 mm)

500 mils Al

(12.7 mm)

> MeV #/cm2 #/cm2 #/cm2 #/cm2 #/cm2

0.100 1.33E+11 7.88E+10 4.35E+10 2.52E+10 1.72E+10

0.126 1.33E+11 7.88E+10 4.35E+10 2.52E+10 1.72E+10

0.158 1.33E+11 7.88E+10 4.35E+10 2.52E+10 1.72E+10

0.200 1.33E+11 7.88E+10 4.35E+10 2.52E+10 1.72E+10

0.251 1.33E+11 7.88E+10 4.35E+10 2.52E+10 1.72E+10

0.316 1.33E+11 7.88E+10 4.33E+10 2.52E+10 1.72E+10

0.398 1.33E+11 7.88E+10 4.33E+10 2.52E+10 1.72E+10

0.501 1.33E+11 7.87E+10 4.33E+10 2.52E+10 1.72E+10

0.631 1.33E+11 7.87E+10 4.33E+10 2.52E+10 1.72E+10

0.794 1.33E+11 7.85E+10 4.33E+10 2.52E+10 1.72E+10

1.000 1.32E+11 7.85E+10 4.33E+10 2.50E+10 1.72E+10

1.260 1.32E+11 7.84E+10 4.32E+10 2.50E+10 1.72E+10

1.580 1.31E+11 7.81E+10 4.32E+10 2.50E+10 1.72E+10

2.000 1.30E+11 7.78E+10 4.30E+10 2.50E+10 1.72E+10

2.510 1.29E+11 7.73E+10 4.29E+10 2.49E+10 1.71E+10

3.160 1.27E+11 7.67E+10 4.28E+10 2.49E+10 1.71E+10

3.980 1.24E+11 7.58E+10 4.24E+10 2.48E+10 1.71E+10

5.010 1.20E+11 7.44E+10 4.19E+10 2.46E+10 1.69E+10

6.310 1.15E+11 7.25E+10 4.13E+10 2.43E+10 1.68E+10

7.940 1.08E+11 6.97E+10 4.03E+10 2.39E+10 1.66E+10

10.000 9.93E+10 6.60E+10 3.91E+10 2.34E+10 1.64E+10

12.600 8.86E+10 6.09E+10 3.70E+10 2.27E+10 1.60E+10

15.800 7.59E+10 5.47E+10 3.45E+10 2.16E+10 1.53E+10

20.000 6.27E+10 4.73E+10 3.12E+10 2.01E+10 1.44E+10

25.100 4.95E+10 3.92E+10 2.72E+10 1.83E+10 1.34E+10

31.600 3.71E+10 3.08E+10 2.27E+10 1.60E+10 1.19E+10

39.800 2.67E+10 2.31E+10 1.80E+10 1.32E+10 1.02E+10

50.100 1.83E+10 1.64E+10 1.34E+10 1.04E+10 8.32E+09

63.100 1.20E+10 1.10E+10 9.42E+09 7.66E+09 6.35E+09

79.400 7.45E+09 7.00E+09 6.22E+09 5.27E+09 4.54E+09

100.000 4.41E+09 4.21E+09 3.85E+09 3.41E+09 3.04E+09

126.000 2.52E+09 2.43E+09 2.27E+09 2.06E+09 1.88E+09

158.000 1.33E+09 1.29E+09 1.23E+09 1.13E+09 1.05E+09

200.000 6.16E+08 6.02E+08 5.73E+08 5.34E+08 4.96E+08

251.000 1.95E+08 1.90E+08 1.79E+08 1.61E+08 1.46E+08

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-4

Table A4

Integral Electron Fluences (#/cm2) Behind Solid Sphere Aluminum Shields

15-Year Worst-Case Mission (10 Years At 135W + 5 Years At 105W)

Shielded Electron Fluences

50 mils Al

(1.27 mm)

100 mils Al

(2.54 mm)

200 mils Al

(5.08 mm)

350 mils Al

(8.89 mm)

500 mils Al

(12.7 mm)

