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NEAR of the 21 Lunar Landings, 19—All of the U.S

NEAR of the 21 Lunar Landings, 19—All of the U.S

Copyrights Prof Marko Popovic 2021

NEAR Of the 21 lunar landings, 19—all of the U.S. and Russian landings—occurred between 1966 and 1976. Then humanity took a 37-year break from landing on the moon before China achieved its first lunar touchdown in 2013. Take a look at the first 21 successful lunar landings on this interactive map https://www.smithsonianmag.com/science-nature/interactive-map-shows-all-21-successful-moon-landings-180972687/ Moon 1

The near side of the Moon with the major maria (singular mare, vocalized mar-ray) and lunar craters identified. Maria means "seas" in Latin. The maria are basaltic lava plains: i.e., the frozen seas of lava from lava flows. The maria cover ∼ 16% (30%) of the lunar surface (near side).

Light areas are Lunar Highlands exhibiting more impact craters than Maria.

The far side is pocked by ancient craters, mountains and rugged terrain, largely devoid of the smooth maria we see on the near side. The Lunar Reconnaissance Moon 2 is a NASA robotic spacecraft currently orbiting the Moon in an eccentric polar mapping . LRO data is essential for planning NASA's future human and robotic missions to the Moon. Launch date: June 18, 2009 : 2 hours Orbit height: 31 mi Speed on orbit: 0.9942 miles/s Cost: 504 million USD (2009) The Moon is covered with a gently rolling layer of powdery soil with scattered rocks called the regolith; it is made from debris blasted out of the Lunar craters by the meteor impacts that created them. Lunar soil is composed of grains 1 cm in diameter or less. Lunar dust are particles less than 10-50 μm in diameter.

The dust is electrically charged and sticks to any surface with which it comes in contact. Density of lunar regolith is about 1.5 g/cm3. Dirt becomes very dense beneath the top layer of regolith.

The major processes involved in the formation of lunar soil are: Comminution: mechanical breaking of rocks and minerals into smaller particles by meteorite and micrometeorite impacts; Agglutination: welding of mineral and rock fragments together by micrometeorite-impact-produced glass; Solar wind sputtering and cosmic ray spallation caused by impacts of ions and high energy particles. These processes constitute space weathering.

Lunar rocks are igneous ("fire-formed rocks"); rich in calcium (Ca), Aluminum (Al), and Titanium (Ti), poor in light elements like hydrogen (H), and with high abundance of elements like Silicon (Si) and (O).

The Maria are mostly composed of dark basalts. The Highlands rocks are largely Anorthosite. Why do we always observe the same side of the moon? Imagine is fixed in space and is orbiting in … 푀 푀 3 2 푅1 < 퐺 2 푅1휔 < 퐺 2 휔 푅1 푚1 푅 < 푅 1 0 3 푀 2 푀 푅0 = 퐺 2 푅0휔 = 퐺 2 휔 푅 푅0 0 푚0 푀 푅3 > 퐺 푀 2 휔2 푅2 > 푅0 푀 2 푚2 푅2휔 > 퐺 2 푅2

Satellite consists of main body with mass 푚0 and two much smaller, rigidly connected bodies with masses 푚1 and 푚2. All three bodies have the same orbital angular speed 휔 about the Earth, mass 푀. 2 푀푚0 While the main body satisfies the stable orbital equation 푚0푅0휔 = 퐺 2 the other two do not. 푅0 푀푚 After applying 2nd Newton’s law 푚푎 − 푚휔2푅 = −퐺 one obtains 푎 푟푎푑𝑖푎푙 푅2 푟푎푑𝑖푎푙

푎푟푎푑𝑖푎푙 1

푅1 < 푅0

푅0 푀 푅2 > 푅0

푎푟푎푑𝑖푎푙 2

So far we have ignored the rigid connection between three masses. And hence, quite erroneously, it appears here that separation between masses would increase. (NEXT SLIDE) 퐹Ԧ1 푅1 < 푅0 푎1

푅0

퐹Ԧ2 푀 푎2 푅2 > 푅0

The resulting 푎1 and 푎2 of small masses 푚1 and 푚2 are perpendicular to the line connecting all three masses. The component of inertial forces parallel to that line is cancelled by attractive forces pushing masses 푚1 and 푚2 toward mass 푚0 and hence keeping distances between masses constant. Eventually the 3-body Moon reaches…

Stable configuration

푀

Moon’s rotation and orbital periods are tidally locked with each other .

The far side of the Moon was not seen until 1959 when it was photographed by the Soviet spacecraft Luna 3. Now you know

푀

But, we still need to explain the terminology “tidal locking”… Gravitation induced of droplet on the surface of the Earth

푚0 퐺 2 푅0 푀 퐺 2 푚0 푅퐸 푚0 퐺 2 퐺 2 푅0 + 푅퐸 푅0 − 푅퐸 푚0 푀 푀 푀 퐺 2 퐺 2 푅퐸 푅퐸

푚0 퐺 2 푅0 Gravitation induced acceleration of water droplet on the surface of the Earth

푚 Earth is actually also orbiting about the Earth-Moon (!!!) 퐺 0 2 풎ퟎ푴 ퟐ 풎ퟎ 푅0 푮 ퟐ = 푴흎 푹ퟎ 푹 푴 + 풎ퟎ 푀 ퟎ 퐺 2 푚0 푅퐸 푚0 퐺 2 퐺 2 푅0 + 푅퐸 푅0 − 푅퐸 푚0 푀 푀 푀 퐺 2 퐺 2 Please be careful here…Earth is also orbiting due to Moon’s pull. 푅퐸 푅 퐸 풎ퟎ Hence one needs to subtract 푮 ퟐ which accounts for Earth’s acceleration… 푹ퟎ 푚0 퐺 2 푅0 Gravitation induced acceleration of water droplet on the surface of the Earth

푚0 퐺 2 푅0 푀 퐺 2 푚0 푅퐸 푚0 퐺 2 퐺 2 푅0 + 푅퐸 푅0 − 푅퐸 푚0 푀 푀 푀 퐺 퐺 푅2 2 풎ퟎ 퐸 푅퐸 Lets subtract 푮 ퟐ to accounts for Earth’s acceleration… 푹ퟎ

1 1 푅퐸 2푅0 − 푅퐸 2푅퐸 푚 퐺푚0 − = 퐺푚0 ≅ 퐺푚0 0 푅 − 푅 2 푅2 푅 − 푅 2푅2 푅3 퐺 2 0 퐸 0 0 퐸 0 0 푅0 1 1 푅퐸 2푅0 + 푅퐸 2푅퐸 퐺푚0 2 − 2 = −퐺푚0 2 2 ≅ −퐺푚0 3 푅0 + 푅퐸 푅0 푅0 + 푅퐸 푅0 푅0 Gravitation induced acceleration of water droplet on the surface of the Earth and corrected by the orbiting motion of the Earth

Yes, once again, Earth is orbiting about the Earth-Moon center of mass 푀 2푅퐸 퐺 퐺푚 2 2푅 0 푅3 푅퐸 퐸 0 퐺푚0 3 푅0 푚0 푀 푀 푀 퐺 2 퐺 2 푅퐸 푅퐸 1 1 푅퐸 2푅0 − 푅퐸 2푅퐸 퐺푚0 2 − 2 = 퐺푚0 2 2 ≅ 퐺푚0 3 푅0 − 푅퐸 푅0 푅0 − 푅퐸 푅0 푅0

1 1 푅퐸 2푅0 + 푅퐸 2푅퐸 퐺푚0 2 − 2 = −퐺푚0 2 2 ≅ −퐺푚0 3 푅0 + 푅퐸 푅0 푅0 + 푅퐸 푅0 푅0 Gravitation induced acceleration of water droplet on the surface of the Earth and corrected by the orbiting motion of the Earth

𝑔Τ𝑔푀표표푛 푡𝑖푑푎푙 3 푀 2푅퐸 푀 푅0 81 3 푀 퐺 2൘퐺푚0 3 = ≈ 60 ≅ 8,748,000 2푅퐸 퐺 푅퐸 푅0 2푚0 푅퐸 2 퐺푚 2 2푅 0 푅3 푅퐸 퐸 0 퐺푚0 3 푅0 푚0 푀 푀 푀 퐺 2 퐺 2 푅퐸 푅퐸

These two bulges explain why in one day there are two high tides and two low tides, as the Earth's surface rotates through each of the bulges once a day. It is surprising that this tiny moon acceleration

𝑔푀표표푛 푡𝑖푑푎푙 = 𝑔Τ8,748,000

is able to create such a huge effect…

For example, in the Bay of Fundy, between New Brunswick and Nova Scotia, high tides are often 12 m above low tides (!)

