S/C in Heliosynchronous Orbit - Spacecraft Environment Analysis

S/C in Heliosynchronous Orbit
Spacecraft Environment Analysis
Preliminary Design Review
Coll Ortega, Jordi
Molas Roca, Pau
9th of March 2018, Kiruna
Lule˚a University of Technology
The present document contains an extended study of the general hazards a spacecraft
would face in a heliosynchronous orbit. Particularly, the radiation environment is deeply
characterized. The main emphasis is made on the eﬀects of radiation on two sensitive
devices projected to be on-board.
Nomenclature
CMOS Complementary Metal Oxide Semiconductor
DD Displacement Damage
EM Electromagnetic
ESA European Space Agency
EUV Extrem-Ultra-Violet radiation
eV Electron Volt → 1.602 · 1019
C
GCRs Galactic Cosmic Rays
LET Linear Energy Transfer ( Stopping power)
LEO Low Earth Orbit, range of altitude 100km - 1000km
PEO Polar Earth Orbit (High inclined LEO orbit)
S/C Spacecraft
SAA South Atlantic Anomaly
SEL Single Event Latch-up
SEU Single Event Upset
SPENVIS SPace ENVironment Information System
TTL Transistor Transistor Logic
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S/C in Heliosynchronous Orbit - PDR

I. Introduction
A. Background
The space environment interacts with a spacecraft in a way that it can strongly affect its operation and
lifetime. Therefore, a clear understanding of this environment is essential during the design phase of a
spacecraft in order to avoid as much problems as possible in a situation where maintenance and repairing or
upgrading of damaged components is usually not possible.
The present document is focused on the effects of the radiation environment and how the particles interact
with the spacecraft, its electronic devices and its solar cells. It must be considered that spacecrafts, and also
astronauts, are exposed to a large level of radiation when in space and a risk analysis relative to extended
radiation exposures in missions and space travel in general is basic to achieve success in such inhospitable
environment.
However, a brief overview of the neutral, plasma and particulate environments will be given to provide a
better understanding of the critical phases of the whole mission and the worst scenarios that the spacecraft
will face during its operation.
B. Content
• Characterization the space radiation environment of the mission.
• Estimation worst possible cases.
• Identification of critical phases of the mission.
• Study of the radiation environment effects on an electronic device of the satellite.
• Study of the performance degradation of solar arrays.
C. Mission definition
The chosen mission for the current study is based on a spacecraft orbiting the Earth in an heliosynchronous
orbit at 800 km of altitude. In an heliosynchronous orbit, also known as Sun-synchronous orbit, the spacecraft
passes over a given latitude of the Earths surface at the same local mean solar time.
This kind of orbit is quite similar to a polar orbit regarding to its inclination respect the Earth’s equator.
However, it is normally used for very specific applications, such as imaging, spy and weather satellites,
because of its constant surface illumination angle at each location.
For instance, several missions have been using this orbit in previous missions, taking advantage of a
specific orbit case called down/dusk orbit, which allows a nearly continuous view of the Sun. Among them it
can be mentioned Yohkoh, TRACE, Hinode and PROBA2, that were mean to be solar-observing scientific
satellites.
The mission is meant to start the 1st of January of 2018 at midnight where Equator and Prime Meridian
intersect: 0◦
Latitude and 0◦
Longitude. Such an orbit is another special case, also known as a noon/midnight
orbit.
Table 1: Mission definition summary.
Orbit type Heliosynchronous
Mission start 01/01/2018 at 00:00:00 hours
Mission duration 1 year
Altitude 800 km
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D. Orbitography
In Table 2, the main characteristics of the orbit are defined. Among them, the Semi Major Axis value,
which places the spacecraft in a Low Earth Orbit environment, and the Eccentricity, which defines a perfect
circular orbit, are of great relevance.
Table 2: Orbital parameters.