> MeV #/cm2 #/cm2 #/cm2 #/cm2 #/cm2

0.010 3.18E+14 4.77E+13 2.33E+12 2.82E+10 1.03E+03

0.013 3.18E+14 4.77E+13 2.33E+12 2.82E+10 7.60E+03

0.016 3.18E+14 4.77E+13 2.33E+12 2.80E+10

0.020 3.18E+14 4.77E+13 2.32E+12 2.80E+10

0.025 3.16E+14 4.75E+13 2.31E+12 2.80E+10 6.21E+03

0.032 3.16E+14 4.74E+13 2.31E+12 2.80E+10 3.07E+03

0.040 3.14E+14 4.73E+13 2.31E+12 2.78E+10

0.050 3.12E+14 4.69E+13 2.29E+12 2.76E+10

0.063 3.08E+14 4.67E+13 2.29E+12 2.74E+10

0.079 3.04E+14 4.61E+13 2.27E+12 2.72E+10

0.100 2.98E+14 4.55E+13 2.23E+12 2.66E+10

0.126 2.89E+14 4.46E+13 2.20E+12 2.60E+10

0.158 2.77E+14 4.32E+13 2.16E+12 2.52E+10

0.200 2.60E+14 4.15E+13 2.08E+12 2.41E+10

0.251 2.38E+14 3.91E+13 1.99E+12 2.25E+10

0.316 2.11E+14 3.60E+13 1.87E+12 2.06E+10

0.398 1.78E+14 3.21E+13 1.71E+12 1.80E+10

0.501 1.41E+14 2.74E+13 1.51E+12 1.47E+10

0.631 1.05E+14 2.20E+13 1.28E+12 1.09E+10

0.794 7.11E+13 1.63E+13 1.02E+12 6.94E+09

1.000 4.31E+13 1.06E+13 7.49E+11 3.21E+09

1.260 2.43E+13 5.85E+12 4.89E+11 8.21E+08

1.580 1.14E+13 2.78E+12 2.63E+11 6.11E+07

2.000 3.86E+12 1.25E+12 1.01E+11 4.76E+05

2.510 1.33E+12 4.80E+11 2.08E+10

3.160 3.80E+11 9.59E+10 5.68E+07

3.980 4.22E+10 7.89E+08

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-5

Table A5

Integral Electron Fluences (#/cm2) Behind Solid Sphere Aluminum Shields

15 Year Mission (10 Years At 75W + 5 Years At 105W)

Shielded Electron Fluences

50 mils Al

(1.27 mm)

100 mils Al

(2.54 mm)

200 mils Al

(5.08 mm)

350 mils Al

(8.89 mm)

500 mils Al

(12.7 mm)

> MeV #/cm2 #/cm2 #/cm2 #/cm2 #/cm2

0.010 2.19E+14 3.14E+13 1.39E+12 1.97E+10 1.03E+03

0.013 2.19E+14 3.14E+13 1.39E+12 1.97E+10 2.70E+03

0.016 2.19E+14 3.14E+13 1.39E+12 1.96E+10

0.020 2.18E+14 3.14E+13 1.38E+12 1.96E+10

0.025 2.18E+14 3.14E+13 1.38E+12 1.96E+10

0.032 2.17E+14 3.13E+13 1.38E+12 1.95E+10 1.84E+03

0.040 2.16E+14 3.12E+13 1.38E+12 1.94E+10

0.050 2.14E+14 3.09E+13 1.37E+12 1.93E+10

0.063 2.12E+14 3.07E+13 1.36E+12 1.92E+10

0.079 2.09E+14 3.04E+13 1.35E+12 1.90E+10

0.100 2.05E+14 2.99E+13 1.34E+12 1.87E+10

0.126 1.99E+14 2.93E+13 1.32E+12 1.82E+10

0.158 1.90E+14 2.85E+13 1.29E+12 1.77E+10

0.200 1.79E+14 2.72E+13 1.25E+12 1.69E+10

0.251 1.65E+14 2.55E+13 1.19E+12 1.58E+10

0.316 1.46E+14 2.35E+13 1.12E+12 1.44E+10

0.398 1.23E+14 2.08E+13 1.03E+12 1.26E+10

0.501 9.78E+13 1.76E+13 9.13E+11 1.03E+10

0.631 7.20E+13 1.40E+13 7.78E+11 7.67E+09

0.794 4.81E+13 1.01E+13 6.24E+11 4.76E+09

1.000 2.85E+13 6.46E+12 4.62E+11 2.18E+09

1.260 1.55E+13 3.46E+12 3.04E+11 5.38E+08

1.580 6.92E+12 1.64E+12 1.68E+11 3.68E+07

2.000 2.27E+12 7.62E+11 6.94E+10

2.510 8.06E+11 2.97E+11 1.46E+10

3.160 2.38E+11 6.83E+10 3.64E+07

3.980 3.02E+10 4.44E+08

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-6

Table A6

Total Ionizing Dose (RadsSi) at the Center of Solid Aluminum Spheres for 15-Year Worst-Case Mission

(10 Years At 135W + 5 Years At 105W, 11 Active Years At 90% Confidence Level)