We should have expected this result; stable configuration is ‘elongated’ along axis that connects two gravitating bodies. Tidal locking configuration Stable configuration Is there any similar effect caused by the Sun’s ?

3 푀 2푅퐸 푀 푅0 81 3 Recall 𝑔Τ𝑔푀표표푛 푡𝑖푑푎푙 = 퐺 2ൗ퐺푚0 3 = ≈ 60 ≅ 8,748,000 푅퐸 푅0 2푚0 푅퐸 2

In the case of Sun

3 푀 2푅퐸 푀 푅푆 1 3 𝑔Τ𝑔푆푢푛 푡𝑖푑푎푙 = 퐺 2൘퐺푀푆 3 = ≈ 23,455 ≅ 19,375,000 푅퐸 푅푆 2푀푆 푅퐸 2 × 333,000 or 𝑔푆푢푛 푡𝑖푑푎푙Τ𝑔푀표표푛 푡𝑖푑푎푙 ≈ 0.45

Hence, yes, Sun’s gravity also causes tides… Moon and Sun combined But that is not all…

There is an average 40 minutes delay between 푀 the Moon’s transit and the following high tide. (!) 퐺 2 푅퐸 2푅퐸 표 퐺푚0 3 10 푅0 푚0 푀 푀 푀 2푅퐸 퐺 퐺 퐺푚0 2 2 푅3 푅퐸 푅퐸 0 If this is the northern hemisphere Moon is orbiting counterclockwise.

Imagine that this is the northern hemisphere Bulge is pushed “forward” due to the Earth’s -> Earth is rotating counterclockwise. rapid rotation and viscous forces. Tidal Braking

This is tidal braking that slows down Earth’s rotational while it 푀 increases Moon’s orbital angular momentum. 퐺 2 Sum of Earth’s rotational angular momentum and 푅퐸 2푅퐸 표 Moon’s orbital angular momentum is constant. 퐺푚0 3 10 푅0 푚0 푀 푀 푀 2푅퐸 퐺 퐺 퐺푚0 2 2 푅3 Moon is orbiting counterclockwise. 푅퐸 푅퐸 0 Increased Moon’s orbital angular momentum means that the Moon’s distance must increase too.

Earth is rotating counterclockwise. Moon is receding from the Earth at a rate of ퟒ 풄풎/풚풆풂풓

And the duration of Earth’s day increases by 0.0016푠 every century. 휃

(Please note >0 outward and <0 inward)

G푀 푚Τ2 푚 푚 − = − 휔2 푅 + 0.5퐻 cos 휃 + 푎 To understand this, we will 푅 + 0.5퐻 cos 휃 2 2 2 푟푎푑𝑖푎푙 chop the into two Τ G푀 푚Τ2 halves. G푀 푚 2 퐺푀푚/2 푚 − 2 = − 3 푅 + 0.5퐻 cos 휃 + 푎푟푎푑𝑖푎푙 푅 + 0.5퐻 cos 휃 2 푅 + 0.5퐻 cos 휃 푅 2

2 푅 푚Τ2 퐺푀푚/2 1 퐻 cos 휃 푚 퐺푀푚 − − 1 + = 푎푟푎푑𝑖푎푙 H 푚휔2푅 = 푅2 퐻 cos 휃 2 2푅 2 2 1 + 푚Τ2 푅 2푅 G푀 푚Τ2 푅 − 0.5퐻 cos 휃 2 퐺푀푚/2 4퐻 cos 휃 푚 2퐺푀퐻 cos 휃 푚 − − = = 푎 푅2 2푅 2 푅3 2 푟푎푑𝑖푎푙

푚 2퐺푀퐻 cos 휃 2 푚 2퐺푀퐻 cos 휃 sin 휃 → Stretching inertial force & Rotating inertial force 2 푅3 2 푅3 휃

(Please note >0 outward and <0 inward)

G푀 푚Τ2 푚 푚 − = − 휔2 푅 − 0.5퐻 cos 휃 + 푎 To understand this, we will 푅 − 0.5퐻 cos 휃 2 2 2 푟푎푑𝑖푎푙 chop the astronaut into two Τ G푀 푚Τ2 halves. G푀 푚 2 퐺푀푚/2 푚 − 2 = − 3 푅 − 0.5퐻 cos 휃 + 푎푟푎푑𝑖푎푙 푅 + 0.5퐻 cos 휃 2 푅 − 0.5퐻 cos 휃 푅 2

2 푅 푚Τ2 퐺푀푚/2 1 퐻 cos 휃 푚 퐺푀푚 − − 1 − = 푎푟푎푑𝑖푎푙 H 푚휔2푅 = 푅2 퐻 cos 휃 2 2푅 2 2 1 − 푚Τ2 푅 2푅 G푀 푚Τ2 푅 − 0.5퐻 cos 휃 2 퐺푀푚/2 4퐻 cos 휃 푚 2퐺푀퐻 cos 휃 푚 − = − = 푎 푅2 2푅 2 푅3 2 푟푎푑𝑖푎푙

푚 2퐺푀퐻 cos 휃 2 푚 2퐺푀퐻 cos 휃 sin 휃 → Stretching inertial force & Rotating inertial force 2 푅3 2 푅3 퐻 퐹 cos 휃 2 푔 푅

퐻 퐻 퐹 cos 휃 sin 휃 푅 휃 푔 푅

퐻 퐹 cos 휃 sin 휃 푔 푅 퐺푀푚 푅 2 퐹 = = 푚𝑔 푅 = 푚 × 9.83푚/푠2 × 퐸 푔 푅2 푅 퐻 2 퐹 cos 휃 2 푔 푅 Professor, 9.83푚/푠 ? Surface gravitational acceleration

Gravitational acceleration can be obtained from Newton’s law of gravity as 푀푚 푀 4휋휌 푅 퐺 = 푚𝑔 → 퐺 = 퐺 푎푣푒 = 𝑔 푅2 푅2 3 with R radius of the , M mass of the planet, and 휌푎푣푒 average density of the planet.

Hence, by substitution of density and radius for Earth with universal gravitational constant G=6.67408 × 10-11 m3 kg-1 s-2

3 3 푘𝑔 6 − 3 4휋휌 푅 11 m 4휋 ∙ 5.52 ∙ 10 ∙ 6.37 ∙ 10 푚 풎 푎푣푒 퐸 m If Earth stood still 𝑔 = 퐺 ≈ 6.67408 ∙ 10 2 = ퟗ. ퟖퟑ ퟐ . 3 kg s 3 풔 (no spinning) This pure gravitational pull is affected by Earth daily rotations by amount 2휋 2 ∙ 6.37 ∙ 106푚 −휔2푅 푐표푠휃 = − 푐표푠휃 = −0.0337푐표푠휃 퐸 퐸 86400푠 2 With 휃 being latitude. For 휃 = 45표 2 푚 푚 𝑔 = 9.83 − 0.0337 ≈ 9.81 . 푒푓푓 2 푠2 푠2 퐻 2 −7 퐻 2 For (LEO), assume 푅 ≈ 푅퐸 and ≈ = 3.1 × 10 . 퐹 cos 휃 푅 6,371,000 푔 푅

퐻 퐻 퐹 cos 휃 sin 휃 푅 휃 푔 푅

퐻 퐹 cos 휃 sin 휃 푔 푅 2 퐺푀푚 2 푅퐸 퐹푔 = 2 = 푚𝑔 푅 = 푚 × 9.83푚/푠 × 퐻 푅 푅 퐹 cos 휃 2 푔 푅 Can we create artificial “tide” and spin big wheel in space?

