Marspedia - User contributions [en]

Landing on Mars

2019-08-17T12:21:26Z

Pb:

A typical mission to land on Mars consists of atmospheric entry phase using an aeroshell, a parachute descent phase, a terminal descent phase, and ends with the spacecraft touching down somewhere within a predetermined [[Landing Ellipse|landing ellipse]]. The spacecraft can enter the atmosphere either directly from the [[Earth-Mars_Transfer_Trajectory|Earth-Mars transfer orbit]] or it can descend after first being captured into Mars orbit. The parachute descent phase begins in the supersonic flight regime and normally utilizes ring-sail or disk-gap-band supersonic parachute. The terminal descent phase has been achieved with retrorockets and lanyards/skycranes, and spacecraft have been cushioned on touchdown with foam and airbags.

To date there have been 10 successful soft landings: Mars 3, Vikings 1 and 2, Pathfinder, Beagle 2, MER A and B (Spirit and Opportunity), Phoenix, MSL (Curiosity), and InSight. Mars 3 and Beagle 2 have been declared as successful soft landings though the spacecraft did not operate correctly once on the surface. There have been 4 failed landing attempts: Mars 2, Mars 6, Mars Polar Lander, and the Schiaparelli EDM lander.

Wind tunnels, rocket sleds, and drop tests have been used to test equipment for entry, descent, and landing. Work continues on new technology such as supersonic retropropulsion, ballutes, low-density-supersonic-decelerators (and other expandable entry shields), biconic heat shields (and other entry vehicle shapes), and new ablative or non-ablative heat shield materials.

==Concepts==

Marspedia contributors have supplied the following concepts for EDL technology.

If we need a four hundred foot diameter parachute manufactured in space out of aluminum oxide fiber and sent to Mars in stiff deployed condition instead of being packed, we will not learn about it unless we see a need to experiment. Such a parachute might merit investigation. It would avoid opening shock and might be sufficiently heat resistant to maintain structural integrity during the entire descent in Mars' low gravity well. The larger the diameter of the parachute, the less the max g loading. So let us be honest with ourselves about all necessary colonization technology.

The expected max temperature for ballistic entry into Mars atmosphere is expected to be a thousand or more Kelvin degrees above the melting point of aluminum oxide so coating course aluminum oxide fibers with potassium oxide which decomposes at 490 Centigrade might protect the fibers through atmospheric entry by ablative cooling or it might not. A mixture of potassium and sodium oxides as a coating or Teflon as a coating are things that are conceivable. Engineers in this specialty would have a better idea.

===High Lift Vertical Landing Vehicle===

Another alternative with a greater probability of working, but possibly high cost, is a delta winged entry vehicle or lifting body with insulation like that on the space shuttle. The insulation would be somewhat cheaper because Mars atmospheric entry is less demanding than Earth reentry. After losing most of its orbital velocity to the atmosphere by heating the atmosphere in passing, this vehicle would fly supersonic close to the ground then ignite its rockets for landing. Then it would perform a Pugachev's Cobra<ref>http://en.wikipedia.org/wiki/Pugachev%27s_Cobra</ref> maneuver losing horizontal velocity by drag and by rocket thrust. It would then touch down on its tail. Rocket thrust directly into the supersonic slipstream of Mars' atmosphere will not work to safely land on Mars because the supersonic slipstream that the lander flies into would carry the noise of the rocket exhaust right back to the lander. The potential for the chaotic forces of this rocket noise to destabilize the lander's orientation and damage its structure rule out this technique. In the Pugachev's Cobra maneuver, rocket thrust is never directed directly into the supersonic slipstream. The rocket thrust always has a vertical component while the slipstream moves horizontally until the slip stream velocity is reduced to a negligible value.

This sort of vehicle might approach the point of entering a Pugachev's Cobra maneuver by flying horizontally near Mars' surface while increasing angle of attack to maintain lift while killing velocity. At a pitch attitude of 45 degrees there is little lift left to be gained by increasing angle of attack. This should occur at about Mach 2.5, which is about 600 meters per second on Mars. Then the rockets are ignited generating two Mars gravities of acceleration and the angle of attack is further increased past 90 degrees to generate negative lift and keep the vehicle in horizontal flight. As the speed decreases and negative lift generated by the wings decreases, the pitch angle is increased to reduce the component of rocket thrust in the vertical direction and increase the component of rocket thrust directed to braking. As the vehicle eventually slows to a stop in horizontal motion, a combination of throttling and thrust deflection reduces thrust to about 1 Mars gravity, the vehicle moves to a 90 degree pitch angle and settles on its tail. A guesstimate of the required rocket delta V for killing the last 600 meters per second and landing in this way is about 850 meters per second. This includes the amount of speed lost to atmospheric drag and very substantial gravity losses.

===Another Alternative is the Sky Crane===

<blockquote>the 2009 Mars Science Laboratory (MSL) rover, weighing 775 kilograms (versus MER at 175.4 kilograms each) requires an entirely new landing architecture. Too massive for airbags, the small-car sized rover will use a landing system dubbed the Sky Crane. "Even though some people laugh when they first see it, my personal view is that the Sky Crane is actually the most elegant system we've come up with yet, and the simplest," said Manning. MSL will use a combination of a rocket-guided entry with a heat shield, a parachute, then thrusters to slow the vehicle even more, followed by a crane-like system that lowers the rover on a cable for a soft landing directly on its wheels. Depending on the success of the Sky Crane with MSL, it's likely that this system can be scaled for larger payloads, but probably not the size needed to land humans on Mars. (See Ref #1) </blockquote>

===A Sure Way to Land on Mars===
A sure but expensive way to land on Mars with a ten metric ton vehicle is to build a heat shield in orbit around Earth and send it to Mars as part of the spacecraft. After the heat shield slows the spacecraft, rockets bring it to a safe stop on Mars. Since Mars' atmosphere at the surface is one hundredth the density of Earth's atmosphere at the surface, make the heat shield proportionally bigger. Considering that the 12,250 pound Apollo command module was 12.8 feet in diameter, a ten metric ton Mars lander should have a 52 meter diameter heat shield. Assembled from 127 roughly hexagonal pieces about 4 meters in diameter, this would be a hexagonal heat shield instead of a round one. That should do. Each hexagonal piece would have a layer of ablative material on one side of a hexagon of aircraft grade aluminum. Aluminum t cross section extrusions would be fastened to the Aluminum sheet as stiffeners. In orbit, two hexagon sections would have their ablative sections butted against each other, protrusions fitting into cavities. A small gap would remain between the aluminum sheets. A 2 inch strip along the edge of each aluminum sheet would be pre-coated with brazing material. A 4 inch wide strip of aluminum to join them would likewise be coated with brazing material on one side. A ridge on the joining strip would fit in the gap between the aluminum hexagons. Then an iron heated to the right temperature would be placed on the joining strip and left for the right time. When the iron is removed and the piece cools the two hexagons make one piece with brazing material partially filling the gap between the two hexagons and rounding out the corner where the hexagons meet the joining strip. Likewise, a trusswork joining the stiffeners of all of the hexagons would be assembled by the same brazing technique and make the whole heat shield one strong rigid light weight piece. Some work has already been done considering robotic truss assembly on orbit. [http://www.ri.cmu.edu/publication_view.html?pub_id=1691 Light-Weight Mobile Robot for Space Station Trusswork] A cone section for the sphere-cone reentry vehicle would also need to be built or a somewhat different shaped section serving the same aerodynamic function, avoiding excessive, uncontrolled and chaotic side slip.

The advantage of sending up a ten ton vehicle, many pieces of heat shield and a robotic assembly station two make a big heat shield as compared to sending up a vehicle with heat shield and parachutes on an Ares V is that the big assembled-on-orbit heat shield would allow a 10 ton vehicle to land cargo safely on Mars while the Ares V scheme would not land cargo or people safely on Mars. Mars direct would do no better. See [http://www.universetoday.com/2007/07/17/the-mars-landing-approach-getting-large-payloads-to-the-surface-of-the-red-planet/ The mars landing Approach: Getting Large Payloads to the Surface of the Red Planet]

Just as all economic activity in orbit so far has been done by robots, assembling a spacecraft to go to Mars should be done by robots and setting up the infrastructure for people to survive on Mars should be done by robots. There are some technical difficulties with this approach that must be addressed, but they seem likely to be amenable to solution.