Aluminum Shield Thickness

Total

Dose

Solar

Proton

Dose

Bremsstrahlung

Photon

Electron g/cm2 mm mils RadsSi RadsSi RadsSi RadsSi

0.03 0.11 4.37 3.91E+08 8.26E+05 6.39E+04 3.90E+08

0.04 0.15 5.83 3.09E+08 6.44E+05 6.91E+04 3.08E+08

0.05 0.19 7.29 2.45E+08 5.37E+05 7.13E+04 2.44E+08

0.06 0.22 8.75 1.97E+08 4.54E+05 7.14E+04 1.97E+08

0.08 0.30 11.67 1.38E+08 3.54E+05 6.88E+04 1.38E+08

0.10 0.37 14.58 1.01E+08 2.88E+05 6.45E+04 1.00E+08

0.15 0.56 21.87 5.04E+07 1.96E+05 5.27E+04 5.01E+07

0.20 0.74 29.16 2.96E+07 1.51E+05 4.31E+04 2.94E+07

0.30 1.11 43.74 1.28E+07 9.98E+04 3.10E+04 1.27E+07

0.40 1.48 58.31 6.34E+06 7.50E+04 2.37E+04 6.24E+06

0.50 1.85 72.91 3.47E+06 5.73E+04 1.92E+04 3.39E+06

0.60 2.22 87.48 2.03E+06 4.78E+04 1.61E+04 1.96E+06

0.80 2.96 116.65 7.97E+05 3.40E+04 1.23E+04 7.51E+05

1.00 3.70 145.83 3.40E+05 2.59E+04 1.00E+04 3.04E+05

1.25 4.63 182.28 1.25E+05 1.99E+04 8.26E+03 9.67E+04

1.50 5.56 218.74 5.92E+04 1.57E+04 7.12E+03 3.64E+04

1.75 6.48 255.16 3.38E+04 1.27E+04 6.35E+03 1.47E+04

2.00 7.41 291.61 2.18E+04 1.07E+04 5.70E+03 5.48E+03

2.50 9.26 364.53 1.33E+04 8.02E+03 4.91E+03 4.06E+02

3.00 11.11 437.40 1.04E+04 6.15E+03 4.29E+03 3.42E+00

3.50 12.96 510.24 8.77E+03 4.92E+03 3.84E+03

4.00 14.81 583.07 7.58E+03 4.08E+03 3.50E+03

4.50 16.67 656.30 6.66E+03 3.42E+03 3.24E+03

5.00 18.52 729.13 5.94E+03 2.93E+03 3.02E+03

5.50 20.37 801.97 5.38E+03 2.53E+03 2.85E+03

6.00 22.22 874.80 4.88E+03 2.18E+03 2.70E+03

6.50 24.07 947.64 4.46E+03 1.93E+03 2.53E+03

7.00 25.93 1020.87 4.15E+03 1.73E+03 2.42E+03

7.50 27.78 1093.70 3.85E+03 1.54E+03 2.31E+03

8.00 29.63 1166.54 3.59E+03 1.38E+03 2.21E+03

8.50 31.48 1239.37 3.35E+03 1.25E+03 2.10E+03

9.00 33.33 1312.20 3.13E+03 1.12E+03 2.01E+03

9.50 35.19 1385.43 2.96E+03 1.01E+03 1.95E+03

10.00 37.04 1458.27 2.73E+03 8.74E+02 1.86E+03

12.50 46.30 1822.83 2.20E+03 6.38E+02 1.56E+03

15.00 55.56 2187.40 1.77E+03 4.63E+02 1.30E+03

20.00 74.07 2916.14 1.17E+03 2.77E+02 8.92E+02

25.00 92.59 3645.28 8.45E+02 1.83E+02 6.62E+02

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-7

Solar

Proton

Dose

Bremsstrahlung g/cm2 mm mils RadsSi RadsSi RadsSi RadsSi

30.00 111.10 4374.02 5.81E+02 1.23E+02 4.58E+02

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-8

Table A7

Total Ionizing Dose (RadsSi) at the Center of Solid Aluminum Spheres for 15 Year Mission

(10 Years At 75W + 5 Years At 105W, 11 Active Years At 90% Confidence Level)