Sure, why not. For example we can have 푡1 masses that move in radial directions to mimic asymmetric tidal bulges.

휔

푡2

휏Ԧ

Use green energy 푡3 Psssst! Easier way to do that is with…

푡1 Robotic arm moving in intelligent way relative to the wheel.

휔

푡2

휏Ԧ

Use green energy 푡3 From and of we obtain angular acceleration

2 푟푤ℎ푒푒푙 휏 𝑔 휏 ≈ 4푚 g Say 퐼 = 4푚 푟2 푏푎푙푙 푅 푟𝑖푚 푏푎푙푙 푤ℎ푒푒푙 → 훼 = = 퐸 퐼 푅퐸

휔

푚푏푎푙푙

푚푏푎푙푙

휏Ԧ

Time for wheel to reach 1 revolution per second. 휔 2휋 푠−1 2휋 푠−1푅 푡 = = = 퐸 ≈ 4 × 106푠 ≈ 47 days 훼 훼 𝑔 Maybe have two wheels rotating in opposite directions for stability

휔 −휔

Applications: artificial gravity, energy storage, etc. Recently in the news: “Orbital Assembly Corp. aims to have luxury space hotel open in 2027” Voyager Station: 11,600 m2 (125,000 sf) of habitable space in modules and access tubes 200m in overall diameter (ISS is 73m long and 109m wide). Estimated mass of 2,418 metric tons (ISS: 419 tons). Estimated volume of 51,104 m3 (ISS pressurized volume: 915 m3) They plan to use thrusters to turn the wheel.

What is the angular speed that will provide

Earth like gravity? Period ~ퟐퟎ풔 푚 9.81 2 푟푎푑 휔2푟 = 𝑔 → 휔 = 푠 ≈ 0.313 The Voyager Station by Orbital Assembly Corp. (1 minute, 36 seconds) 100 푚 푠 https://www.youtube.com/watch?v=Ao5o3EgQ-sU

Moon like gravity? Period ~ퟒퟗ풔 푚 1.62 2 푟푎푑 Should they worry about and stability? 휔2푟 = 𝑔 → 휔 = 푠 ≈ 0.127 100 푚 푠 The International (ISS) is the second brightest object in the night sky after the moon https://www.nasa.gov/feature/facts-and-figures The ISS's orientation is controlled through the use of Control-Moment Gyroscopes (CMG), with chemical thrusters used as backup when demands exceed CMG authority or to desaturate CMGs.

Nadir (white arrow) along line connecting ISS and center of the Earth. Image courtesy of the Earth Science and Remote Sensing Unit, NASA https://www.nasa.gov/directorates/heo/scan/services/missions/other/ISS.html The gravitational torque is zero for ISS orientation but it is not a stable equilibrium. Hence active control of orientation is necessary. Apogee altitude: 422 km (262.2 mi) Perigee altitude: 418 km (259.7 mi) Orbital decay: 2 km/month Orbital period: 92.68 minutes

Let’s check altitude (fun exercise):

2 2 퐺푀 푅퐸 휔 푅퐸 + ℎ = 2 = 𝑔′ ℎ 푅퐸 + ℎ 푅퐸 + ℎ

ℎ 3 𝑔′푇2 3 𝑔′푇2 1 + = 2 → ℎ = 푅퐸 2 − 1 푅퐸 4휋 푅퐸 4휋 푅퐸

3 9.83 92.682 602 ′ 퐺푀 푚 2휋 풉 = 6.371 × 106푚 − 1 ≅ ퟒퟏퟖ 풌풎 (used 𝑔 = 2 = 9.83 2 & 휔 = ) 2 6 푅퐸 푠 푇 4휋 6.371 × 10 Going back to Voyager Space Station and stability vs Coriolis Force. It depends on orientation of the station. If station axis is aligned with Nadir and station is orbiting If the station is within the equatorial plane and orbiting in the equatorial plane, then Coriolis Force will generate in the equatorial plane, then Coriolis Force will not torque that will tend to rotate station about the orbital generate torque. Perhaps that could work? path.

휔퐸 & 휔표푟푏𝑖푡𝑖푛푔 푠푡푎푡𝑖표푛 휔퐸 & 휔표푟푏𝑖푡𝑖푛푔 푠푡푎푡𝑖표푛

퐿푤ℎ푒푒푙 퐿푤ℎ푒푒푙

퐹Ԧ퐶표푟𝑖표푙𝑖푠 = −2푚 휔 × 푣Ԧ′ Tricky business…Maybe it is best to have two wheels next to each other and rotating in opposite directions? Geostationary orbit

Consider orbit for which are always positioned above the same point on Earth. In this case angular speed of satellites and of the Earth coincide 푣 = 휔 푅 퐸 also for any stationary orbit gravitational pull must be equal to mass times radial acceleration 푚푀 푣2 퐺 퐸 = 푚 = 푚휔2푅 . 푅2 푅 퐸 Hence 푚 2 3 9.81 ∙ 86400푠 2 3 퐺푀퐸 3 𝑔푅퐸 3 𝑔 푠2 푅 = 2 = 2 = 2 푅퐸 = 2 6 푅퐸 = 6.63 ∙ 푅퐸 휔퐸 휔퐸 휔퐸푅퐸 2휋 ∙ 6.37 ∙ 10 푚 or 푅 = 6.63 ∙ 6370 푘푚 ≈ 42200푘푚 . With speed ~ 3.1 푘푚/푠 https://upload.wikimedia.org/wikipedia/commons/b/b4/Comparison_satellite_navigation_orbits.svg

Comparison of geostationary Earth orbit with GPS, GLONASS, Galileo and Compass (medium Earth orbit) satellite navigation system with the International Space Station, Hubble and Iridium constellation orbits, and the nominal size of the Earth. The Moon's orbit is around 9 times larger (in radius and length) than geostationary orbit. ISS is moving pretty fast with speed of 7.667 km/s or 27600 km/h. Recall ℎ ≈ 420푘푚. For comparison, the ℎ ≈ 0 푘푚 satellite would need to orbit with 7.9 km/s (first cosmic speed).

Both speeds are very close to or second cosmic speed.