===Terminal Velocity of the Big Heat Shield Vehicle===
For a rough estimate take as a starting point the estimated terminal velocity for the Apollo reentry vehicle. Estimate that the square of the terminal velocity is directly proportional to mass and the force of gravity and inversely proportional to air density and cross sectional area. The 10 metric ton vehicle with a 52 meter diameter heat shield should then have a terminal velocity less than 56 meters per second (125 miles per hour) at Mars' surface. If near the end of this descent a hole is burst through the bottom of the heat shield in the center right under the payload and hinges swing that portion of the heat shield out of the way, then retro rockets can fire at an altitude of 160 meters for 5.1 seconds with an acceleration of 11 meters per second squared and bring the payload to a stop about 17 meters above the heat shield that crashes into Mars. A few more seconds for horizontal maneuvering and throttling the rocket motor brings the payload safely to Mars on landing legs. A guesstimate of the required rocket delta V for this maneuver is about 68 meters per second.

===Supersonic retro-propulsion landing===
SpaceX has proposed supersonic retro-propulsive landing as a method to land large payloads on Mars.<ref>https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170008725.pdf</ref>

*An initial speed reduction using aerodynamic breaking.
*Retro propulsion in the supersonic flight regime is used to support and slow the vehicle during part of the flight path.<ref>Youtube video of Mars landing, SpaceX https://www.youtube.com/watch?v=LQTnWEHl5qU</ref>
*More aerodynamic breaking and conventional retro propulsion is used to land.
*SpaceX has demonstrated the capability of very precise landing using these methods on Earth, that could be transposed to Mars.

==References==
<references />

[[Category:Human Mission Architecture]]

Concrete

2019-03-19T03:40:51Z

Pb: some sulfur-regolith details

'''Concrete''' is a well known material for building [[house]]s and [[infrastructure]] elements. It has excellent characteristics for protection against [[radiation]] and small [[meteorites]]. Possibly, concrete can be made in situ on [[Mars]], using [[local resources]].

==Hydraulic cement==
Unlike on Earth's Moon there seems to be plenty of water on Mars, but hydraulic cement also requires [[calcium]], [[silicon]] and [[aluminum]]. It is unclear whether these substances can be found on Mars in a form that allows a simple processing.

==Waterless concrete==
There are ideas of making waterless concrete from [[sulfur]] and [[regolith]]<ref>https://arxiv.org/pdf/1512.05461.pdf "A Novel Material for In Situ Construction on Mars:
Experiments and Numerical Simulations." Lin Wan et al. 2016</ref>.

The ultimate strength and tensile strength was found to be best at a mixing ratio of 50% sulfur and 50% JSC Mars-1A regolith simulant sieved to a maximum particle size of 1 mm. The concrete was found to have a compression strength of > 50 MPa, a flexural strength of 1.75 MPa, and a splitting tensile strength of 3.9 MPa.

Utilizing sulfur-regolith concrete is possible on Mars, but not the Moon. On the moon, the concrete mass would be gradually lost due to sublimation of sulfur in vacuum, and the large temperature swings between lunar day and night which compromise the structure. Sulfur-regolith concrete is stable under martian conditions and would not experience a loss in mass due to sublimation.

Mars is considered a sulfur-rich planet, but it in unclear where sulfur may be and if it is present in a form suitable for the production of sulfur concrete.

==Reinforcement==
The stability of concrete can be increased significantly by [[tension glass fibers]] or reinforcing [[steel]].

==See also==
*[[Sintered regolith]]
*[[Brick]]

==External links==
*[http://scitation.aip.org/getabs/servlet/GetabsServlet?prog=normal&id=JAEEEZ000020000004000220000001&idtype=cvips&gifs=yes Journal of Aerospace Engineering - Analysis of Lunar-Habitat Structure Using Waterless Concrete and Tension Glass Fibers]

[[Category:Materials]]

Deep Space Network

2018-12-20T05:54:29Z

Pb:

The '''Deep Space Network''' (DSN) is an international network of antennas owned and operated by NASA located on [[Earth]] and dedicated to the communication with interplanetary missions and astronomical radio observations. Currently, there are three sites based approximately 120° apart in the Mojave Desert, California (USA), Madrid (Spain) and Canberra (Australia). The dispersion of antennas allow mission controllers to keep constant watch on all interplanetary missions regardless on Earth's rotation.

This network may be used for an [[internet|interplanetary internet link]].

[[File:CCSDS_DDOR_Observation_Geometry.png|right|thumb|320px| Geometry of Delta-DOR for navigation]]

==Use for Communication==
Typically, one or more of the 34 meter diameter antennas is used to communicate with spacecraft at Mars. The data rate of the downlink from Mars can be anywhere from a few bits/second to on the order of 1 Mpbs. Most missions use the DSN at X-band, but a small number of missions are using S-band or Ka-band.

==Use for Navigation==
The DSN can provide ranging and doppler information for use in orbit determination of spacecraft. Long two-way doppler tracking is commonly used.<ref>https://descanso.jpl.nasa.gov/monograph/series1/Descanso1_C03.pdf</ref>

Delta-DOR is a navigation technique that is similar to Very Long Baseline Interferometry. Delta-DOR requires two base stations separated by a mile or two, a quasar, and a target spacecraft which must emit several tones spanning a few MHz. From this, angular measurement is obtained. <ref>https://public.ccsds.org/Pubs/500x1g1.pdf</ref>

==Downtime==

The DSN requires a line of sight between the antenna and the vehicle, which imposes obvious constraints on the system. For rovers and landers, in order to use a Direct to Earth (DTE) connection, the vehicle must be on the Earth-facing side of the planet, which means that DTE communication is only possible for about half of the Martian day. DTE connections for rovers and landers are uncommon, and instead the surface asset will use a UHF proximity link to transmit data to an orbiter, which relays the signal to DSN over X-band or Ka-band.

There are six orbiters around Mars. Orbiters are occulted when their orbit takes them behind Mars from the viewpoint of the Earth station.

Interference from the solar radiation can cause loss of signal when the angle between the Earth, Sun, and Mars is less than 5°. This loss of signal occurs during a superior conjunction, which occurs once per 26 month synodic period, and can last from a day to over a week depending on the relative geometry due to the non-co-planar orbits of Earth and Mars.

==Loading Considerations==

The Deep Space Network supports approximately 35 missions as of 2018, and significant planning is required to schedule time on the DSN for all space assets that need it. Mars vehicles represent only a fraction of the total users, but when non-Mars missions appear close in the sky to Mars there can be a significant bottleneck. One technique that is used to relieve strain on the system imposed by the many Mars missions is Multiple Spacecraft per Antenna (MSPA). With this technique, one 34 meter antenna can be used to simultaneously downlink data from up to four spacecraft, and uplink to one spacecraft.<ref>https://arc.aiaa.org/doi/pdf/10.2514/6.2018-2502</ref>

==Alternatives==

The Sardinia Radio Telescope, a 64 meter aperture steerable antenna in Italy, can be used for deep space communications. NASA may potentially be able to use this if needed.

The former Soviet Union constructed a number of large aperture antennas for deep-space communication to support their Mars missions. More recently, this network was to be used to support the 2011 Fobos-Grunt mission, but the mission failed before reaching deep space.

China has several large antennas. Most are located in mainland China, but a new groundstation was recently constructed in Argentina. The network was used to track Chang'e 2, which had a mission profile similar to NASA's Clementine mission, with operations in cis-lunar space followed by operations around an asteroid.

The American company Atlas Space Operations is building a commercial capability which they claim will be equivalent to one of the DSN's 70 meter dishes. Atlas LINKS (TM) will be an array of small antennas, not a large aperture antenna like those the DSN utilizes.

==Open issues==
*How long is the gap and how often does it occur?

==External links==
*[http://deepspace.jpl.nasa.gov/dsn/ NASA site for the DSN]

{{stub}}

[[category:Communication]]

File:CCSDS DDOR Observation Geometry.png

2018-12-20T05:51:38Z

Pb: Figure representing the geometry when performing Delta-DOR measurements for spacecraft navigation. Obtained from the CCSDS May 2013 Green Book, Delta-DOR Technical Characteristics and Performance. Available here: https://public.ccsds.org/Pubs/500x1g1.pdf

== Summary ==
Figure representing the geometry when performing Delta-DOR measurements for spacecraft navigation.