Solar

Proton

Dose

Bremsstrahlung g/cm2 mm mils RadsSi RadsSi RadsSi RadsSi

0.03 0.11 4.37 3.26E+08 8.26E+05 5.45E+04 3.26E+08

0.04 0.15 5.83 2.56E+08 6.44E+05 5.84E+04 2.55E+08

0.05 0.19 7.29 2.01E+08 5.37E+05 5.98E+04 2.00E+08

0.06 0.22 8.75 1.61E+08 4.54E+05 5.97E+04 1.60E+08

0.08 0.30 11.67 1.11E+08 3.54E+05 5.71E+04 1.11E+08

0.10 0.37 14.58 8.01E+07 2.88E+05 5.31E+04 7.98E+07

0.15 0.56 21.87 3.92E+07 1.96E+05 4.31E+04 3.90E+07

0.20 0.74 29.16 2.20E+07 1.51E+05 3.49E+04 2.18E+07

0.30 1.11 43.74 8.80E+06 9.98E+04 2.47E+04 8.68E+06

0.40 1.48 58.31 4.41E+06 7.50E+04 1.88E+04 4.31E+06

0.50 1.85 72.91 2.41E+06 5.73E+04 1.52E+04 2.33E+06

0.60 2.22 87.48 1.39E+06 4.78E+04 1.27E+04 1.33E+06

0.80 2.96 116.65 5.26E+05 3.40E+04 9.68E+03 4.82E+05

1.00 3.70 145.83 2.20E+05 2.59E+04 7.88E+03 1.87E+05

1.25 4.63 182.28 8.39E+04 1.99E+04 6.49E+03 5.75E+04

1.50 5.56 218.74 4.32E+04 1.57E+04 5.58E+03 2.19E+04

1.75 6.48 255.16 2.68E+04 1.27E+04 4.94E+03 9.12E+03

2.00 7.41 291.61 1.87E+04 1.07E+04 4.45E+03 3.55E+03

2.50 9.26 364.53 1.21E+04 8.02E+03 3.80E+03 2.80E+02

3.00 11.11 437.40 9.45E+03 6.15E+03 3.30E+03 2.35E+00

3.50 12.96 510.24 7.92E+03 4.92E+03 3.00E+03

4.00 14.81 583.07 6.82E+03 4.08E+03 2.74E+03

4.50 16.67 656.30 5.94E+03 3.42E+03 2.52E+03

5.00 18.52 729.13 5.27E+03 2.93E+03 2.34E+03

5.50 20.37 801.97 4.72E+03 2.53E+03 2.20E+03

6.00 22.22 874.80 4.26E+03 2.18E+03 2.08E+03

6.50 24.07 947.64 3.88E+03 1.93E+03 1.95E+03

7.00 25.93 1020.87 3.60E+03 1.73E+03 1.87E+03

7.50 27.78 1093.70 3.31E+03 1.54E+03 1.77E+03

8.00 29.63 1166.54 3.05E+03 1.38E+03 1.67E+03

8.50 31.48 1239.37 2.85E+03 1.25E+03 1.60E+03

9.00 33.33 1312.20 2.64E+03 1.12E+03 1.52E+03

9.50 35.19 1385.43 2.47E+03 1.01E+03 1.47E+03

10.00 37.04 1458.27 2.28E+03 8.74E+02 1.41E+03

12.50 46.30 1822.83 1.78E+03 6.38E+02 1.14E+03

15.00 55.56 2187.40 1.43E+03 4.63E+02 9.72E+02

20.00 74.07 2916.14 9.75E+02 2.77E+02 6.98E+02

25.00 92.59 3645.28 6.46E+02 1.83E+02 4.64E+02

30.00 111.10 4374.02 4.60E+02 1.23E+02 3.37E+02

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-9

Table A8

Integral LET for Interplanetary Galactic Cosmic Rays (Z=1-92)