Escape velocity is related to necessary kinetic energy of ballistic object to leave gravitational pull, and in the context of Earth can be expressed as 1 1 푣2 퐺푚푀퐸 − = 푚 푅퐸 ∞ 2 or

2퐺푀퐸 3 푚 푘푚 3 푘푚 푣 = = 2𝑔푅퐸 = 11.2 ∙ 10 = 11.2 = 40.2 ∙ 10 . 푅퐸 푠 푠 ℎ referred to as second cosmic speed. [9] [9] Location with respect to Ve (km/s) Location with respect to Ve (km/s) on the Sun the Sun's gravity 617.5 on Mercury Mercury's gravity 4.3[10]:230 at Mercury the Sun's gravity 67.7 on Venus's gravity 10.3 at Venus the Sun's gravity 49.5 on Earth Earth's gravity 11.2[10]:200 at the Earth/Moon the Sun's gravity 42.1 on the Moon the Moon's gravity 2.4 at the Moon the Earth's gravity 1.4 on Mars' gravity 5.0[10]:234 at Mars the Sun's gravity 34.1 on Ceres Ceres's gravity 0.51 on Jupiter Jupiter's gravity 59.6[10]:236 at Jupiter the Sun's gravity 18.5 on Io Io's gravity 2.558 on Europa Europa's gravity 2.025 on Ganymede Ganymede's gravity 2.741 on Callisto Callisto's gravity 2.440 on Saturn Saturn's gravity 35.6[10]:238 at Saturn the Sun's gravity 13.6 on Titan Titan's gravity 2.639 on Uranus Uranus' gravity 21.3[10]:240 at Uranus the Sun's gravity 9.6 on Neptune Neptune's gravity 23.8[10]:240 at Neptune the Sun's gravity 7.7 on Triton's gravity 1.455 on Pluto Pluto's gravity 1.2 at Solar the Milky Way's gravity 492–594[11][12] System galactic radius on the event horizon a 's gravity 299,792 (speed of light LEO

A low Earth orbit (LEO) is an orbit around Earth with an altitude between 160 kilometers (99 mi) (orbital period of about 88 minutes), and 2,000 kilometers (1,200 mi) (about 127 minutes). Objects below approximately 160 kilometers (99 mi) will experience very rapid orbital decay and altitude loss. The orbital velocity needed to maintain a stable low Earth orbit is about 7.8 km/s, but it reduces with increased orbital altitude. With the exception of the manned lunar flights of the Apollo program, all human have taken place in LEO (or were suborbital). The International Space Station conducts operations in LEO. The altitude record for a human in LEO was with an apogee of 1,374.1 kilometers (853.8 mi). All manned space stations to date, as well as the majority of satellites, have been in LEO. LEO

Objects in LEO encounter atmospheric drag from gases in the thermosphere (approximately 80–500 km up) or exosphere (approximately 500 km and up), depending on orbit height. Due to atmospheric drag, satellites do not usually orbit below 300 km. Objects in LEO orbit Earth between the denser part of the and below the inner Van Allen radiation belt. The orbital velocity needed to maintain a stable low Earth orbit is about 7.8 km/s, but it reduces with increased orbital altitude. Calculated for circular orbit of 200 km it is 7.79 km/s and for 1500 km it is 7.12 km/s. The delta-v needed to achieve low Earth orbit starts around 9.4 km/s. Atmospheric and gravity drag associated with launch typically adds 1.3–1.8 km/s to the launch vehicle delta-v required to reach normal LEO orbital velocity of around 7.8 km/s (28,080 km/h). Equatorial low Earth orbits (ELEO) are a subset of LEO. These orbits, with low inclination to the Equator, allow rapid revisit times and have the lowest delta-v requirement (i.e., fuel spend) of any orbit. Orbits with a high inclination angle to the equator are usually called polar orbits. Satellites and LEO

Higher orbits include medium Earth orbit (MEO), sometimes called intermediate circular orbit (ICO), and further above, geostationary orbit (GEO). Orbits higher than low orbit can lead to early failure of electronic components due to intense radiation and charge accumulation. A low Earth orbit is simplest and cheapest for satellite placement. It provides high bandwidth and low communication time lag (latency), but satellites in LEO will not be visible from any given point on the Earth at all times. Earth observation satellites and spy satellites use LEO as they can see the surface of the Earth more clearly as they are not so far away. They are also able to traverse the surface of the Earth. A majority of artificial satellites are placed in LEO, making one complete revolution around the Earth in about 90 minutes. The International Space Station is in a LEO about 400 km (250 mi) above the Earth's surface. Since it requires less energy to place a satellite into a LEO and the LEO satellite needs less powerful amplifiers for successful transmission, LEO is still used for many communication applications. Because these LEO orbits are not geostationary, a network (or "constellation") of satellites is required to provide continuous coverage. Space debris encompasses both natural (meteoroid) and artificial (man-made) particles. Meteoroids are in orbit about the sun, while most artificial debris is in orbit about the Earth. Hence, the latter is more commonly referred to as orbital debris. Orbital debris is any man-made object in orbit about the Earth which no longer serves a useful function. Such debris includes nonfunctional spacecrafts, abandoned launch vehicle stages, mission-related debris and fragmentation debris. There are more than 20,000 pieces of debris larger than a softball orbiting the Earth. They travel at speeds up to 17,500 mph, fast enough for a relatively small piece of orbital debris to damage a satellite or a spacecraft. There are 500,000 pieces of debris the size of a marble or larger. There are many millions of pieces of debris that are so small they can’t be tracked. Even tiny paint flecks can damage a spacecraft when traveling at these velocities. In fact, a number of windows have been replaced because of damage caused by material that was analyzed and shown to be paint flecks. Space Debris

In 1996, a French satellite was hit and damaged by debris from a French rocket that had exploded a decade earlier. On Feb. 10, 2009, a defunct Russian satellite collided with and destroyed a functioning U.S. Iridium commercial satellite. The collision added more than 2,000 pieces of trackable debris to the inventory of space junk. China's 2007 anti-satellite test, which used a missile to destroy an old weather satellite, added more than 3,000 pieces to the debris problem. NASA and the DoD cooperate and share responsibilities for characterizing the satellite (including orbital debris) environment. DoD’s Space Surveillance Network tracks discrete objects as small as 2 inches (5 centimeters) in diameter in low Earth orbit and about 1 yard (1 meter) in geosynchronous orbit. Currently, about 15,000 officially cataloged objects are still in orbit. The total number of tracked objects exceeds 21,000. Using special ground-based sensors and inspections of returned satellite surfaces, NASA statistically determines the extent of the population for objects less than 4 inches (10 centimeters) in diameter. Collision risks are divided into three categories depending upon size of threat. For objects 4 inches (10 centimeters) and larger, conjunction assessments and collision avoidance maneuvers are effective in countering objects which can be tracked by the Space Surveillance Network. Objects smaller than this usually are too small to track and too large to shield against. Debris shields can be effective in withstanding impacts of particles smaller than half an inch (1 centimeter). RemoveDEBRIS satellite

https://www.youtube.com/watch?v=5MX-1a2VZHk

A British satellite, designed to test out ways to clean up debris in space, successfully ensnared a simulated piece of junk in orbit using a big net. In September 2018, the vehicle, known as the RemoveDEBRIS satellite, deployed its onboard net, which then captured a nearby target probe that the vehicle had released a few seconds earlier. The demonstration shows that a simple idea like a net may be an effective way to clean up all the material orbiting Earth.

The RemoveDEBRIS satellite is meant to try out numerous different methods for cleaning up space junk, which has become a growing problem ever since we started launching rockets into orbit. Thousands of dead, uncontrollable objects linger in orbit, including defunct satellites, spent launch vehicles, and other pieces of debris that have come off other spacecrafts. And all of this junk is moving fast, at upwards of 17,000 miles per hour. The more debris we have in orbit, the higher the chance that these pieces might collide at break-neck speeds, creating even more debris that could pose a threat to other spacecraft.

THE MORE DEBRIS WE HAVE IN ORBIT, THE HIGHER THE CHANCE THAT THESE PIECES MIGHT COLLIDE Lets go back to two wheels rotating in opposite directions and see if we can use them to give that extra (11.2-7.7) km/s=3.5 km/s kick WIND UP MECHANISM? Need mechanism that provide reasonable acceleration, say less than 5g to keep “comfy” (more realistic number is probably just 2g), long enough (probably elastic) cable, and simultaneous lunch of two vehicles…Hmmm

휔 −휔 푣2 35002 푠 = = ≅ 125 푘푚 2 × 5𝑔 10 × 9.81 Need better idea…

Extra vehicle, too much acrobatics, very long cable… NOTE: when you see NAAAH gray slide background, please take it with a grain of salt. Multiple kicks? … NAAAH Any other way to give that extra (11.2-7.7) km/s=3.5 km/s kick?