Ω: Right Ascension of Ascending Node (RAAN) 100.21 ◦
a: Semi Major Axis 7178.16 km
e: Eccentricity 0
i: Inclination 98.6 ◦
ω: Argument of Perigee 0 ◦
ν0: True Anomaly 0 ◦
t0: Epoch 2018
The other key value, and the most characteristic in this case, is the Orbital Inclination: the angle
between the orbital plane and the Earth’s equator. Values larger than 90◦
, associated with Polar orbits,
gives to the spacecraft a retrograde motion around the Earth. In particular, values about 98◦
and 99◦
places
the spacecraft in an orbital plane that moves at the same rate as the Earth orbits around the Sun. The latter
means that the spacecraft will orbit with a constant Sun angle in the aforesaid Heliosynchronous orbit.
A simulation for the ground track of the spacecraft after 6 swipes of the defined orbit can be seen in
Figure 1. Another characteristic of this orbit is that it doesn’t pass through the same location every orbit.
This fact, together with its almost polar inclination, means that this orbit allows to cover most of the Earth’s
surface.
Figure 1: Ground track of the satellite after 6 orbits around the Earth.
According to the previously mentioned parameters, the spacecraft will have an Orbital Period of 1.68
hours, which means that will perform about 14.25 orbits per day around the Earth.
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(a) Altitude of the spacecraft at each instant of time during
one orbit.
(b) Orbital velocity of the spacecraft at each altitude.
Figure 2: Orbital motion characteristics of the spacecraft during the Heliosynchronous orbit.
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II. Overview of Spacecraft Environment
During the whole mission, the spacecraft will exclusively fly in LEO with a circular orbit. The latter
means that the mission will consist in only two phases: 1) Launch and 2) Orbit at 800km. Actually, most
of the lifetime of the spacecraft will occur during the second phase, which, since the launch phase will not
last long, the main study effort shall be focused on the LEO environment.
Among the primary physical components relative to space environment in Low Earth Orbits, it is possible
to expect a dense and supersonic neutral atmosphere, a cold, dense and ionospheric plasma environment,
solar ultra-violet radiation, the effects of the South Atlantic Anomaly, and a significant density of orbital
debris.
A. The Neutral Atmosphere
The neutral environment in LEO is produced by the upper layers of the Earths atmosphere, which are
mainly composed by mono-atomic oxygen over most of the altitude range. In Figure 3, the distribution of
the mono-atomic oxygen at 800km of altitude can be seen.
Figure 3: Global distribution of atomic oxygen density at 800km of altitude.
When the spacecraft passes through this environment, a drag force is produced due to the impact of the
atmospheric particles on the vehicle surfaces. As a result, some torques appear, which must be considered by
the attitude control system. Eventually, this drag can slow down the spacecraft. Therefore, small thrusters
are required to maintain the orbit during the mission time.
On the other hand, the impact of the ambient particles can also produce physical and chemical degradation
to the spacecraft surface. Figure 4a shows the density of neutral particles over the altitude. Even if these
are usually not enough energetic to remove material or erode the spacecraft surfaces, oxidation could change
the thermal properties of the external layers.
In addition to chemical effects, neutral particles also contribute to the appearance of diffuse glows which
have been detected above surfaces oriented towards the spacecraft ram direction. In Figure 4b, the density
of ion flux in that direction at different LEO altitudes is shown.
Among other possible effects that can result in neutral particles interactions with the spacecraft subsys-
tems, are changes in cover-glass transmittance in the power system, and coat and contamination of sensors
and surfaces.
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(a) Neutral particle density. (b) Ion ram particle flux.
Figure 4: Main neutral environment characteristics at 800km.
B. The Plasma Environment
A spacecraft is affected by plasma environment in any orbit. At LEO altitudes, it is cold and dense, but
the mean energy is lower than in GEO. This environment is generated by ionization of the neutral gas by
EUV and X-ray radiation, and by hyper-velocity impacts with the spacecraft surfaces or rests from plasma
thrusters and arc discharges.
Plasma is a collection of charged particles that responds to magnetic-field variations. Therefore, the
spacecraft interaction with the plasma environment is strongly dependent on its location respect to Van
Allen Belts and the Earths magnetosphere. Figure 5 shows the intensity of the Earth magnetic field at
800km of altitude.
Figure 5: Global magnetic field intensity at 800km.
