ISS Eclipse Determination – Anson’s Projects

Determining the eclipses a satellite will encounter is a major driving factor when designing a mission in space. Thermal and power budgets have to be made with the fact that a satellite will periodically be in the complete darkness of space where it will receive no solar radiation to power the solar panels and keep the spacecraft from freezing.

What is an Eclipse

{fig-alt=“A diagram of an eclipse. The sun is shown as a large yellow circle on the left. A smaller blue circle labeled”Body” is to the right of the sun. A spacecraft is shown above the body. The umbra and penumbra are labeled.”}

The above image is a simple representation of what an eclipse is. First, you’ll notice the Umbra is complete darkness, then the Penumbra, which is a shadow of varying darkness, and then the rest of the orbit is in full sunlight. For this example, I will be using the ISS, which has a very low orbit, so the Penumbra isn’t much of a problem. However, you can tell by looking at the diagram that higher altitude orbits would spend more time in the Penumbra.

A diagram expanding on the last figure, but with distances marked for the radii of the sun and body, and the distance between the body and spacecraft.

Body Radius’s and Position Vectors

Here is a more detailed view of the eclipse that will make it easier to explain what is going on. There are 2 Position vectors, and 2 radius that need to be known for simple eclipse determination. More advanced cases where the atmosphere of the body your orbiting can significantly affect the Umbra and Penumbra, and other bodies could also potentially block the Sun. However, we will keep it simple for this example since they have minimal effect on the ISS’s orbit. Rsun and Rbody are the radius of the Sun and Body (In this case Earth), respectively. r_sun_body is a vector from the center of the Sun to the center of the target body. For this example I will only be using one vector, but for more rigorous eclipse determination it is important to calculate the ephemeris at least once a day since it does significantly change over the course of a year. The reason that I am ignoring it at the moment is because there is currently no good way to calculate Ephemerides in Julia but the package is being worked on so I may revisit this and do a more rigorous analysis in the future. r_body_sc is a position vector from the center of the body being orbited, to the center of our spacecraft.

The Code

Imports

using Unitful
using LinearAlgebra
using SatelliteToolbox
using Plots
using Colors
theme(:ggplot2)

To get the orbit for the ISS, I used a Two-Line Element which is a data format for explaining orbits. The US Joint Space Operations Center makes these widely available, but https://live.ariss.org/tle/ makes the TLE for the ISS way more accessible (“ARISS TLE,” n.d.). The Julia Package SatelliteToolbox.jl makes it super easy to turn a TLE into an orbit that can be propagated. Simply putting the TLE in a string and using the tle string macro like below, we now have access to the information to start making our ISS orbit.

ISS = tle"""
ISS (ZARYA)
1 25544U 98067A   21103.84943184  .00000176  00000-0  11381-4 0  9990
2 25544  51.6434 300.9481 0002858 223.8443 263.8789 15.48881793278621
"""

TLE:
                      Name : ISS (ZARYA)
          Satellite number : 25544
  International designator : 98067A
        Epoch (Year / Day) : 21 / 103.84943184 (2021-04-13T20:23:10.911)
        Element set number : 999
              Eccentricity :   0.00028580
               Inclination :  51.64340000 deg
                      RAAN : 300.94810000 deg
       Argument of perigee : 223.84430000 deg
              Mean anomaly : 263.87890000 deg
           Mean motion (n) :  15.48881793 revs / day
         Revolution number : 27862
                        B* :   1.1381e-05 1 / er
                     ṅ / 2 :     1.76e-06 rev / day²
                     n̈ / 6 :            0 rev / day³

Now that we have the TLE, we can pass that into SatelliteToolbox’s orbit propagator. Before propagating the orbit, we need to have a range of time steps to pass into the propagator. The TLE gives the mean motion, n, which is the revolutions per day, so using that, we can calculate the amount of time required for one orbit, which is all that we’re worried about for this analysis. The propagator returns a tuple containing the Orbital elements, a position vector with units meters, and a velocity vector with units meters per second. For this analysis were only worried about the position vector.