2 2 2 TWIRLING MECHANISM? While making sure that 푎푟푎푑𝑖푎푙 + 푎푡푎푛푔푒푛푡𝑖푎푙 ≤ 푛 𝑔 푣Ԧ푡푎푛푔푒푛푡𝑖푎푙 with say 푛 = 5. (  I know…) NOTE: when you see gray slide background, please take it with a For with 푟 = 푐표푛푠푡. and 푣2 grain of salt. 푎 = 0 one has 푇 = 푚 푡푎푛푔푒푛푡푖푎푙 ≤ 푛𝑔 푡푎푛푔푒푛푡𝑖푎푙 푟 푟 푣Ԧ푟푎푑𝑖푎푙 Here, however

2 2 푣푡푎푛푔푒푛푡𝑖푎푙 푑푣 푣푡푎푛푔푒푛푡𝑖푎푙 푇 = 푚 − 푚 푟푎푑𝑖푎푙 ≤ 푚 푟 푑푡 푟

On previous slide we obtained length of cable ≈ 125 푘푚

125푘푚 Here we could do much better and require ‘only’ ~ ≈ ퟐퟎ 풌풎 2휋

Perhaps feasible Kepler’s Laws

In the early 1600s, Johannes Kepler proposed three laws of planetary motion. Kepler was able to summarize the carefully collected data of his mentor - Tycho Brahe - with three statements that described the motion of in a sun-centered . Kepler's efforts to explain the underlying reasons for such motions are no longer accepted; nonetheless, the actual laws themselves are still considered an accurate description of the motion of any planet and any satellite. Kepler's three laws of planetary motion can be described as follows:

 The path of the planets about the sun is elliptical in shape, with the center of the sun being located at one focus. (The Law of Ellipses)

 An imaginary line drawn from the center of the sun to the center of the planet will sweep out equal areas in equal intervals of time. (The Law of Equal Areas)

 The ratio of the squares of the periods of any two planets is equal to the ratio of the cubes of their average distances from the sun. (The Law of Harmonies) The Law of Ellipses

Kepler's first law - sometimes referred to as the law of ellipses - explains that planets are orbiting the sun in a path described as an ellipse. An ellipse can easily be constructed using a pencil, two tacks, a string, a sheet of paper and a piece of cardboard. Tack the sheet of paper to the cardboard using the two tacks. Then tie the string into a loop and wrap the loop around the two tacks. Take your pencil and pull the string until the pencil and two tacks make a triangle (see diagram at the right). Then begin to trace out a path with the pencil, keeping the string wrapped tightly around the tacks. The resulting shape will be an ellipse. An ellipse is a special curve in which the sum of the distances from every point on the curve to two other points is a constant. The two other points (represented here by the tack locations) are known as the foci of the ellipse. The closer together that these points are, the more closely that the ellipse resembles the shape of a circle. In fact, a circle is the special case of an ellipse in which the two foci are at the same location. Kepler's first law is rather simple - all planets orbit the sun in a path that resembles an ellipse, with the sun being located at one of the foci of that ellipse. The Law of Equal Areas

Kepler's second law - sometimes referred to as the law of equal areas - describes the speed at which any given planet will move while orbiting the sun. The speed at which any planet moves through space is constantly changing. A planet moves fastest when it is closest to the sun and slowest when it is furthest from the sun. Yet, if an imaginary line were drawn from the center of the planet to the center of the sun, that line would sweep out the same area in equal periods of time. For instance, if an imaginary line were drawn from the earth to the sun, then the area swept out by the line in every 31-day month would be the same. This is depicted in the diagram below. As can be observed in the diagram, the areas formed when the earth is closest to the sun can be approximated as a wide but short triangle; whereas the areas formed when the earth is farthest from the sun can be approximated as a narrow but long triangle. These areas are the same size. Since the base of these triangles are shortest when the earth is farthest from the sun, the earth would have to be moving more slowly in order for this imaginary area to be the same size as when the earth is closest to the sun. The Law of Harmonies

Kepler's third law - sometimes referred to as the law of harmonies - compares the orbital period and radius of orbit of a planet to those of other planets. Unlike Kepler's first and second laws that describe the motion characteristics of a single planet, the third law makes a comparison between the motion characteristics of different planets. The comparison being made is that the ratio of the squares of the periods to the cubes of their average distances from the sun is the same for every one of the planets.

Additionally, the same law that describes the T2/R3 ratio for the planets' orbits about the sun also accurately describes the T2/R3 ratio for any satellite (whether a moon or a man-made satellite) about any planet. The Law of Harmonies

Period Average T2/R3 Planet (yr) Distance (au) (yr2/au3)

Mercury 0.241 0.39 0.98

Venus .615 0.72 1.01

Earth 1.00 1.00 1.00

Mars 1.88 1.52 1.01

Jupiter 11.8 5.20 0.99

Saturn 29.5 9.54 1.00

Uranus 84.0 19.18 1.00

Neptune 165 30.06 1.00

Pluto 248 39.44 1.00 Hohmann

In orbital , the (/ˈhoʊmən/) is an elliptical orbit used to transfer between two circular orbits of different radii around a central body in the same plane. The Hohmann transfer often uses the lowest possible amount of propellant in traveling between these orbits, but bi-elliptic transfers can use less in some cases.

An example of a Hohmann transfer orbit between Earth and Mars, as used by the NASA InSight probe. Hohmann transfer orbit When used for traveling between celestial bodies, a Hohmann transfer orbit requires that the starting and destination points be at particular locations in their orbits relative to each other. Space missions using a Hohmann transfer must wait for this required alignment to occur, which opens a so-called launch window. For a space mission between Earth and Mars, for example, these launch windows occur every 26 months. A Hohmann transfer orbit also determines a fixed time required to travel between the starting and destination points; for an Earth-Mars journey this travel time is about 9 months.

An example of a Hohmann transfer orbit between Earth and Mars, as used by the NASA InSight probe. Bi-elliptic transfer orbit

In astronautics and , the bi-elliptic transfer is an orbital maneuver that moves a spacecraft from one orbit to another and may, in certain situations, require less delta-v than a Hohmann transfer maneuver.

While they require one more engine burn than a Hohmann transfer and generally require a greater travel time, some bi-elliptic transfers require a lower amount of total delta-v than a Hohmann transfer when the ratio of final to initial semi-major axis is 11.94 or greater, depending on the intermediate semi-major axis chosen.

A bi-elliptic transfer from a low circular starting orbit (blue) to a higher circular orbit (red). A boost at 1 makes the craft follow the green half-ellipse. Another boost at 2 brings it to the orange half-ellipse. A negative boost at 3 makes it follow the red orbit. At the Lagrange (also Lagrangian, L-, or libration) points the gravitational forces of the two large bodies cancel out in such a way that a small object placed in orbit there is in equilibrium in at least two directions relative to the center of mass of the large bodies.

Satellites which are destined for geosynchronous (GSO) or geostationary orbit (GEO) are (almost) always put into a highly elliptic GTO as an intermediate step for reaching their final orbit.

∆푣 needed for various orbital manoeuvers using C3 = Escape orbit conventional rockets; red arrows show where optional GEO=Geosynchronous orbit aerobraking can be performed in that particular direction, GTO=Geostationary transfer orbit black numbers give ∆푣 in km/s that apply in either direction. L4/5 = Earth–Moon L4L5 Lagrangian point Lower- ∆푣 transfers than shown can often be achieved, but LEO = Low Earth orbit involve rare transfer windows or take significantly longer.

A gravitational slingshot, gravity assist maneuver, or swing-by is the use of the relative movement and gravity of a planet or other astronomical object to alter the path and speed of a spacecraft, typically to save propellant / reduce expense.