The orbit performed by the spacecraft will be located in the ionosphere, below the plasmasphere. At
these altitudes, the electron concentration at different altitudes can be seen in Figure 6a. Next to it, in
Figure 6b, it can be seen how the electron temperature increase with the altitude. Usually, this tends to be
a factor of two greater than the neutral particles.
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(a) Electron density relative to the altitude.
(b) Electron temperature relative to the altitude.
Figure 6: Electron distribution characteristics at the ionosphere.
Some of the main effects of the interactions between a spacecraft and this plasma environment are
shift ground, attraction of contaminants, arc damage, changes in the surface properties, EMI, and sensor
interferences. These could seriously damage the vehicle or decrease the quality of the scientific payload
measurements.
In addition to the near-Earth effects related to a LEO environment, in a Sun-Synchronous orbit it must
be considered that with a high orbital inclination, the spacecraft crosses through the Aurora regions and
the polar cap. In direct relationship to these zones there is a high-energy plasma component that produces
significant ionization which increases the density of the thermal component. These corpuscular events are
not constant and directly dependant on the magnetic activity, usually measured by Kp.
C. The Radiation Environment
The radiation environment is generated by two main sources, electromagnetic and corpuscular radiations.
The first one includes the ambient solar photon flux, electromagnetic waves from the plasma environment,
and the electromagnetic interference resultant of the spacecraft systems or arcing. On the other hand, the
flux of particles such as electrons, protons, heavy ions and neutrons, belong to the corpuscular radiation.
For a better understanding of this environment, a detailed study of different radiation sources is required.
It should be mentioned that unlike the solar radiation, which is quite dependent on time and also on the
current solar cycle, the trapped radiation and the galactic cosmic radiation are quite constant sources with
changes on long timescales.
1. Influence of the Sun
The Sun, through the emission of electromagnetic flux and charged particles, is the main energy source for
space environment in the solar system. Although Earth’s magnetosphere provides good shielding against
the charged-particle environment, the electromagnetic flux can penetrate to the atmosphere and even to the
surface at certain wavelengths.
Regarding LEO orbits, the solar activity plays a role in short-term variations through solar flares and
geomagnetic storms. The resulting energetic particles, coupled with changes in solar Extreme-Ultra-Violet
flux that heat the atmosphere, occur mainly during solar maximum and last from a few minutes to a few
hours. However, according to the 11-years solar cycles, 2018 will occur during a solar minimum activity
period, which means that this effects will be much less critical.
During a solar minimum period, the electron density is lower because of the drop in UV/EUV fluxes. On
the contrary, the proton density increases with the galactic cosmic radiation, which should be considered,
specially near the South Atlantic Anomaly.
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2. Trapped radiation
Trapped radiation is based mainly in energetic protons and electrons, with some lower quantities of heavy
ions such as atomic oxygen, which are stuck inside the toroidal Van Allen Belts around the Earth. Figures
7 and 8 show the distribution of protons and electrons at 800 km of altitude.
(a) Trapped electron flux. (b) Average spectra of trapped electrons.
Figure 7: Trapped proton simulation at 800km for Solar Minimum.
(a) Trapped proton flux. (b) Average spectra of trapped protons.
Figure 8: Trapped electron simulation at 800km for Solar Minimum.
In particular, it is in the inner belt, which goes from hundreds of km to 6000km, where the spacecraft
will operate. In there, a special concern for low orbits is the South Atlantic Anomaly. As it can be seen in
the aforesaid Figures 7 and 8, and also in the Figure 5 regarding the intensity magnetic field of the Earth,
the SAA is a region where the Van Allen belts are weaker and comes closer to the Earths surface. This
becomes into higher fluxes of energetic particles in that zone which can produce unexpected failures if not
considered properly.
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3. Galactic Cosmic Rays
Galactic cosmic radiation is principally composed by interplanetary protons and ionized heavy nuclei. Elec-
trons are also part of the GCRs, but their intensities are much lower than that of the protons and, therefore,
are usually ignored. In Figure 9 the GCRs spectrum for some heavy ions such as Helium and atomic Oxygen
are plotted.
For LEO altitudes, the Earths magnetic field shield against many of the low-energy particles. However,
in the polar regions, all particles can enter almost parallel to the magnetic field. As a result, higher and
more directional flux can be expected to reach the spacecraft, as well as higher variations in the energy
distribution.