ISS.mean_motion

15.48881793

orbit = Propagators.init(Val(:SGP4), ISS);
time = 0:0.1:((24/ISS.mean_motion).*60*60);
propagated = Propagators.propagate!.(orbit, time);
r = first.(propagated); # Get distance from propagator

We just need to use the radii and vectors discussed earlier to determine if the ISS is in the penumbra or umbra. This is a lot of trigonometry and vector math that I won’t bore anyone with. However, using the diagrams above and following the code in the sunlight function, you should follow what’s happening. For a rigorous discussion, check out (Vallado 1997).

function sunlight(Rbody, r_sun_body, r_body_sc)
    Rsun = 695_700u"km"

    hu = Rbody * norm(r_sun_body) / (Rsun - Rbody)

    θe = acos((r_sun_body ⋅ r_body_sc) / (norm(r_sun_body) * norm(r_body_sc)))

    θu = atan(Rbody / hu)
    du = hu * sin(θu) / sin(θe + θu)

    θp = π - atan(norm(r_sun_body) / (Rsun + Rbody))
    dp = Rbody * sin(θp) / cos(θe - θp)

    S = 1
    if (θe < π / 2) && (norm(r_body_sc) < du)
        S = 0
    end
    if (θe < π / 2) && ((du < norm(r_body_sc)) && (norm(r_body_sc) < dp))
        S = (norm(r_body_sc .|> u"km") - du) / (dp - du) |> ustrip
    end

    return S
end

sunlight (generic function with 1 method)

Then we can pass all the values we’ve gathered into the function we just made.

S = r .|> R -> sunlight(6371u"km", [0.5370, 1.2606, 0.5466] .* 1e8u"km", R .* u"m");

Plotting the Results

The sunlight function returns values from 0 to 1, 0 being complete darkness, 1 being complete sunlight, and anything between being the fraction of light being received. Again since the ISS has a very low orbit, the amount of time spent in the penumbra is almost insignificant.

Code

# Get fancy with the line color. 
light_range = range(colorant"black", stop=colorant"orange", length=101);
light_colors = [light_range[unique(round(Int, 1 + s * 100))][1] for s in S];

plot(
    LinRange(0, 24, length(S)),
    S .* 100,
    linewidth=5,
    legend=false,
    color=light_colors,
);

xlabel!("Time (hr)");
ylabel!("Sunlight (%)");
title!("ISS Sunlight Over a Day")

ISS Sunlight

Looking at the plot, the vertical transition from 0% to 100% makes it pretty clear that the time in the penumbra is limited. Still, almost counterintuitively, it also looks like the ISS gets more sunlight than it does darkness. So, using the raw sunlight data, we can calculate precisely how much time is spent in each region.

Time in Sun:

sun = length(S[S.==1]) / length(S) * 100

62.04757721886597

Time in Darkness:

umbra = length(S[S.==0]) / length(S) * 100

37.627951167918546

Time in Penumbra:

penumbra = 100 - umbra - sun

0.32447161321548634

This means that even with the ISS’s low orbit, it still gets sunlight ~62% of the time and spends almost no time in the penumbra. This would vary a few percent depending on the time of year, but in a circular orbit like the ISS, the amount of sunlight would remain pretty constant. There are other orbits like a polar orbit, lunar orbit, or highly elliptic earth orbits that can have their time in the sunlight vary widely by the time of year.

References

“ARISS TLE.” n.d. Amateur Radio on the International Space Station. https://live.ariss.org/tle/.

Vallado, David A. 1997. Fundamentals of Astrodynamics and Applications, 2nd. Ed. Edited by Wiley Larson. Dordrecht: Microcosm, Inc.