Obtained from the CCSDS May 2013 Green Book, Delta-DOR Technical Characteristics and Performance.
Available here: https://public.ccsds.org/Pubs/500x1g1.pdf

Deep Space Network

2018-12-20T05:49:21Z

Pb: /* Use for Navigation */

The '''Deep Space Network''' (DSN) is an international network of antennas owned and operated by NASA located on [[Earth]] and dedicated to the communication with interplanetary missions and astronomical radio observations. Currently, there are three sites based approximately 120° apart in the Mojave Desert, California (USA), Madrid (Spain) and Canberra (Australia). The dispersion of antennas allow mission controllers to keep constant watch on all interplanetary missions regardless on Earth's rotation.

This network may be used for an [[internet|interplanetary internet link]].

==Use for Communication==
Typically, one or more of the 34 meter diameter antennas is used to communicate with spacecraft at Mars. The data rate of the downlink from Mars can be anywhere from a few bits/second to on the order of 1 Mpbs. Most missions use the DSN at X-band, but a small number of missions are using S-band or Ka-band.

==Use for Navigation==
The DSN can provide ranging and doppler information for use in orbit determination of spacecraft. Long two-way doppler tracking is commonly used.<ref>https://descanso.jpl.nasa.gov/monograph/series1/Descanso1_C03.pdf</ref>

Delta-DOR is a navigation technique that is similar to Very Long Baseline Interferometry. Delta-DOR requires two base stations separated by a mile or two, a quasar, and a target spacecraft which must emit several tones spanning a few MHz. From this, angular measurement is obtained. <ref>https://public.ccsds.org/Pubs/500x1g1.pdf</ref>

==Downtime==

The DSN requires a line of sight between the antenna and the vehicle, which imposes obvious constraints on the system. For rovers and landers, in order to use a Direct to Earth (DTE) connection, the vehicle must be on the Earth-facing side of the planet, which means that DTE communication is only possible for about half of the Martian day. DTE connections for rovers and landers are uncommon, and instead the surface asset will use a UHF proximity link to transmit data to an orbiter, which relays the signal to DSN over X-band or Ka-band.

There are six orbiters around Mars. Orbiters are occulted when their orbit takes them behind Mars from the viewpoint of the Earth station.

Interference from the solar radiation can cause loss of signal when the angle between the Earth, Sun, and Mars is less than 5°. This loss of signal occurs during a superior conjunction, which occurs once per 26 month synodic period, and can last from a day to over a week depending on the relative geometry due to the non-co-planar orbits of Earth and Mars.

==Loading Considerations==

The Deep Space Network supports approximately 35 missions as of 2018, and significant planning is required to schedule time on the DSN for all space assets that need it. Mars vehicles represent only a fraction of the total users, but when non-Mars missions appear close in the sky to Mars there can be a significant bottleneck. One technique that is used to relieve strain on the system imposed by the many Mars missions is Multiple Spacecraft per Antenna (MSPA). With this technique, one 34 meter antenna can be used to simultaneously downlink data from up to four spacecraft, and uplink to one spacecraft.<ref>https://arc.aiaa.org/doi/pdf/10.2514/6.2018-2502</ref>

==Alternatives==

The Sardinia Radio Telescope, a 64 meter aperture steerable antenna in Italy, can be used for deep space communications. NASA may potentially be able to use this if needed.

The former Soviet Union constructed a number of large aperture antennas for deep-space communication to support their Mars missions. More recently, this network was to be used to support the 2011 Fobos-Grunt mission, but the mission failed before reaching deep space.

China has several large antennas. Most are located in mainland China, but a new groundstation was recently constructed in Argentina. The network was used to track Chang'e 2, which had a mission profile similar to NASA's Clementine mission, with operations in cis-lunar space followed by operations around an asteroid.

The American company Atlas Space Operations is building a commercial capability which they claim will be equivalent to one of the DSN's 70 meter dishes. Atlas LINKS (TM) will be an array of small antennas, not a large aperture antenna like those the DSN utilizes.

==Open issues==
*How long is the gap and how often does it occur?

==External links==
*[http://deepspace.jpl.nasa.gov/dsn/ NASA site for the DSN]

{{stub}}

[[category:Communication]]

Sintered regolith

2018-12-20T05:33:35Z

Pb: /* Methods */

[[Image:UniversalBricks02.jpg|thumb|right|220px|Construction elements from sintered regolith]]

'''Sintered regolith''' has been proposed as a construction material on the Moon. This technology might be extended to Mars. [[Sintering]] is the fusion of mineral particles through the application of heat. The particles are heated just enough to induce fusion, but not enough to fully melt.

==Methods==
===Laser Sintering===
Laser sintering is used in [[rapid prototyping]] applications.

===Solar Sintering===
Solar energy can be focused on the regolith to achieve the 1300 degree C temperature required for sintering.

===Microwave Sintering===
A microwave can be used to heat regolith.

===Furnace Sintering===
Large kilns are used to heat the [[regolith]] as it is held in a mould.

==Use==
===Construction===
Sintered Regolith blocks are a possible construction material. Further testing is needed to determine the structural characteristics of such blocks. Possibilities include [[Universal bricks]] and [[arch segments]].
===Rapid Prototyping===
Laser sintering and additive manufacturing can produce custom items. Laser sintering comprises thin, sequential layers of media are laid down, and sintered together to form the item. Additive manufacturing consists of building up products layer by layer in a semi-molten state until they set into stabilize in solid form.

==See Also==
* [[lunarp:Sintered regolith|Sintered regolith on Lunarpedialunarp]].

[[Category:Synthesis]]

Earth-Mars Transfer Trajectory

2018-10-31T13:23:55Z

Pb: /* Leaving Earth */

[[Image: InSight Trajectory.jpg|thumb|right|px|Earth-Mars transfer trajectory of NASA's InSight lander, including planned trajectory correction maneuvers (TCMs)]]

An Earth-Mars transfer trajectory is an orbital path which a spacecraft follows to travel between Earth and Mars. Several types of trajectories have been studied, but all will satisfy the following conditions:

* The starting point must be near the Earth in its orbit around the sun
* The ending point must intersect Mars in its orbit around the sun
* The intervening trajectory must be heliocentric, though one or more gravitational swing-bys of other bodies are allowed

Not all Mars transfer orbits are Hohmann transfers. This is due to the difference in the plane of Earth and Mars's orbit, and can also be due to constraints on launch windows.
There are many variations on this theme, such as whether the spacecraft ends in a Mars-centered orbit or if the spacecraft directly enters the atmosphere from the heliocentric transfer orbit. Other less common variations include the starting orbit at Earth, which could be a low Earth orbit or could instead the mission could begin from geostationary transfer orbit (GTO) as in the MEGA proposals of the late 90's, or one could vary the heliocentric portion of the flight to include a low-thrust trajectory.

[[Image: Porkchopplot.png|thumb|left|px|Porkchop plot for 2018 launch opportunity <ref>https://trs.jpl.nasa.gov/bitstream/handle/2014/44336/13-0679_A1b.pdf?sequence=1</ref>]]

==Launch Period==

A launch period is a span of days during which a launch vehicle can place the spacecraft in the desired Earth-Mars transfer orbit. A launch '''period''' is different from a launch '''window''' which is a specific time that a launch can take place on a particular day in the launch period. There are many launch windows in a launch period. Sometimes the phrase '''launch opportunity''' is used to refer to the specific year in which a launch period takes place.

Launch periods are generally constrained by the power of the launch vehicle whereas launch windows are also constrained by launch geometry. One way to visualize an acceptable launch period for various values of delta-v is a porkchop plot, which plots contours of constant delta-V on top of launch dates and landing dates. Pick a value of delta-V, and then use that contour to determine the launch period by observing the earliest launch date and latest launch date.

The gap in the porkchop plot is caused by non-planar delta-v in the transfer burns. Since the Earth and Mars orbit in slightly different planes, the most expensive time to launch is when the earth and Mars are at points where their planes are separated by the maximum amount. Conversely, the cheapest time to launch is when their planes intersect.