(a) GCR ion spectrum for Helium. (b) GCR ion spectrum for atomic Oxygen.
Figure 9: Galactic Cosmic Rays Spectrum for some heavy ions.
4. Effects of radiation in the spacecraft devices
Both corpuscular and photon radiations can produce several damages to the spacecraft devices through
many different ways, and it can be either temporary or permanent. Temporary damage happens when a
high-energy particle is introduced into an electrical component and modify its state, which is known as single-
event effect. On the other hand, sometimes, permanent damage can result from burnout of the integrated
circuit.
This effects are unleash by dielectric charging, radiofrequency interferences, thermal balance alteration,
dust and light glinting over the spacecraft surfaces, and so on.
9 of 19

D. The Particulate Environment
The particulate environment is composed by meteoroids, orbital debris and spacecraft components released
during or after their operation. All these together create an hostile environment with impact source particles
ranging from dust size to bigger elements which could damage or totally destroy a spacecraft. In Figure 10
it can be seen the flux of potential particles that could collide with the spacecraft at 800 km of altitude. On
the left, there is the simulation related with the Micrometeoroids and, on the right, the simulation regarding
the space debris flux.
(a) Micrometeroid flux. (b) Space debris flux.
Figure 10: Particulate flux density at 800km.
The damage that these particles can produce not only depend on its size or mass, but also to the impact
velocity. Figure 11 shows the required impact velocity to perforate an Aluminum plate with a thickness of
1 cm.
Figure 11: Critical particle dimensions dependent on the collision velocity.
Besides all these destructive effects, very small particulates can be trapped near the vehicle and degrade
the performance of spaceborne optical systems, damaging sensitive sensors or appearing as clutter in the
FOV of the instruments.
10 of 19

E. Critical phases and worst scenarios
Concerning the study of the different environments through which the spacecraft will have to operate during
the mission time, it has been identified some critical scenarios that must be considered specifically during
the design phase of the probe.
1. Most critical environments
Due to the proximity of the orbit to the Earth, the outer atmosphere still produce a strong influence to a
spacecraft that orbit at such high velocities.
On the other hand, because of the Heliosynchronous orbit has a very high inclination, the radiation
environment becomes into a significant interaction in certain regions such as the polar caps.
Therefore, the most critical envrionments to consider are the Neutral Atmosphere, and the Radiation
Environment.
2. Most critical areas
Strongly related with the previous section, two main areas have been identified as potentially problematic
in terms of radiation exposure, due to the geometry of the Van Allen radiation belts.
Thus, the most critical areas in which the spacecraft will have to operate every few orbits or in all the
orbits are the South Atlantic Anomaly region, and the Auroral region.
3. Most critical periods
Regarding the short-term solar particle fluxes, it has been simulated the worst periods of solar activity. In
Figure 12, it can be seen the energy levels for protons and ions.
(a) Solar protons. (b) Solar ions.
Figure 12: Worst week in short-term solar particle fluxes.
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III. Life time and Performance Degradation
Several critical components of the spacecraft orbiting in an heliosynchronous orbit may be considerably
affected by the previously presented environment. Hence, numerical simulations have been carried in order
to determine the danger they would face.
Solar arrays performance degradation is analyzed as well as four different kinds of memory devices. The
research shall lead to the selection of the most suitable device configuration and the specific shielding required
for the mission.
This section deals with protective measurements for the solar arrays and the memory device that are
on-board the spacecraft.
A. Environmental Flux
The radiation environment is tightly linked with the solar activity. The properties of plasma particle at the
orbiting altitude have been obtain through simulation in SPENVIS. Since the altitude has little variations
the values remain constant. The relevant ones are presented in Table 3.
Table 3: Properties at 800km altitude.
Neutral particle density 8.01012
m−3
Electron density 2.01011
m−3
Electron Temperature 0.259eV
Electron Thermal Velocity 3.015105
ms−1
Ion Temperature 0.198eV
Ion Thermal Velocity 1715ms−1
Average ion mass 2.1610−26
kg
B. Solar Arrays
The aim is to devise the shielding for the solar panels with the constrain of payload power delivery of 95% of
the initial power Pmax. NIEL based damage equivalent fluences for solar cells (MC-SCREAM) in SPENVIS.