Reuse

CC BY 4.0

--- title: "ISS Eclipse Determination" description: | Calculate sunlight exposure for orbiting spacecraft like the ISS. This Julia project demonstrates how to determine eclipse times, considering umbra, penumbra, and full sunlight. Visualizations and code included. Learn about orbital mechanics and mission design considerations. date: 2021-05-01 date-modified: 2024-02-292 categories: - Julia - Astrodynamics - Code - Aerospace - Notes - Space - Math creative_commons: CC BY image: preview.png image-alt: A diagram illustrating the umbra and penumbra regions cast by a celestial body. format: html: code-tools: true code-fold: false freeze: true --- Determining the eclipses a satellite will encounter is a major driving factor when designing a mission in space. Thermal and power budgets have to be made with the fact that a satellite will periodically be in the complete darkness of space where it will receive no solar radiation to power the solar panels and keep the spacecraft from freezing. ## What is an Eclipse ![Geometry of an Eclipse](geometry.svg){fig-alt="A diagram of an eclipse. The sun is shown as a large yellow circle on the left. A smaller blue circle labeled "Body" is to the right of the sun. A spacecraft is shown above the body. The umbra and penumbra are labeled."} The above image is a simple representation of what an eclipse is. First, you'll notice the Umbra is complete darkness, then the Penumbra, which is a shadow of varying darkness, and then the rest of the orbit is in full sunlight. For this example, I will be using the ISS, which has a very low orbit, so the Penumbra isn't much of a problem. However, you can tell by looking at the diagram that higher altitude orbits would spend more time in the Penumbra. ![Body Radius's and Position Vectors](vectors_radiuss.svg){fig-alt="A diagram expanding on the last figure, but with distances marked for the radii of the sun and body, and the distance between the body and spacecraft."} Here is a more detailed view of the eclipse that will make it easier to explain what is going on. There are 2 Position vectors, and 2 radius that need to be known for simple eclipse determination. More advanced cases where the atmosphere of the body your orbiting can significantly affect the Umbra and Penumbra, and other bodies could also potentially block the Sun. However, we will keep it simple for this example since they have minimal effect on the ISS’s orbit. <code style="color:#0b7285">Rsun</code> and <code style="color:#c92a2a">Rbody</code> are the radius of the Sun and Body (In this case Earth), respectively. <code style="color:#5f3dc4">r_sun_body</code> is a vector from the center of the Sun to the center of the target body. For this example I will only be using one vector, but for more rigorous eclipse determination it is important to calculate the ephemeris at least once a day since it does significantly change over the course of a year. The reason that I am ignoring it at the moment is because there is currently no good way to calculate [Ephemerides](https://ssd.jpl.nasa.gov/?ephemerides) in Julia but the package is being worked on so I may revisit this and do a more rigorous analysis in the future. <code style="color:#5c940d">r_body_sc</code> is a position vector from the center of the body being orbited, to the center of our spacecraft. ## The Code ```{julia imports, code_folding=TRUE} #| code-fold: true #| code-summary: Imports using Unitful using LinearAlgebra using SatelliteToolbox using Plots using Colors theme(:ggplot2) ``` To get the orbit for the ISS, I used a [Two-Line Element](https://en.wikipedia.org/wiki/Two-line_element_set) which is a data format for explaining orbits. The US Joint Space Operations Center makes these widely available, but https://live.ariss.org/tle/ makes the TLE for the ISS way more accessible [@ariss]. The Julia Package [SatelliteToolbox.jl](https://github.com/JuliaSpace/SatelliteToolbox.jl) makes it super easy to turn a TLE into an orbit that can be propagated. Simply putting the TLE in a string and using the `tle` string macro like below, we now have access to the information to start making our ISS orbit. ```{julia ISS, results='show'} ISS = tle""" ISS (ZARYA) 1 25544U 98067A 21103.84943184 .00000176 00000-0 11381-4 0 9990 2 25544 51.6434 300.9481 0002858 223.8443 263.8789 15.48881793278621 """ ``` Now that we have the TLE, we can pass that into SatelliteToolbox's orbit propagator. Before propagating the orbit, we need to have a range of time steps to pass into the propagator. The TLE gives the mean motion, n, which is the revolutions per day, so using that, we can calculate the amount of time required for one orbit, which is all that we're worried about for this analysis. The propagator returns a tuple containing the Orbital elements, a position vector with units meters, and a velocity vector with units meters per second. For this analysis were only worried about the position vector. ```{julia mean-motion, result='show'} ISS.mean_motion ``` ```{julia orbit-propagation} orbit = Propagators.init(Val(:SGP4), ISS); time = 0:0.1:((24/ISS.mean_motion).*60*60); propagated = Propagators.propagate!.(orbit, time); r = first.(propagated); # Get distance from propagator ``` We just need to use the radii and vectors discussed earlier to determine if the ISS is in the penumbra or umbra. This is a lot of trigonometry and vector math that I won't bore anyone with. However, using the diagrams above and following the code in the sunlight function, you should follow what's happening. For a rigorous discussion, check out [@vallado]. ```{julia sunlight-function} function sunlight(Rbody, r_sun_body, r_body_sc) Rsun = 695_700u"km" hu = Rbody * norm(r_sun_body) / (Rsun - Rbody) θe = acos((r_sun_body ⋅ r_body_sc) / (norm(r_sun_body) * norm(r_body_sc))) θu = atan(Rbody / hu) du = hu * sin(θu) / sin(θe + θu) θp = π - atan(norm(r_sun_body) / (Rsun + Rbody)) dp = Rbody * sin(θp) / cos(θe - θp) S = 1 if (θe < π / 2) && (norm(r_body_sc) < du) S = 0 end if (θe < π / 2) && ((du < norm(r_body_sc)) && (norm(r_body_sc) < dp)) S = (norm(r_body_sc .|> u"km") - du) / (dp - du) |> ustrip end return S end ``` Then we can pass all the values we've gathered into the function we just made. ```{julia} S = r .|> R -> sunlight(6371u"km", [0.5370, 1.2606, 0.5466] .* 1e8u"km", R .* u"m"); ``` ## Plotting the Results The `sunlight` function returns values from 0 to 1, 0 being complete darkness, 1 being complete sunlight, and anything between being the fraction of light being received. Again since the ISS has a very low orbit, the amount of time spent in the penumbra is almost insignificant. ```{julia} #| code-fold: true #| fig-cap: ISS Sunlight #| alt-text: A graph titled "ISS Sunlight Over a Day" showing the percentage of sunlight the ISS receives over a 24-hour period. The x-axis represents time in hours, and the y-axis represents sunlight percentage. The graph shows a period of near-total sunlight followed by a period of darkness, and the cycle repeats. # Get fancy with the line color. light_range = range(colorant"black", stop=colorant"orange", length=101); light_colors = [light_range[unique(round(Int, 1 + s * 100))][1] for s in S]; plot( LinRange(0, 24, length(S)), S .* 100, linewidth=5, legend=false, color=light_colors, ); xlabel!("Time (hr)"); ylabel!("Sunlight (%)"); title!("ISS Sunlight Over a Day") ``` Looking at the plot, the vertical transition from 0% to 100% makes it pretty clear that the time in the penumbra is limited. Still, almost counterintuitively, it also looks like the ISS gets more sunlight than it does darkness. So, using the raw sunlight data, we can calculate precisely how much time is spent in each region. Time in Sun: ```{julia sun-time, results='show'} sun = length(S[S.==1]) / length(S) * 100 ``` Time in Darkness: ```{julia dark-time, results='show'} umbra = length(S[S.==0]) / length(S) * 100 ``` Time in Penumbra: ```{julia penum-time, results='show'} penumbra = 100 - umbra - sun ``` This means that even with the ISS's low orbit, it still gets sunlight ~62% of the time and spends almost no time in the penumbra. This would vary a few percent depending on the time of year, but in a circular orbit like the ISS, the amount of sunlight would remain pretty constant. There are other orbits like a polar orbit, lunar orbit, or highly elliptic earth orbits that can have their time in the sunlight vary widely by the time of year.