Animation of Voyager 1's trajectory Animation of Voyager 2's trajectory from 5 September 1977 to 30 December 1981 from 20 August 1977 to 30 December 2000 Voyager 1 · Earth · Jupiter · Saturn · Sun Voyager 2 · Earth · Jupiter · Saturn · Uranus · Neptune · Sun When designing vehicle trajectory in addition to gravitational forces one also needs to take into account the effects of solar pressure, magnetic forces, and mechanical drag forces Overall effect of solar pressure depends on exact geometry and reflective/absorbing properties of surface materials of the vehicle, as well as vehicle position and orientation in space. This time varying effect introduces both non-zero forces and that when double integrated over long enough period of time can introduce quite substantial changes of vehicle’s trajectory.

푁 per second 푁 per second Total absorption Total reflection

ℎ휈 휃 휃 Ԧ 퐹Ԧ푁 퐹푁

휈 is the light frequency in the vehicle’s rest frame ℎ휈 퐹Ԧ Photon momentum is 푝 = hence 푇 푐 ℎ휈 ℎ휈 ℎ휈 퐹 = 푁 sin 휃 & 퐹 = 푁 cos 휃 퐹푁 = 2푁 sin 휃 & 퐹푇 = 0 푁 푐 푇 푐 푐 When designing vehicle trajectory in addition to gravitational forces one also needs to take into account the effects of solar pressure, magnetic forces, and mechanical drag forces Overall effect of drag forces depends on the exact geometry and properties of surface materials (for skin drag) of the vehicle, vehicle position and orientation in space relative to fluid, as well as properties of the fluid. This time varying effect introduces quite substantial forces and torques (especially if fluid speed in vehicle’s rest frame and fluid density are large).

Neither absorption nor elastic reflection 푁 per second

If fluid’s particle density (in fluid’s rest frame) is 푛 then one 푚푣 expects 푛푣 sin 휃 ∆푡 surface collisions per unit area per time ∆푡

휃 푣 is the fluid velocity in the vehicle’s rest frame 퐹Ԧ푁 Particle momentum is 푝 = 푚푣 hence 푪푫 ퟐ Normal drag 푭푵 = 푁푚푣 sin 휃 = 푛푣 sin 휃 푚푣 sin 휃 퐴 = 흆 풗 풔풊풏 휽 푨 (with 퐶퐷 ≅ 2) 퐹Ԧ푇 ퟐ Skin drag 퐹푇 = phenomenological model. Tricky to model theoretically. Typically 푚푣′ 퐹푁 ≫ 퐹푇 . However, important. One can not obtain cruising speed w/o 퐹푇. Drag force provides most of the lift force to aircraft

Typical values of angle of attack for standard aircraft are in the 1.5표 to 40 range depending on geometry and speed. Drag force provides most of the lift force to aircraft

Simplify by assuming zero path angle; angle of attack = pitch angle Here, we will ignore skin or drag Ԧ Ԧ 퐹푁 푑푟푎푔 푣Ԧ 퐹푇ℎ푟푢푠푡 휃

In a steady state:

Horizontal 퐹 cos 휃 = 퐹 sin 휃 푇ℎ푟푢푠푡 푁 푑푟푎푔 푚𝑔Ԧ

Vertical 퐹푇ℎ푟푢푠푡 sin 휃 + 퐹푁 푑푟푎푔 cos 휃 = 푚𝑔 퐿퐼퐹푇 Say for 3표 AoA, 푑푟푎푔 = 99.7% 퐿퐼퐹푇

퐿퐼퐹푇푑푟푎푔 퐹푁 푑푟푎푔 cos 휃 1 1 = = = = cos2 휃 퐹 sin 휃 2 퐿퐼퐹푇 퐹푇ℎ푟푢푠푡 sin 휃 + 퐹푁 푑푟푎푔 cos 휃 푇ℎ푟푢푠푡 + 1 tan 휃 + 1 퐹푁 푑푟푎푔 cos 휃 Drag force provides most of the lift force to aircraft

Simplify by assuming zero path angle; angle of attack = pitch angle Here, we will ignore skin or friction drag Ԧ Ԧ 퐹푁 푑푟푎푔 푣Ԧ 퐹푇ℎ푟푢푠푡 휃

In a steady state: Horizontal 퐹푇ℎ푟푢푠푡 cos 휃 = 퐹푁 푑푟푎푔 sin 휃 Vertical 퐹푇ℎ푟푢푠푡 sin 휃 + 퐹푁 푑푟푎푔 cos 휃 = 푚𝑔 푚𝑔Ԧ sin2 휃 퐹 + 퐹 cos 휃 = 푚𝑔 → 퐹 = 푚𝑔 cos 휃 푁 푑푟푎푔 cos 휃 푁 푑푟푎푔 푁 푑푟푎푔 휃 푣 퐶 2푚𝑔 cot 휃 Hence 퐷 퐴휌푣2 sin2 휃 = 푚𝑔 cos 휃 i.e. 푣 = 2 퐶퐷퐴휌 sin 휃 휌 푣 Drag force is also important for parachutes and vertical landing rockets

Initial state known 푟1

푥0, 푦0, 휃0, 푣푥0, 푣푦0, 휔0 푟2

휃퐿 휃푅 퐿2

Center of Mass 푥, 푦

퐿1

푗Ƹ

푖Ƹ

Landing site 0,0 푚𝑔Ԧ 휃 Altitude and human and other factors

Altitude Time of useful consciousness 5.5 km 20-30 min 6.7 km 10 min 9.1 km 1-2 min 10.7 km 30-60 s 12.2 km 15-20 s

The space zone starts at 13.7 km (8.5 푚𝑖푙푒푠). Even 100% oxygen intake at ambient pressure is insufficient to support lungs’ function (퐶푂2 and water vapor blocks enough oxygen from reaching lungs and blood). The pressurized cabin or suit is necessary for survival.

At about ퟏퟗ. ퟐ 풌풎 (12 푚𝑖푙푒푠) the fluid in body starts to vaporize (read “boil”) as the ambient pressure is equal to the of water at the body temperature of 37표퐶. This altitude is known as the Armstrong limit. At about ퟐퟒ 풌풎 (15 푚𝑖푙푒푠) it starts to become increasingly difficult to compress enough oxygen and nitrogen to supply aircraft pressurization system. Also, a huge problem is the large concentration of ozone which is poisonous to the and can not be separated from the ambient atmosphere. Hence, oxygen must be supplied separately (that is not from the ambient atmosphere).

Turbojets cease to function at about ퟑퟐ 풌풎 (20 푚𝑖푙푒푠). Ramjets cease to function at about ퟒퟓ 풌풎 (28 푚𝑖푙푒푠). No = No (EVA)

Sudden exposure to space without space suit will cause death by asphyxiation, i.e. or oxygen depravation. It takes about 15 seconds for O2 deprived blood to reach brain and which point one pass out and soon (after minute or two) dies.

Holding breath won’t help much to prevent hypoxia in the case of sudden decompression. In normal conditions, human lungs use muscular force for inhalation and exhalation is primary driven by tissue elasticity. Without ambient pressure the gas in lungs will suddenly expand in process that primarily depends on tissue elasticity. This will cause sudden drop in oxygen pressure and regular O2 diffusion from alveoli to bloodstream will be greatly reduced.

Many sources claim that sudden lung volume change can also result in lungs taring, rupturing, and allowing very large air bubbles to enter the bloodstream. These bubble can then block the blood vessels (embolism) and cause cardiac arrest, stroke, etc.