100 mils Aluminum Shielding

LET LET Fluence LET LET Fluence

MeV*sqcm/mg #/sqcm/day MeV*sqcm/mg #/sqcm/day

Solar Minimum Solar Maximum

1.00E-03 4.25E+05 1.00E-03 1.54E+05

1.65E-03 4.24E+05 1.65E-03 1.54E+05

1.69E-03 3.29E+05 1.69E-03 1.07E+05

1.70E-03 3.04E+05 1.70E-03 9.42E+04

1.72E-03 2.84E+05 1.72E-03 8.46E+04

1.77E-03 2.54E+05 1.77E-03 7.02E+04

1.81E-03 2.30E+05 1.81E-03 5.98E+04

1.85E-03 2.12E+05 1.85E-03 5.20E+04

1.91E-03 1.90E+05 1.91E-03 4.34E+04

1.98E-03 1.72E+05 1.98E-03 3.75E+04

2.01E-03 1.67E+05 2.01E-03 3.59E+04

2.13E-03 1.46E+05 2.13E-03 3.05E+04

2.28E-03 1.27E+05 2.28E-03 2.69E+04

2.53E-03 1.07E+05 2.53E-03 2.39E+04

3.01E-03 8.29E+04 3.01E-03 2.11E+04

3.54E-03 6.87E+04 3.54E-03 1.98E+04

4.52E-03 5.55E+04 4.52E-03 1.88E+04

5.56E-03 4.90E+04 5.56E-03 1.83E+04

6.54E-03 4.58E+04 6.54E-03 1.82E+04

7.52E-03 2.76E+04 7.52E-03 7.46E+03

8.55E-03 2.13E+04 8.55E-03 5.04E+03

9.60E-03 1.75E+04 9.60E-03 3.97E+03

1.97E-02 7.02E+03 1.97E-02 1.88E+03

2.96E-02 5.07E+03 2.96E-02 1.63E+03

4.00E-02 4.33E+03 4.00E-02 1.55E+03

5.04E-02 3.81E+03 5.04E-02 1.43E+03

6.00E-02 3.50E+03 6.00E-02 1.36E+03

6.97E-02 2.91E+03 6.97E-02 1.08E+03

8.01E-02 2.66E+03 8.01E-02 1.01E+03

9.00E-02 2.40E+03 9.00E-02 9.12E+02

1.01E-01 2.23E+03 1.01E-01 8.74E+02

2.00E-01 9.84E+02 2.00E-01 3.59E+02

4.02E-01 4.33E+02 4.02E-01 1.52E+02

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-10

Table A8 (Continued)

Integral LET for Interplanetary Galactic Cosmic Rays (Z=1-92)

LET LET Fluence LET LET Fluence

MeV*sqcm/mg #/sqcm/day MeV*sqcm/mg #/sqcm/day

Solar Minimum Solar Maximum

6.03E-01 2.90E+02 6.03E-01 1.10E+02

7.96E-01 2.23E+02 7.96E-01 8.84E+01

1.00E+00 1.79E+02 1.00E+00 7.22E+01

2.01E+00 3.39E+01 2.01E+00 5.88E+00

3.02E+00 1.43E+01 3.02E+00 2.03E+00

3.99E+00 7.76E+00 3.99E+00 1.02E+00

5.03E+00 4.59E+00 5.03E+00 5.81E-01

5.99E+00 3.07E+00 5.99E+00 3.80E-01

8.00E+00 1.55E+00 8.00E+00 1.90E-01

1.01E+01 9.00E-01 1.01E+01 1.10E-01

1.11E+01 7.17E-01 1.11E+01 8.75E-02

1.20E+01 5.76E-01 1.20E+01 7.04E-02

1.30E+01 4.67E-01 1.30E+01 5.71E-02

1.40E+01 3.85E-01 1.40E+01 4.72E-02

1.50E+01 3.16E-01 1.50E+01 3.88E-02

1.60E+01 2.61E-01 1.60E+01 3.20E-02

1.70E+01 2.20E-01 1.70E+01 2.71E-02

1.80E+01 1.85E-01 1.80E+01 2.27E-02

1.91E+01 1.54E-01 1.91E+01 1.89E-02

2.00E+01 1.30E-01 2.00E+01 1.60E-02

2.49E+01 4.45E-02 2.49E+01 5.50E-03

3.00E+01 6.27E-04 3.00E+01 8.18E-05

3.49E+01 6.86E-05 3.49E+01 1.06E-05

4.01E+01 4.18E-05 4.01E+01 6.50E-06

4.50E+01 2.83E-05 4.50E+01 4.42E-06

5.00E+01 2.00E-05 5.00E+01 3.13E-06

5.06E+01 1.92E-05 5.06E+01 3.00E-06

5.55E+01 1.34E-05 5.55E+01 2.11E-06

6.02E+01 9.38E-06 6.02E+01 1.49E-06

6.53E+01 6.32E-06 6.53E+01 1.01E-06

7.00E+01 4.40E-06 7.00E+01 7.01E-07

7.50E+01 2.83E-06 7.50E+01 4.52E-07

8.04E+01 1.65E-06 8.04E+01 2.63E-07

8.52E+01 7.71E-07 8.52E+01 1.23E-07

9.03E+01 1.94E-07 9.03E+01 3.10E-08

9.57E+01 2.88E-08 9.57E+01 4.60E-09

1.00E+02 1.19E-08 1.00E+02 1.89E-09

1.01E+02 5.27E-09 1.01E+02 8.41E-10

1.03E+02 2.54E-09 1.03E+02 4.05E-10

Effective Date: August 14, 2006 417-R-RPT-0027

Responsible Organization: GOES-R/Code 417 Baseline Version

A-11

Table A9

Integral LET for the October 1989 Solar Particle Event (Z=1-92)

LET LET Fluence LET Fluence LET…