One drawback of porkchop plots is that they are only for single-arc transfers, which is why they have such large gaps as a result of launch and arrival plane changes. A different transfer trajectory could be constructed which uses a mid-course plane change maneuver at the intersection of the Earth and Mars orbital planes. However, most launch vehicles would not offer this capability.

There are other real-world considerations which affect the launch period; the ability of MRO to be at the right place at the right time to serve as a relay satellite during entry, descent, and landing constrained the end of Insight's launch period.

==Leaving Earth==

Typically a mission will first launch into a relatively low Earth-centered parking orbit, then it will coast in that orbit for a variable amount of time, and finally the second or third stage of the launch vehicle will inject the spacecraft into a Mars transfer orbit. This injection can either take place all at once, as with the 8 minute burn for [[Curiosity|Curiosity's]] launch, or it can take place over several orbits with gradual apogee-raising maneuvers as in the case of Mars Orbiter Mission.

The parking orbit can be of any inclination. A heavy spacecraft like Curiosity may need to launch into a lower inclination parking orbit so it can take more advantage of the Earth's rotation at launch, however a small spacecraft like Insight might not. Insight launched south from Vandenberg into a polar orbit, coasted for 3/4 of a parking orbit, and the second stage reignited its engines approximately over Alaska to place Insight on the Mars transfer orbit.

At the completion of this Mars transfer insertion burn, the spacecraft will not be exactly on its final course. The spacecraft will be set on a course which intentionally misses Mars so that the non-sterilized launch vehicle does not accidentally hit and contaminate the surface of Mars. After the Mars-transfer orbit burn is completed, the spacecraft will separate from the launch vehicle. The spacecraft will eventually perform a trajectory correction maneuver which will set it on the correct course, but the expended rocket parts will remain on their initial trajectory and miss Mars.

[[Image: GTO-to-Mars.png|thumb|right|px|MEGA scheme for ride-sharing on a GTO launch and then achieving escape to Mars transfer orbit.]]

There are other ways of leaving Earth. One such way is the Moon and Earth Gravity Assist (MEGA) scheme, which has never been performed. In this scheme, a small (~200 kg) spacecraft hitches a ride as a secondary payload on a launch to geostationary transfer orbit (GTO). Once in GTO, the spacecraft will perform a burn at perigee which will raise its apogee enough to take it out beyond the Moon, but not quite escape Earth's gravitational pull. At this far apogee the spacecraft might perform a small burn to target a lunar gravity assist as it comes back towards the Earth. Finally, at the next perigee, the spacecraft will perform its Mars transfer orbit insertion burn.<ref>Paul Penzo, "Mission design for Mars mission using the Ariane ASAP launch capability," 1999. https://trs.jpl.nasa.gov/bitstream/handle/2014/16879/99-0288.pdf?sequence=1</ref>

Another way to escape Earth would be a low-thrust trajectory, where the spacecraft slowly spirals out away from Earth until its eventual escape into heliocentric orbit.

Regardless of the escape scheme, practical considerations impose constraints on the trajectory and timing of maneuvers. For instance, a mission might need to time things so that spacecraft separation occurs in view of a [[Deep Space Network|deep space tracking station]] such as Goldstone.

[[Image: B Plane Targeting.png|thumb|left|px|B Plane used for targeting trajectory correction maneuvers]]

==Targeting Mars==

I'm not sure what is used for targeting the initial transfer orbit insertion burn, but the trajectory correction maneuver burns target a point on the "B-plane" of Mars. The B-plane is defined in a JPL glossary as the "plane perpendicular to the asymptote of the incoming hyperbolic trajectory of the object relative to the Earth."<ref>https://cneos.jpl.nasa.gov/glossary/b_plane.html</ref> I think of it as the "bullseye" plane, which is the plane of the dartboard from the perspective of the person throwing darts. B-plane targeting was originally developed for gravity-assist maneuvers, but it has come to be used for missions where the destination is the planet itself as well.

==topics to elaborate on==
* opposition vs conjunction class transfers
* plane changes
* low-thrust trajectories
* earth orbit part, launch sites, equatorial vs polar parking orbits (or lack of difference between them)
* GTO to Mars transfer scheme
* mars capture schemes: aerobraking, ballistic capture

==References==

* https://blog.adafruit.com/2018/07/26/how-porkchop-plots-determine-earth-to-mars-trajectories-nasa/
* https://trs.jpl.nasa.gov/bitstream/handle/2014/44336/13-0679_A1b.pdf?sequence=1 Ryan C. Woolley and Charles W. Whetsel "On the nature of Earth-Mars porkchop plots" 2014 AAS

Earth-Mars Transfer Trajectory

2018-10-25T03:17:08Z

Pb: /* Leaving Earth */

[[Image: InSight Trajectory.jpg|thumb|right|px|Earth-Mars transfer trajectory of NASA's InSight lander, including planned trajectory correction maneuvers (TCMs)]]

An Earth-Mars transfer trajectory is an orbital path which a spacecraft follows to travel between Earth and Mars. Several types of trajectories have been studied, but all will satisfy the following conditions:

* The starting point must be near the Earth in its orbit around the sun
* The ending point must intersect Mars in its orbit around the sun
* The intervening trajectory must be heliocentric, though one or more gravitational swing-bys of other bodies are allowed

Not all Mars transfer orbits are Hohmann transfers. This is due to the difference in the plane of Earth and Mars's orbit, and can also be due to constraints on launch windows.
There are many variations on this theme, such as whether the spacecraft ends in a Mars-centered orbit or if the spacecraft directly enters the atmosphere from the heliocentric transfer orbit. Other less common variations include the starting orbit at Earth, which could be a low Earth orbit or could instead the mission could begin from geostationary transfer orbit (GTO) as in the MEGA proposals of the late 90's, or one could vary the heliocentric portion of the flight to include a low-thrust trajectory.

[[Image: Porkchopplot.png|thumb|left|px|Porkchop plot for 2018 launch opportunity <ref>https://trs.jpl.nasa.gov/bitstream/handle/2014/44336/13-0679_A1b.pdf?sequence=1</ref>]]

==Launch Period==

A launch period is a span of days during which a launch vehicle can place the spacecraft in the desired Earth-Mars transfer orbit. A launch '''period''' is different from a launch '''window''' which is a specific time that a launch can take place on a particular day in the launch period. There are many launch windows in a launch period. Sometimes the phrase '''launch opportunity''' is used to refer to the specific year in which a launch period takes place.

Launch periods are generally constrained by the power of the launch vehicle whereas launch windows are also constrained by launch geometry. One way to visualize an acceptable launch period for various values of delta-v is a porkchop plot, which plots contours of constant delta-V on top of launch dates and landing dates. Pick a value of delta-V, and then use that contour to determine the launch period by observing the earliest launch date and latest launch date.

The gap in the porkchop plot is caused by non-planar delta-v in the transfer burns. Since the Earth and Mars orbit in slightly different planes, the most expensive time to launch is when the earth and Mars are at points where their planes are separated by the maximum amount. Conversely, the cheapest time to launch is when their planes intersect.

One drawback of porkchop plots is that they are only for single-arc transfers, which is why they have such large gaps as a result of launch and arrival plane changes. A different transfer trajectory could be constructed which uses a mid-course plane change maneuver at the intersection of the Earth and Mars orbital planes. However, most launch vehicles would not offer this capability.

There are other real-world considerations which affect the launch period; the ability of MRO to be at the right place at the right time to serve as a relay satellite during entry, descent, and landing constrained the end of Insight's launch period.

==Leaving Earth==

Typically a mission will launch from Earth into a relatively low Earth-centered parking orbit, then it will coast in that orbit for a variable amount of time, and finally the second or third stage of the launch vehicle will inject the spacecraft into a Mars transfer orbit. This injection can either take place all at once, as with the 8 minute burn for [[Curiosity|Curiosity's]] launch, or it can take place over several orbits with gradual perigee-raising maneuvers as in the case of Mars Orbiter Mission.

The parking orbit can be of any inclination. A heavy spacecraft like Curiosity may need to launch into a lower inclination parking orbit so it can take more advantage of the Earth's rotation at launch, however a small spacecraft like Insight might not. Insight launched south from Vandenberg into a polar orbit, coasted for 3/4 of a parking orbit, and the second stage reignited its engines approximately over Alaska to place Insight on the Mars transfer orbit.