Sudden decompression can also cause when reduced ambience pressure reduces boiling point and cause formation of water vapor bubbles in the bloodstream and soft tissues. This can cause swelling and bruising due to the formation of the water vapor under the skin, and eventually embolism. The reduced boiling point will cause water vapor loss through nostril and mouth if opened. And after some time (likely ~30 seconds) freezing/drying effects in those areas; these will propagate through respiratory tract and cause eventual collapse of the lungs. Space suit A space suit must perform several functions to allow its occupant to work safely and comfortably, inside or outside a spacecraft. It must provide: A stable internal pressure. This can be less than Earth's atmosphere, as there is usually no need for the space suit to carry nitrogen (which comprises about 78% of Earth's atmosphere and is not used by the body). Lower pressure allows for greater mobility, but requires the suit occupant to breathe pure oxygen for a time before going into this lower pressure, to avoid decompression sickness. (At above the Armstrong limit, around 19,000 m (62,000 ft), water boils at body temperature and pressurized suits are needed.) Mobility. Movement is typically opposed by the pressure of the suit; mobility is achieved by careful joint design. Supply of breathable oxygen and elimination of carbon dioxide; these gases are exchanged with the spacecraft or a Portable Life Support System (PLSS) Temperature regulation. Unlike on Earth, where heat can be transferred by convection to the atmosphere, in space, heat can be lost only by thermal radiation or by conduction to objects in physical contact with the exterior of the suit. Since the temperature on the outside of the suit varies greatly between sunlight and shadow, the suit is heavily insulated, and air temperature is maintained at a comfortable level.

A communication system, with external electrical connection to the spacecraft or PLSS Means of collecting and containing solid and liquid bodily waste (such as a Maximum Absorbency Garment) Other requirements:

Shielding against ultraviolet radiation Limited shielding against particle radiation Means to maneuver, dock, release, and/or tether onto a spacecraft

Protection against small micrometeoroids, some traveling at up to 27,000 kilometers per hour, provided by a puncture-resistant Thermal Micrometeoroid Garment, which is the outermost layer of the suit. Experience has shown the greatest chance of exposure occurs near the gravitational field of a moon or planet, so these were first employed on the Apollo lunar EVA suits (see United States suit models next slide).

Generally, to supply enough oxygen for respiration, a space suit using pure oxygen must have a pressure of about 32.4 kPa (240 Torr; 4.7 psi), equal to the 20.7 kPa (160 Torr; 3.0 psi) of oxygen in the Earth's atmosphere at , plus 5.3 kPa (40 Torr; 0.77 psi) CO2 and 6.3 kPa (47 Torr; 0.91 psi) water vapor pressure, both of which must be subtracted from the alveolar pressure to get alveolar oxygen partial pressure in 100% oxygen , by the alveolar gas equation.

When space suits below a specific operating pressure are used from craft that are pressurized to normal (such as the ISS), this requires astronauts to "pre-breathe" (meaning pre-breathe pure oxygen for a period) before donning their suits and depressurizing in the air lock. This procedure purges the body of dissolved nitrogen, so as to avoid decompression sickness due to rapid depressurization from a nitrogen-containing atmosphere. Apollo/ A7L EVA and Moon suits. The Block II Apollo suit was the primary worn for eleven Apollo flights, three Skylab flights, and the US astronauts on the Apollo–Soyuz Test Project between 1968 and 1975. The pressure garment's nylon outer layer was replaced with fireproof Beta cloth after the Apollo 1 fire. This suit was the first to employ a liquid- cooled inner garment and outer micrometeroid garment. Beginning with the Apollo 13 mission, it also introduced "commander's stripes" so that a pair of space walkers will not appear identical on camera.

Extravehicular Mobility Unit (EMU) used on both the Space Shuttle and International Space Station (ISS). The EMU is an independent anthropomorphic system that provides environmental protection, mobility, life support, and communications for a Space Shuttle or ISS crew member to perform an EVA in Earth orbit. Used from 1982 to present, but only available in limited sizing as of 2019. More on space suits

The new suit that will be worn on Artemis missions is called the Exploration Extravehicular Mobility Unit, or xEMU. https://www.space.com/nasa-xemu-moon-spacesuit-2023-test.html

The lunar soil is composed of tiny glass-like shards, so the new suit has a suite of dust-tolerant features to prevent inhalation or contamination of the suit’s life support system or other spacecraft. The suit also is built to withstand temperature extremes of minus 250 degrees Fahrenheit in the shade and up to 250 degrees in the sun.

The Portable Life Support System is the familiar backpack astronauts wear on spacewalks that houses the suit’s power and breathable air and removes exhaled carbon dioxide and other toxic gasses, odors and moisture from the suit. It also helps regulate temperature and monitors overall suit performance, emitting warnings if resources fall low, or if there is a system failure. Miniaturization of electronics and plumbing systems have made it possible to build in duplicates for much of the system, making some failures less of a concern.

The pressure garment is the human shaped portion of the spacesuit that enables astronaut mobility and protects their body. The new lower torso includes advanced materials and joint bearings that allow bending and rotating at the hips, increased bending at the knees, and hiking-style boots with flexible soles. On the upper torso, in addition to the updated shoulder placement, other shoulder enhancements allow astronauts to move their arms more freely and easily lift objects over their heads or reach across their body in the pressurized suit. Apollo shoulder mobility was enabled by pleats in the fabric with cable pulleys that provided mechanical advantage to move the shoulders up and down but limited the ability to rotate the joint. The new shoulders minimize the effort required for full mobility and include bearings that allow full rotation of the arm from shoulder to wrist. More on space suits

Feitian EVA space suit. New generation indigenously developed Chinese-made EVA space suit also used for the Shenzhou 7 mission. The suit was designed Sokol ("falcon") suits worn by Orlan ("bald eagle") suits is for a spacewalk mission of up to seven hours. Soyuz crew members during Russia's current EVA suit. Chinese astronauts have been training in the out-of- launch and reentry They were first Used from 1977 to present. capsule space suits since July 2007, and movements worn on Soyuz 12. They have been are seriously restricted in the suits, with a mass of used from 1973 to present. more than 110 kilograms (240 lb) each. More on space suits A mechanical counterpressure (MCP) suit, partial pressure suit, direct compression suit, or space activity suit (SAS) is an experimental spacesuit which applies stable pressure against the skin by means of skintight elastic garments. The SAS is not inflated like a conventional spacesuit: it uses mechanical pressure, rather than air pressure, to compress the human body in low-pressure environments. Development was begun by NASA and the Air Force in the late 1950s and then again in the late 1960s, but neither design was used. Research is under way at MIT on a "Bio-Suit" System which is based on the original SAS concept. Newman has worked extensively in biomechanics, especially in the field of computerized measurement of human movement. As with gas-filled suits, Newman has used the principle of "lines of non-extension", a concept originated by Arthur Iberall in the late 1940s, to place the tension elements along lines of the body where the skin does not stretch during most normal movements. The primary structure of the BioSuit is built by placing elastic cords along the lines of non-extension. Thus, whatever pressure they provide will be constant even as the wearer moves. In this way, they can control the mechanical counter-pressure the suit applies. The rest of the suit is then built up from The NASA Paul Webb’s SAS: the spandex lying between the primary pressure cords. complete multi-layer suit and positive-pressure helmet, lacking Bio-Suit by Professor only the backpack. (taken ~1971) Dava Newman, MIT Effects of microgravity on human health Effects of microgravity on human health Effects of microgravity on human health

https://www.slideshare.net/FJHScience/2013-space-human-phys-and-anatomy Plasma interactions: sputtering Sputtering is a phenomenon in which microscopic particles of a solid material are ejected from its surface, after the material is bombarded by energetic particles of a plasma or gas.

Occurs naturally in space and can be an unwelcome source of wear for spacecraft and precision components. Negatively charged regions of solar arrays may suffer erosion. In case of ion-thruster that operate with exhaust ion energies in several hundred to several thousand eV range, sputtering can cause degradation of thruster and nearby surfaces.

However, this phenomenon can be also used to perform precise etching and deposit thin film layers for optical coatings, semiconductor devices and nanotechnology products. It is a physical vapor deposition technique.

Physical sputtering has a well-defined minimum energy threshold, equal to or larger than the ion energy at which the maximum energy transfer from the ion to a target atom equals the binding energy of a surface atom.