The cell type Azur 3G28 used in the mission was considered. Several shielding thicknesses were tested
and checked with the calculated power degradation. After some iterations, the required thickness of the
protective layer to withstand the radiation effect shall be
x = 7.0µm (1)
The above value reduces the electron fluence to the desired limit. Once the surface of the solar panels is
determined, the mass of the shielding could be easily calculated and decide if it is feasible or not.
12 of 19

C. Total Dose and Shielding
The semi-conductor memory device is made mainly out of silicon. The device withstands 25krad of radiation
before it ceases functioning. It is mounted on the front face of the spacecraft in a box with shielding of 1.0mm.
The radiation dose received during the mission have been computed and are presented in Figure 13.
The proposed shielding thickness will receive a total dose of 2.42krad during the one year mission. The
latter conﬁrms the eligibility of the projected thickness. Nonetheless, if mass had to be reduced, the shielding
could go down to 0.3mm and still having a dose lower than 25krad.
Figure 13: Radiation dose levels during the one-year mission. Credit: SPENVIS.
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D. Single Event Upsets
A single event upset (SEU) is a change of state in a semiconductor device caused by an intruding high-energy
particle. The SEU rate for several possible choices of on-board volatile memory, such as NMOS2164, CMOS
R160-25, Bipolar 93L422, and SMJ329C50GFAM66, has been analyzed.
Simulation outputs of Bipolar 93L422 and NMOS2164, obtained in SPENVIS, are presented in Figures
14 and 15.
Figure 14: Short-term of SEU rates and LET spectra Bipolar for the 93L422 1K SRAM.
Figure 15: Short-term of SEU rates and LET spectra for the NMOS 2164.
Regarding SMJ329C50GFAM66, in order to obtain the data required for the comparative study, numerical
studies were carried and are presented next.
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1. Linear Energy Transfer Spectrum
To estimate the SEU rate, shielding thickness of 1g/cm2
(Aluminum equivalent) is assumed. Maximum
ion range from hydrogen to uranium is considered and the simulation is carried out under peak composite
worst-case flare flux and worst-case composition, taking trapped protons into account. Having specified
this environment, SPENVIS provides the total mission differential flux f(L) for values of the linear energy
transfer within a broad range.
2. Cross-Section and Components characteristics
The influence of the LET value on the SEU rate has been modeled as a Weibull distribution function. The
sensitivity function is assumed as
L →



0 L ≤ 0
Cs(1 − exp[−(L−Lo
W )s
]) L > 0
(2)
The four parameters Lo, Cs, W and s depend on the specific material. The values corresponding to all
the devices are stated in Table 4.
Sensitivity of SMJ329C50GFAM66
Numerical calculations were conducted in order to obtain the constant values characterizing the memory
device, see Table 4. To do so, a least-square fit to the data points was employed.
In Figure 16 below, both the data points and the function σ(L) are ploted. The MATLAB code has been
developed to perform the least-square fit can be found in the second code in Appendix A.
Figure 16: Approximation result for the sensitivity of the SMJ329C50GFAM66. The red dots represent the
data derived from Table 17 in Appendix B.
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3. SEU Estimation
The SEU rate ρ is the value of the following integral. Recall that the sensitivity function σ(L) stated in
equation 2 is determined by the four parameters Lo, Cs, W and s that depend on the specific material.
ρ =
∞
0
f(L)σ(L)dL (3)
Using numerical integration, a good approximation of the SEU rate ρ as defined in equation 3 for all four
devices was calculated, see Table 4.
Table 4: Parameters and SEU rate estimation.