At the completion of this Mars transfer insertion burn, the spacecraft will not be exactly on its final course. The spacecraft will be set on a course which intentionally misses Mars so that the non-sterilized launch vehicle does not accidentally hit and contaminate the surface of Mars. After the Mars-transfer orbit burn is completed, the spacecraft will separate from the launch vehicle. The spacecraft will eventually perform a trajectory correction maneuver which will set it on the correct course, but the expended rocket parts will remain on their initial trajectory and miss Mars.

[[Image: GTO-to-Mars.png|thumb|right|px|MEGA scheme for ride-sharing on a GTO launch and then achieving escape to Mars transfer orbit.]]

There are other ways of leaving Earth. One such way is the Moon and Earth Gravity Assist (MEGA) scheme, which has never been performed. In this scheme, a small (~200 kg) spacecraft hitches a ride as a secondary payload on a launch to geostationary transfer orbit (GTO). Once in GTO, the spacecraft will perform a burn at perigee which will raise its apogee enough to take it out beyond the Moon, but not quite escape Earth's gravitational pull. At this far apogee the spacecraft might perform a small burn to target a lunar gravity assist as it comes back towards the Earth. Finally, at the next perigee, the spacecraft will perform its Mars transfer orbit insertion burn.<ref>Paul Penzo, "Mission design for Mars mission using the Ariane ASAP launch capability," 1999. https://trs.jpl.nasa.gov/bitstream/handle/2014/16879/99-0288.pdf?sequence=1</ref>

Another way to escape Earth would be a low-thrust trajectory, where the spacecraft slowly spirals out away from Earth until its eventual escape into heliocentric orbit.

Regardless of the escape scheme, practical considerations impose constraints on the trajectory and timing of maneuvers. For instance, one may need to launch into a parking orbit such that they can perform their final transfer orbit injection burn and time spacecraft separation to occur in view of a [[Deep Space Network|deep space tracking station]] such as Goldstone.

[[Image: B Plane Targeting.png|thumb|left|px|B Plane used for targeting trajectory correction maneuvers]]

==Targeting Mars==

I'm not sure what is used for targeting the initial transfer orbit insertion burn, but the trajectory correction maneuver burns target a point on the "B-plane" of Mars. The B-plane is defined in a JPL glossary as the "plane perpendicular to the asymptote of the incoming hyperbolic trajectory of the object relative to the Earth."<ref>https://cneos.jpl.nasa.gov/glossary/b_plane.html</ref> I think of it as the "bullseye" plane, which is the plane of the dartboard from the perspective of the person throwing darts. B-plane targeting was originally developed for gravity-assist maneuvers, but it has come to be used for missions where the destination is the planet itself as well.

==topics to elaborate on==
* opposition vs conjunction class transfers
* plane changes
* low-thrust trajectories
* earth orbit part, launch sites, equatorial vs polar parking orbits (or lack of difference between them)
* GTO to Mars transfer scheme
* mars capture schemes: aerobraking, ballistic capture

==References==

* https://blog.adafruit.com/2018/07/26/how-porkchop-plots-determine-earth-to-mars-trajectories-nasa/
* https://trs.jpl.nasa.gov/bitstream/handle/2014/44336/13-0679_A1b.pdf?sequence=1 Ryan C. Woolley and Charles W. Whetsel "On the nature of Earth-Mars porkchop plots" 2014 AAS

File:GTO-to-Mars.png

2018-10-25T03:04:28Z

Pb: typo

== Summary ==
GTO to Mars using the 2-burn Moon-Earth gravity assist. From "Mission design for Mars missions using the Ariane ASAP launch capability," by Paul Penzo, 1999.

https://trs.jpl.nasa.gov/bitstream/handle/2014/16879/99-0288.pdf?sequence=1

Earth-Mars Transfer Trajectory

2018-10-25T03:03:27Z

Pb: escaping earth

[[Image: InSight Trajectory.jpg|thumb|right|px|Earth-Mars transfer trajectory of NASA's InSight lander, including planned trajectory correction maneuvers (TCMs)]]

An Earth-Mars transfer trajectory is an orbital path which a spacecraft follows to travel between Earth and Mars. Several types of trajectories have been studied, but all will satisfy the following conditions:

* The starting point must be near the Earth in its orbit around the sun
* The ending point must intersect Mars in its orbit around the sun
* The intervening trajectory must be heliocentric, though one or more gravitational swing-bys of other bodies are allowed

Not all Mars transfer orbits are Hohmann transfers. This is due to the difference in the plane of Earth and Mars's orbit, and can also be due to constraints on launch windows.
There are many variations on this theme, such as whether the spacecraft ends in a Mars-centered orbit or if the spacecraft directly enters the atmosphere from the heliocentric transfer orbit. Other less common variations include the starting orbit at Earth, which could be a low Earth orbit or could instead the mission could begin from geostationary transfer orbit (GTO) as in the MEGA proposals of the late 90's, or one could vary the heliocentric portion of the flight to include a low-thrust trajectory.

[[Image: Porkchopplot.png|thumb|left|px|Porkchop plot for 2018 launch opportunity <ref>https://trs.jpl.nasa.gov/bitstream/handle/2014/44336/13-0679_A1b.pdf?sequence=1</ref>]]

==Launch Period==

A launch period is a span of days during which a launch vehicle can place the spacecraft in the desired Earth-Mars transfer orbit. A launch '''period''' is different from a launch '''window''' which is a specific time that a launch can take place on a particular day in the launch period. There are many launch windows in a launch period. Sometimes the phrase '''launch opportunity''' is used to refer to the specific year in which a launch period takes place.

Launch periods are generally constrained by the power of the launch vehicle whereas launch windows are also constrained by launch geometry. One way to visualize an acceptable launch period for various values of delta-v is a porkchop plot, which plots contours of constant delta-V on top of launch dates and landing dates. Pick a value of delta-V, and then use that contour to determine the launch period by observing the earliest launch date and latest launch date.

The gap in the porkchop plot is caused by non-planar delta-v in the transfer burns. Since the Earth and Mars orbit in slightly different planes, the most expensive time to launch is when the earth and Mars are at points where their planes are separated by the maximum amount. Conversely, the cheapest time to launch is when their planes intersect.

One drawback of porkchop plots is that they are only for single-arc transfers, which is why they have such large gaps as a result of launch and arrival plane changes. A different transfer trajectory could be constructed which uses a mid-course plane change maneuver at the intersection of the Earth and Mars orbital planes. However, most launch vehicles would not offer this capability.

There are other real-world considerations which affect the launch period; the ability of MRO to be at the right place at the right time to serve as a relay satellite during entry, descent, and landing constrained the end of Insight's launch period.

==Leaving Earth==

Typically a mission will launch from Earth into a relatively low Earth-centered parking orbit, then it will coast in that orbit for a variable amount of time, and finally the second or third stage of the launch vehicle will inject the spacecraft into a Mars transfer orbit. This injection can either take place all at once, as with the 8 minute burn for Curiosity's launch, or it can take place over several orbits with gradual perigee-raising maneuvers as in the case of Mars Orbiter Mission.

The parking orbit can be of any inclination. A heavy spacecraft like Curiosity may need to launch into a lower inclination parking orbit so it can take more advantage of the Earth's rotation at launch, however a small spacecraft like Insight might not. Insight launched south from Vandenberg into a polar orbit, coasted for 3/4 of a parking orbit, and the second stage reignited its engines approximately over Alaska to place Insight on the Mars transfer orbit.

At the completion of this Mars transfer insertion burn, the spacecraft will not be exactly on its final course. The spacecraft will be set on a course which intentionally misses Mars so that the non-sterilized launch vehicle does not accidentally hit and contaminate the surface of Mars. After the Mars-transfer orbit burn is completed, the spacecraft will separate from the launch vehicle. The spacecraft will eventually perform a trajectory correction maneuver which will set it on the correct course, but the expended rocket parts will remain on their initial trajectory and miss Mars.

[[Image: GTO-to-Mars.png|thumb|right|px|MEGA scheme for ride-sharing on a GTO launch and then achieving escape to Mars transfer orbit.]]