The average number of atoms ejected from the target per incident ion is called the "sputter yield". The sputter yield depends on several things: the angle at which ions collide with the surface of the material, how much energy they strike it with, their masses, the masses of the target atoms, and the target's surface binding energy. If the target possesses a crystal structure, the Xenon is popular fuel for ion thrusters. orientation of its axes with respect to the surface is an important factor. Plasma interactions: Debye radius Gauss law in differential form 2 휌 휌 1 푑𝑖푣 퐸 = 훻 ∙ 퐸= 훻 ∙ −𝑔푟푎푑 푉 = 훻 ∙ −훻푉 = −훻 푉 = −∆푉 = = = ൫휌푝푙푎푠푚푎 + 휀0휀푟 휀 휀 1 1 −∆푉 = ෍ 휌푝푙푎푠푚푎 𝑖 + 휌푑𝑖푠푡푢푟푏푎푛푐푒 = ෍ 푞𝑖 푛𝑖 + 휌푑𝑖푠푡푢푟푏푎푛푐푒 휀 𝑖 휀 𝑖

푞푖푉 푛𝑖,0(푟) spatial i-th concentration without disturbance − 푘푇 This term is zero for electrically 푛𝑖 = 푛𝑖,0푒 푛𝑖(푟) spatial i-th concentration with disturbance macroscopically neutral plasma 2 푞𝑖푉 1 푞𝑖 푛𝑖,0 푛𝑖 ≅ 푛𝑖,0 1 − → −∆푉 = ෍ 푞𝑖 푛𝑖,0 − ෍ + 휌푑𝑖푠푡푢푟푏푎푛푐푒 푘푇 휀 𝑖 𝑖 푘푇 for point like charge Q 푟 푄 − 휀푘푇 휆퐷 휌푑𝑖푠푡푢푟푏푎푛푐푒 = 푄 훿 푟 → 푉 = 푒 푤𝑖푡ℎ 휆퐷 = 2 4휋휀푟 σ𝑖 푞𝑖 푛𝑖,0

휀푘 In case of different temperatures of plasma particle “species” 휆퐷 = 2 σ𝑖 푞𝑖 푛𝑖,0/푇𝑖 Plasma interactions: spacecraft charging

Spacecraft charging is the process by which a spacecraft or selected components accumulate an electric charge from the environment. For a plasma in thermal equilibrium with the same temperature for electrons and protons, the difference in mass results in electrons having a speed in average 43 times that of protons. As a result, the flux or current of electrons onto a spacecraft is much higher than that of the ions, resulting in the potential for negatively charged surfaces relative to the plasma while in shadow. The emergence of the spacecraft into the sun can create positively charged surfaces, due to photoelectric effect (photons kicking electrons out of the spacecraft surface). High negatively charged surfaces relative to the plasma or in proximity to uncharged surfaces, and negatively charged surfaces in proximity to positively charged surfaces, can cause electrostatic discharges. A high negative potential will also accelerate ions, thereby enhancing sputtering. Arcing can produce EM fields that can cause anomalies in operation or damage spacecraft electronics. Arcing can also physically damage spacecraft materials, changing their thermal and electrical properties.

The electric potential 푉푆 of the spacecraft is defined relative to the plasma potential 푉푃, which is assumed to be zero or ground. This potential difference will distort the more mobile electron distribution, which will tend to neutralize ESA EURECA satellite solar the charge at distances from the spacecraft characterized by the Debye length. array sustained arc damage. Plasma interactions: spacecraft charging The Debye lengths for a low-Earth-orbit (LEO) spacecraft (≤ 1000 푘푚 altitude) are on the order of centimeters, while the Debye lengths for geostationary equatorial orbits (GEO) are on the order of tens of meters. The larger volume of distorted plasma in GEO causes a complex disturbed environment including a depleted wake effect, which can support surface charging. At LEO, there are large numbers of plasma particles near the spacecraft that are able to neutralize differentially charged surfaces by providing conductive paths between them. Spacecraft charging is described as either absolute or differential. Absolute charging occurs when the spacecraft as a whole accumulates a net charge so that it has a potential difference relative to its environment. Differential charging occurs when different portions of the spacecraft charge to different electric potentials. Differential charging can be either internal or surface. Potential difference that exceeds breakdown threshold voltage will lead to electrostatic discharging (arcing). The basic process for spacecraft charging is the balancing of currents at the equilibrium potential, when all currents sum to zero. Thermal Control

In spacecraft design, the function of the thermal control system (TCS) is to keep all the spacecraft's component systems within acceptable temperature ranges during all mission phases.

The thermal control subsystem can be composed of both passive and active items and works in two ways:

COOLING: Protects the equipment from overheating, either by thermal insulation from external heat fluxes (such as the Sun or the planetary infrared and albedo flux), or by proper heat removal from internal sources (such as the heat emitted by the internal electronic equipment).

HEATING: Protects the equipment from temperatures that are too low, by thermal insulation from external sinks, by enhanced heat absorption from external sources, or by heat release from internal sources.

Common TCS elements are Coatings, Multilayer insulation, Louvers, Heaters, Radiators, Heat pipes, etc. Thermal Control

Coatings are the simplest and least expensive of the TCS techniques. A coating may be paint or a more sophisticated chemical applied to the surfaces of the spacecraft to lower or increase heat transfer. The characteristics of the type of coating depends on their absorptivity, emissivity, transparency, and reflectivity. The main disadvantage of coating is that it degrades quickly due to the operating environment. Multilayer insulation (MLI) is the most common passive TCS element used on spacecraft. MLI prevents both heat losses to the environment and excessive heating from the environment. Spacecraft components such as propellant tanks, propellant lines, batteries, and solid rocket motors are also covered in MLI blankets to maintain ideal operating temperature. MLI consist of an outer cover layer, interior layer, and an inner cover layer. The outer cover layer needs to be opaque to sunlight, generate a low amount of particulate contaminates (common materials used for the outer layer are fiberglass woven cloth impregnated with PTFE Teflon, PVF reinforced with Nomex bonded with polyester adhesive, and FEP Teflon). The interior layer needs to have a low emittance (commonly used material is Mylar aluminized on one or both sides). The interior layers are usually thin compared to the outer layer to save weight and are perforated to aid in venting trapped air during launch. The inner cover faces the spacecraft hardware and is used to protect the thin interior layers. Inner covers are often not aluminized in order to prevent electrical shorts (materials used for the inner covers are Dacron and Nomex netting. mylar is not used because of flammability concerns). Louvers are active TCS element most commonly placed over external radiators; most commonly used louver is the bimetallic, spring-actuated, rectangular blade louver also known as venetian-blind louver. Louver radiator assemblies consist of five main elements: baseplate, blades, actuators, sensing elements, and structural elements.. A louver in fully open state can reject six times as much heat as it does in fully closed state, with no power required to operate it. Thermal Control Heaters: Patch heater, which consists of an electrical-resistance element sandwiched between two sheets of flexible electrically insulating material, such as Kapton, is the most common heater. Cartridge heater, is often used to heat blocks of material or high-temperature components such as propellants. This heater consists of a coiled resistor enclosed in a cylindrical metallic case. Radioisotope heater unit also known as RHU; at the center of each RHU is a radioactive material, which decays to provide heat. The most commonly used material is plutonium dioxide.

Radiators, removing excess waste heat, come in several different forms, such as spacecraft structural panels, flat- plate radiators mounted to the side of the spacecraft, and panels deployed after the spacecraft is in orbit. Radiators reject heat by infrared (IR) radiation from their surfaces. The radiating power depends on the surface's emittance and temperature. The radiator must reject both the spacecraft waste heat and any radiant-heat loads from the environment. Most radiators are therefore given surface finishes with high IR emittance to maximize heat rejection and low solar absorptance to limit heat from the Sun.

Heat pipes use a closed two-phase liquid-flow cycle with an evaporator and a condenser to transport relatively large quantities of heat from one location to another with very little or no electric power.

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