Lo[MeV cm2
mg ] Cs[cm2
] W[MeV cm2
mg ] s SEU rate
NMOS 2164 0.487 1.71 · 10−5
4.95 1.422 1.152 · 10−4
CMOS R160-25 136.8 1.2 · 10−5
350 3.0 6.52 · 10−8
Bipolar 93L422 0.6 2.6 · 10−5
4.4 0.7 4.811 · 10−4
SMJ329C50GFAM66 1.9604 4.9942 · 10−6
1.1065 2.6050 1.1789 · 10−4
Radiation causes the least number of single event upsets in the CMOS R160-25 device, which is therefore
recommended to incorporate in the design of the satellite. The SEU rates ρ are yielded using the last
MATLAB code in Appendix A. The numerical approximation of the differential flux f(L) provided by
SPENVIS is not included.
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IV. Conclusions
The analysis made should provide the reader with detailed knowledge about the environment the mission
would be involved in. The description has been lately supported with several numerical calculations to
determine the validity of the shielding designed to fly as well as to help on the decision making of the
memory device to be flown.
The results do not give any reason to doubt about the final choice. Regarding the memory device
selection, it has been proven that the radiation causes the least number of single event upsets in the CMOS
R160-25 device. Therefore, in this PDR, CMOS R160-25 device is the one proposed as the one to be chosen.
In shielding terms, it was proven that the shielding thickness is more than enough to withstand all the
radiation interactions during the hole mission. If mass was a constraint for the mission, the shielding could
be reduced down to 0.3mm
Moving to solar array shielding, the environment is not as harmful as initially thought on the first design
phase. Hence, the shielding thickness should be of a minimum of 7µm.
The Preliminary design anlysis concludes that the performance degradation of the studied components
is compatible with the mission duration and its requirements.
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Appendix
A. Numerical Code
The code attached was implemented in Matlab.
% Weibull d i s t r i b u t i o n function
function y=weibull (x , xdata )
y=x(1)∗(1 − exp( −(( xdata−x (4))/ x ( 2 ) ) . ˆ x ( 3 ) ) ) ;
———————————
% S e n s i t i v i t y of the SMJ329C50GFAM66
data=[
% MeV LET Cross Section=NumberofFlips /(F∗ t ∗60)
0.6 2.65 1.067 e−9 %C
0.72 3.05 9.996 e−7 %C
9.6 4.537 3.0 e−6 %C
4.8 7.02 4.703 e−6 %O
20 17.4 4.9 e−6 %Ar
56 27.4 4.9985 e−6 %Fe
84 36.5 5.0 e−6 %Kr
786 55.7 5.012 e −6]; %Xe
x=abs ( l s q c u r v e f i t ( @weibull , [ 3 . 0 e−4 5 1 1 ] , data ( : , 2 ) , data ( : , 3 ) , zeros (2) , ones (2)∗5 , optimset ( ’ TolF
c l f r e s e t ;
xdata=l i n s p a c e (x ( 2 ) , 6 0 , 1 00 ) ;
hold on
plot ( data ( : , 2 ) , data ( : , 3 ) , ’ r ∗ ’ ) ;
plot ( xdata , weibull (x , xdata ) ) ;
xlabel ( ’LET [MeV cmˆ2/mg] ’ )
ylabel ( ’ sigma (LET) [cmˆ 2 ] ’ )
grid on
legend ( ’ Data ’ , ’ Weibull ’ )
hold o f f
———————————
% Numerical i n t e g r a t i o n of the product sigma (LET)∗ f lu x (LET)
load dFlux ;
load LET;
x=LET;
z = [ ] ;
f o r n=1:4
switch n
case 1 % nmos2164
Ln=0.487; Cs=1.71e −5; W=4.95; s =1.422;
case 2 % CMOS R160−25
Ln=136.8; Cs=1.2e −5; W=350; s =3.0;
case 3 % Bipolar 93L422 1K SRAM
Ln=0.6; Cs=2.6e −5; W=4.4; s =0.7;
case 4 % SMJ329C50GFAM66
Ln=1.9604; Cs=4.9942e −6; W=1.1065; s =2.6050;
end
y=Cs∗(1−exp ( −(((x−Ln)/W) . ˆ s ) ) ) . ∗ ( x>Ln )∗0.0001;
z=[z trapz (x , y .∗ dFlux ) ] ;
end
18 of 19

B. Data
Figure 17: Mono-beam experimental results of the SMJ329C50GFAM66 component.
19 of 19

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