There are other ways of leaving Earth. One such way is the Moon and Earth Gravity Assist (MEGA) scheme, which has never been performed. In this scheme, a small (~200 kg) spacecraft hitches a ride as a secondary payload on a launch to geostationary transfer orbit (GTO). Once in GTO, the spacecraft will perform a burn at perigee which will raise its apogee enough to take it out beyond the Moon, but not quite escape Earth's gravitational pull. At this far apogee the spacecraft might perform a small burn to target a lunar gravity assist as it comes back towards the Earth. Finally, at the next perigee, the spacecraft will perform its Mars transfer orbit insertion burn.<ref>Paul Penzo, "Mission design for Mars mission using the Ariane ASAP launch capability," 1999. https://trs.jpl.nasa.gov/bitstream/handle/2014/16879/99-0288.pdf?sequence=1</ref>

Another way to escape Earth would be a low-thrust trajectory, where the spacecraft slowly spirals out away from Earth until its eventual escape into heliocentric orbit.

Regardless of the escape scheme, practical considerations imposing constraints on the timing of maneuvers. For instance, one may need to launch into a parking orbit such that they can perform their final transfer orbit injection burn in view of a tracking station such as Goldstone.

[[Image: B Plane Targeting.png|thumb|left|px|B Plane used for targeting trajectory correction maneuvers]]

==Targeting Mars==

I'm not sure what is used for targeting the initial transfer orbit insertion burn, but the trajectory correction maneuver burns target a point on the "B-plane" of Mars. The B-plane is defined in a JPL glossary as the "plane perpendicular to the asymptote of the incoming hyperbolic trajectory of the object relative to the Earth."<ref>https://cneos.jpl.nasa.gov/glossary/b_plane.html</ref> I think of it as the "bullseye" plane, which is the plane of the dartboard from the perspective of the person throwing darts. B-plane targeting was originally developed for gravity-assist maneuvers, but it has come to be used for missions where the destination is the planet itself as well.

==topics to elaborate on==
* opposition vs conjunction class transfers
* plane changes
* low-thrust trajectories
* earth orbit part, launch sites, equatorial vs polar parking orbits (or lack of difference between them)
* GTO to Mars transfer scheme
* mars capture schemes: aerobraking, ballistic capture

==References==

* https://blog.adafruit.com/2018/07/26/how-porkchop-plots-determine-earth-to-mars-trajectories-nasa/
* https://trs.jpl.nasa.gov/bitstream/handle/2014/44336/13-0679_A1b.pdf?sequence=1 Ryan C. Woolley and Charles W. Whetsel "On the nature of Earth-Mars porkchop plots" 2014 AAS

File:GTO-to-Mars.png

2018-10-25T02:51:36Z

Pb: GOT to Mars using the 2-burn Moon-Earth gravity assist. From "Mission design for Mars mission using the Ariane ASAP launch capability," by Paul Penzo, 1999. https://trs.jpl.nasa.gov/bitstream/handle/2014/16879/99-0288.pdf?sequence=1

== Summary ==
GOT to Mars using the 2-burn Moon-Earth gravity assist. From "Mission design for Mars mission using the Ariane ASAP launch capability," by Paul Penzo, 1999.

https://trs.jpl.nasa.gov/bitstream/handle/2014/16879/99-0288.pdf?sequence=1

Deep Space Network

2018-10-21T16:32:25Z

Pb: /* Alternatives */

The '''Deep Space Network''' (DSN) is an international network of antennas owned and operated by NASA located on [[Earth]] and dedicated to the communication with interplanetary missions and astronomical radio observations. Currently, there are three sites based approximately 120° apart in the Mojave Desert, California (USA), Madrid (Spain) and Canberra (Australia). The dispersion of antennas allow mission controllers to keep constant watch on all interplanetary missions regardless on Earth's rotation.

This network may be used for an [[internet|interplanetary internet link]].

==Use for Communication==
Typically, one or more of the 34 meter diameter antennas is used to communicate with spacecraft at Mars. The data rate of the downlink from Mars can be anywhere from a few bits/second to on the order of 1 Mpbs. Most missions use the DSN at X-band, but a small number of missions are using S-band or Ka-band.

==Use for Navigation==
The DSN can provide ranging and doppler information for use in orbit determination of spacecraft. Long two-way doppler tracking is commonly used.<ref>https://descanso.jpl.nasa.gov/monograph/series1/Descanso1_C03.pdf</ref>

==Downtime==

The DSN requires a line of sight between the antenna and the vehicle, which imposes obvious constraints on the system. For rovers and landers, in order to use a Direct to Earth (DTE) connection, the vehicle must be on the Earth-facing side of the planet, which means that DTE communication is only possible for about half of the Martian day. DTE connections for rovers and landers are uncommon, and instead the surface asset will use a UHF proximity link to transmit data to an orbiter, which relays the signal to DSN over X-band or Ka-band.

There are six orbiters around Mars. Orbiters are occulted when their orbit takes them behind Mars from the viewpoint of the Earth station.

Interference from the solar radiation can cause loss of signal when the angle between the Earth, Sun, and Mars is less than 5°. This loss of signal occurs during a superior conjunction, which occurs once per 26 month synodic period, and can last from a day to over a week depending on the relative geometry due to the non-co-planar orbits of Earth and Mars.

==Loading Considerations==

The Deep Space Network supports approximately 35 missions as of 2018, and significant planning is required to schedule time on the DSN for all space assets that need it. Mars vehicles represent only a fraction of the total users, but when non-Mars missions appear close in the sky to Mars there can be a significant bottleneck. One technique that is used to relieve strain on the system imposed by the many Mars missions is Multiple Spacecraft per Antenna (MSPA). With this technique, one 34 meter antenna can be used to simultaneously downlink data from up to four spacecraft, and uplink to one spacecraft.<ref>https://arc.aiaa.org/doi/pdf/10.2514/6.2018-2502</ref>

==Alternatives==

The Sardinia Radio Telescope, a 64 meter aperture steerable antenna in Italy, can be used for deep space communications. NASA may potentially be able to use this if needed.

The former Soviet Union constructed a number of large aperture antennas for deep-space communication to support their Mars missions. More recently, this network was to be used to support the 2011 Fobos-Grunt mission, but the mission failed before reaching deep space.

China has several large antennas. Most are located in mainland China, but a new groundstation was recently constructed in Argentina. The network was used to track Chang'e 2, which had a mission profile similar to NASA's Clementine mission, with operations in cis-lunar space followed by operations around an asteroid.

The American company Atlas Space Operations is building a commercial capability which they claim will be equivalent to one of the DSN's 70 meter dishes. Atlas LINKS (TM) will be an array of small antennas, not a large aperture antenna like those the DSN utilizes.

==Open issues==
*How long is the gap and how often does it occur?

==External links==
*[http://deepspace.jpl.nasa.gov/dsn/ NASA site for the DSN]

{{stub}}

[[category:communication]][[category:Mission_Control_Centers]][[Category:Technology]]

Deep Space Network

2018-10-21T14:27:49Z

Pb: updating data rate

The '''Deep Space Network''' (DSN) is an international network of antennas owned and operated by NASA located on [[Earth]] and dedicated to the communication with interplanetary missions and astronomical radio observations. Currently, there are three sites based approximately 120° apart in the Mojave Desert, California (USA), Madrid (Spain) and Canberra (Australia). The dispersion of antennas allow mission controllers to keep constant watch on all interplanetary missions regardless on Earth's rotation.

This network may be used for an [[internet|interplanetary internet link]].

==Use for Communication==
Typically, one or more of the 34 meter diameter antennas is used to communicate with spacecraft at Mars. The data rate of the downlink from Mars can be anywhere from a few bits/second to on the order of 1 Mpbs. Most missions use the DSN at X-band, but a small number of missions are using S-band or Ka-band.

==Use for Navigation==
The DSN can provide ranging and doppler information for use in orbit determination of spacecraft. Long two-way doppler tracking is commonly used.<ref>https://descanso.jpl.nasa.gov/monograph/series1/Descanso1_C03.pdf</ref>

==Downtime==

The DSN requires a line of sight between the antenna and the vehicle, which imposes obvious constraints on the system. For rovers and landers, in order to use a Direct to Earth (DTE) connection, the vehicle must be on the Earth-facing side of the planet, which means that DTE communication is only possible for about half of the Martian day. DTE connections for rovers and landers are uncommon, and instead the surface asset will use a UHF proximity link to transmit data to an orbiter, which relays the signal to DSN over X-band or Ka-band.

There are six orbiters around Mars. Orbiters are occulted when their orbit takes them behind Mars from the viewpoint of the Earth station.

Interference from the solar radiation can cause loss of signal when the angle between the Earth, Sun, and Mars is less than 5°. This loss of signal occurs during a superior conjunction, which occurs once per 26 month synodic period, and can last from a day to over a week depending on the relative geometry due to the non-co-planar orbits of Earth and Mars.

==Loading Considerations==

The Deep Space Network supports approximately 35 missions as of 2018, and significant planning is required to schedule time on the DSN for all space assets that need it. Mars vehicles represent only a fraction of the total users, but when non-Mars missions appear close in the sky to Mars there can be a significant bottleneck. One technique that is used to relieve strain on the system imposed by the many Mars missions is Multiple Spacecraft per Antenna (MSPA). With this technique, one 34 meter antenna can be used to simultaneously downlink data from up to four spacecraft, and uplink to one spacecraft.<ref>https://arc.aiaa.org/doi/pdf/10.2514/6.2018-2502</ref>

==Alternatives==

The Sardinia Radio Telescope, a 64 meter aperture steerable antenna in Italy, can be used for deep space communications. NASA may potentially be able to use this if needed.

The former Soviet Union constructed a number of large aperture antennas for deep-space communication to support their Mars missions. More recently, this network was to be used to support the 2011 Fobos-Grunt mission, but the mission failed before reaching deep space.

China has several large antennas. Most are located in mainland China, but a new groundstation was recently constructed in Argentina. The network was used to track Chang'e 2, which had a mission profile similar to NASA's Clementine mission, with operations in cis-lunar space followed by operations around an asteroid.

The American company Atlas Space Operations is building a commercial capability which they claim will be equivalent to one of the DSN's 70 meter dishes. Atlas LINKS (TM) will be an array of small antennas, not a large beam waveguide antenna like those the DSN utilizes.

==Open issues==
*How long is the gap and how often does it occur?

==External links==
*[http://deepspace.jpl.nasa.gov/dsn/ NASA site for the DSN]

{{stub}}

[[category:communication]][[category:Mission_Control_Centers]][[Category:Technology]]

Deep Space Network

2018-10-21T14:01:52Z

Pb: expanding with some details

The '''Deep Space Network''' (DSN) is an international network of antennas owned and operated by NASA located on [[Earth]] and dedicated to the communication with interplanetary missions and astronomical radio observations. Currently, there are three sites based approximately 120° apart in the Mojave Desert, California (USA), Madrid (Spain) and Canberra (Australia). The dispersion of antennas allow mission controllers to keep constant watch on all interplanetary missions regardless on Earth's rotation.

This network may be used for an [[internet|interplanetary internet link]].

==Use for Communication==
Typically, one or more of the 34 meter diameter antennas is used to communicate with spacecraft at Mars. The data rate of the downlink from Mars can be anywhere from a few bits/second to on the order of a hundred kpbs. Most missions use the DSN at X-band, but a small number of missions are using S-band or Ka-band.

==Use for Navigation==
The DSN can provide ranging and doppler information for use in orbit determination of spacecraft. Long two-way doppler tracking is commonly used.<ref>https://descanso.jpl.nasa.gov/monograph/series1/Descanso1_C03.pdf</ref>

==Downtime==

The DSN requires a line of sight between the antenna and the vehicle, which imposes obvious constraints on the system. For rovers and landers, in order to use a Direct to Earth (DTE) connection, the vehicle must be on the Earth-facing side of the planet, which means that DTE communication is only possible for about half of the Martian day. DTE connections for rovers and landers are uncommon, and instead the surface asset will use a UHF proximity link to transmit data to an orbiter, which relays the signal to DSN over X-band or Ka-band.

There are six orbiters around Mars. Orbiters are occulted when their orbit takes them behind Mars from the viewpoint of the Earth station.

Interference from the solar radiation can cause loss of signal when the angle between the Earth, Sun, and Mars is less than 5°. This loss of signal occurs during a superior conjunction, which occurs once per 26 month synodic period, and can last from a day to over a week depending on the relative geometry due to the non-co-planar orbits of Earth and Mars.

==Loading Considerations==

The Deep Space Network supports approximately 35 missions as of 2018, and significant planning is required to schedule time on the DSN for all space assets that need it. Mars vehicles represent only a fraction of the total users, but when non-Mars missions appear close in the sky to Mars there can be a significant bottleneck. One technique that is used to relieve strain on the system imposed by the many Mars missions is Multiple Spacecraft per Antenna (MSPA). With this technique, one 34 meter antenna can be used to simultaneously downlink data from up to four spacecraft, and uplink to one spacecraft.<ref>https://arc.aiaa.org/doi/pdf/10.2514/6.2018-2502</ref>

==Alternatives==

The Sardinia Radio Telescope, a 64 meter aperture steerable antenna in Italy, can be used for deep space communications. NASA may potentially be able to use this if needed.

The former Soviet Union constructed a number of large aperture antennas for deep-space communication to support their Mars missions. More recently, this network was to be used to support the 2011 Fobos-Grunt mission, but the mission failed before reaching deep space.

China has several large antennas. Most are located in mainland China, but a new groundstation was recently constructed in Argentina. The network was used to track Chang'e 2, which had a mission profile similar to NASA's Clementine mission, with operations in cis-lunar space followed by operations around an asteroid.

The American company Atlas Space Operations is building a commercial capability which they claim will be equivalent to one of the DSN's 70 meter dishes. Atlas LINKS (TM) will be an array of small antennas, not a large beam waveguide antenna like those the DSN utilizes.

==Open issues==
*How long is the gap and how often does it occur?

==External links==
*[http://deepspace.jpl.nasa.gov/dsn/ NASA site for the DSN]

{{stub}}

[[category:communication]][[category:Mission_Control_Centers]][[Category:Technology]]

List of Planned Missions

2018-10-21T12:55:15Z

Pb: adding acronyms

==Planned Missions==

These missions have a high likelihood of making it to the launch pad based on funding levels, program history, and the track record of the selected launch vehicle.

{| class="wikitable"
|-
! Mission
! Category
! Planned Launch
! Launch Vehicle
! Organization
|-
| ExoMars Rover and Surface Platform (RSP)
| [[:category:landers|Lander]], [[:category:rovers|Rover]]
| 2020
| Proton
| ESA
|-
| Mars 2020 Rover (M20)
| [[:category:rovers|Rover]]
| 2020
| Atlas V 541
| NASA
|-
| Mars Hope / Emirates Mars Mission (EMM)
| [[:category:orbiters|Orbiter]]
| 2020
| H-IIA
| UAE
|-
| Mars Global Remote Sensing Orbiter and Small Rover
| [[:category:orbiters|Orbiter]], [[:category:rovers|Rover]]
| 2020
| Long March 5
| CNSA
|-
| Mars Orbiter Mission 2 (MOM2)
| [[:category:orbiters|Orbiter]]
| 2022
| GSLV Mark III
| ISRO
|-
| Martian Moons Exploration Phobos Lander with Sample Return (MMX)
| [[:category:landers|Lander]]
| 2024
| H3-24L
| JAXA
|}

==Secondary Payloads on Planned Missions==

As of October 2018, there have been no formal announcements of any secondary payloads launching with any of the above primary missions. There has been word of a [https://en.wikipedia.org/wiki/Mars_Terahertz_Microsatellite Japanese microsatellite] launching as a secondary payload in 2020, but no details have emerged.

==Proposed Missions==

These missions have been proposed but lack in either certainty concerning the launch date, or certainty of launching at all. They are arranged in no particular order (i.e. the first is not necessarily the most likely to launch first).

* SpaceX first BFR cargo mission in 2022, followed by crew and cargo in 2024 (as stated on their website in October 2018).
* NASA's "Next Mars Orbiter" (NeMO) in the late 2020's. This was originally planned for 2022 but has taken a back seat in favor of an "accelerated" sample return budgeting scheme
* NASA's Mars sample return mission. The details of this mission are to be decided in 2019.
* Mars One still exists, though support and confidence have waned
* NASA still presents Mars as a "horizon goal," though according to a report delivered to the United States Congress in September of 2018, no decision will be made concerning the architecture of this mission until 2024. Certain members of Congress, notably Rep. Perlmutter of Colorado, advocate strongly for a 2033 date.

[[Category:Missions]]

Earth-Mars Transfer Trajectory

2018-10-09T00:30:08Z

Pb: adding launch period section

File:Porkchopplot.png

2018-10-09T00:19:06Z

Pb: Example porkchop plot for Earth-Mars transfer in the 2018 window. https://trs.jpl.nasa.gov/bitstream/handle/2014/44336/13-0679_A1b.pdf?sequence=1 Ryan C. Woolley and Charles W. Whetsel "On the nature of Earth-Mars porkchop plots" 2014 AAS

== Summary ==
Example porkchop plot for Earth-Mars transfer in the 2018 window.

https://trs.jpl.nasa.gov/bitstream/handle/2014/44336/13-0679_A1b.pdf?sequence=1

Ryan C. Woolley and Charles W. Whetsel "On the nature of Earth-Mars porkchop plots" 2014 AAS

Earth-Mars Transfer Trajectory

2018-10-03T14:27:16Z

Pb: work in progress

An Earth-Mars transfer trajectory is an orbital path which a spacecraft follows to travel between Earth and Mars. Several types of trajectories have been studied, but all must satisfy the following conditions:

[[Image: InSight Trajectory.jpg|thumb|right|px|Earth-Mars transfer trajectory of NASA's InSight lander, including planned trajectory correction maneuvers (TCMs)]]

* The starting point must intersect the Earth in its orbit around the sun
* The ending point must intersect Mars in its orbit around the sun
* The intervening trajectory must be heliocentric, though one or more gravitational swing-bys of other bodies are allowed

There are many variations on this theme, such as whether the spacecraft ends in a Mars-centered orbit or if the spacecraft directly enters the atmosphere from the heliocentric transfer orbit. Other less common variations include the starting orbit at Earth, which could be a low Earth orbit or could instead the mission could begin from geostationary transfer orbit (GTO) as in the MEGA proposals of the late 90's, or one could vary the heliocentric portion of the flight to include a low-thrust trajectory.

==Targeting==

I'm not sure what is used for targeting the initial transfer orbit insertion burn, but the trajectory correction maneuver burns target a point on the "B-plane" of Mars. The B-plane is defined by the NASA glossary as the "plane perpendicular to the asymptote of the incoming hyperbolic trajectory of the object relative to the Earth."<ref>https://cneos.jpl.nasa.gov/glossary/b_plane.html</ref> I think of it as the "bullseye" plane, which is the plane of the dartboard from the perspective of the person throwing darts. B-plane targeting was originally developed for gravity-assist maneuvers, but it has come to be used for missions where the destination is the planet itself as well.

[[Image: B Plane Targeting.png|thumb|left|px|B Plane used for targeting trajectory correction maneuvers]]

==topics to elaborate on==
* opposition vs conjunction class transfers
* really nail down b-plane stuff with diagrams
* porkchop plots
* launch windows
* plane changes
* low-thrust trajectories
* earth orbit part, launch sites, equatorial vs polar parking orbits (or lack of difference between them)
* GTO to Mars transfer scheme
* mars capture schemes: aerobraking, ballistic capture

==References==

File:B Plane Targeting.png

2018-10-03T14:24:37Z

Pb: Diagram of the "B-Plane" used in targeting for interplanetary maneuvers including gravity assists, orbital insertion, and atmospheric entry. Image courtesy NASA, taken from GMAT tutorial on B-Plane targeting: http://gmat.sourceforge.net/docs/R2018a/he...

== Summary ==
Diagram of the "B-Plane" used in targeting for interplanetary maneuvers including gravity assists, orbital insertion, and atmospheric entry.

Image courtesy NASA, taken from GMAT tutorial on B-Plane targeting: http://gmat.sourceforge.net/docs/R2018a/help.html#Mars_B_Plane_Targeting
== Licensing ==
{{PD}}

Earth-Mars Transfer Trajectory

2018-10-03T13:52:07Z

Pb: adding thumbnail

File:InSight Trajectory.jpg

2018-10-03T13:50:39Z

Pb: The trajectory of NASA's InSight lander, including planned trajectory correction maneuvers (TCMs). Image credit NASA/JPL

== Summary ==
The trajectory of NASA's InSight lander, including planned trajectory correction maneuvers (TCMs).

Image credit NASA/JPL
== Licensing ==
{{PD}}

List of Planned Missions

2018-10-03T03:25:48Z

Pb: category:missions

==Planned Missions==

These missions have a high likelihood of making it to the launch pad based on funding levels, program history, and the track record of the selected launch vehicle.

{| class="wikitable"
|-
! Mission
! Category
! Planned Launch
! Launch Vehicle
! Organization
|-
| ExoMars 2020
| [[:category:landers|Lander]], [[:category:rovers|Rover]]
| 2020
| Proton
| ESA
|-
| Mars 2020
| [[:category:rovers|Rover]]
| 2020
| Atlas V 541
| NASA
|-
| Mars Hope
| [[:category:orbiters|Orbiter]]
| 2020
| H-IIA
| UAE
|-
| Mars Global Remote Sensing Orbiter and Small Rover
| [[:category:orbiters|Orbiter]], [[:category:rovers|Rover]]
| 2020
| Long March 5
| CNSA
|-
| Mars Orbiter Mission 2
| [[:category:orbiters|Orbiter]]
| 2022
| GSLV Mark III
| ISRO
|-
| Martian Moons Exploration Phobos Lander with Sample Return
| [[:category:landers|Lander]]
| 2024
| H3-24L
| JAXA
|}

==Secondary Payloads on Planned Missions==

As of October 2018, there have been no formal announcements of any secondary payloads launching with any of the above primary missions. There has been word of a [https://en.wikipedia.org/wiki/Mars_Terahertz_Microsatellite Japanese microsatellite] launching as a secondary payload in 2020, but no details have emerged.

==Proposed Missions==

These missions have been proposed but lack in either certainty concerning the launch date, or certainty of launching at all. They are arranged in no particular order (i.e. the first is not necessarily the most likely to launch first).

* SpaceX first BFR cargo mission in 2022, followed by crew and cargo in 2024 (as stated on their website in October 2018).
* NASA's "Next Mars Orbiter" (NeMO) in the late 2020's. This was originally planned for 2022 but has taken a back seat in favor of an "accelerated" sample return budgeting scheme
* NASA's Mars sample return mission. The details of this mission are to be decided in 2019.
* Mars One still exists, though support and confidence have waned
* NASA still presents Mars as a "horizon goal," though according to a report delivered to the United States Congress in September of 2018, no decision will be made concerning the architecture of this mission until 2024. Certain members of Congress, notably Rep. Perlmutter of Colorado, advocate strongly for a 2033 date.

[[Category:Missions]]

List of current missions

2018-10-03T03:24:49Z

Pb: link

{{Stub}}

'''Missions currently ''en-route'' to, landed on, or orbiting the planet are listed below.'''

{|
|style="width:200px"|'''Mission'''
|style="width:70px"|'''Type'''
|style="width:100px"|'''Launch date'''
|style="width:100px"|'''Arrival date'''
|style="width:100px"|'''Principal organization'''
|style="width:100px"|'''Location'''
|style="width:200px"|'''Status'''
|-
|'''[[Mars Odyssey|2001 Mars Odyssey]]'''
|[[:category:orbiters|Orbiter]]
|7 April 2001
|24 October 2001
|[[NASA]]
|Polar orbit
|Propellant could last until 2025
|-
|'''[[Mars Express]]'''
|[[:category:orbiters|Orbiter]]
|2 June 2003
|25 December 2003
|[[ESA]]
|Elliptical near-polar orbit
|Funded through 2020
|-
|'''[[Opportunity]]'''
|[[:category:rovers|Rover]]
|7 July 2003
|25 January 2004
|[[NASA]]
|Endeavour Crater
|Hibernation - dust storm
|-
|'''[[Mars Reconnaissance Orbiter]]'''
|[[:category:orbiters|Orbiter]]
|12 August 2005
|10 March 2006
|[[NASA]]
|Sun-synchronous (polar) orbit
|Propellant could last to 2030's
|-
|'''[[Curiosity]]'''
|[[:category:landers|Lander]]
|26 November 2011
|''6 August 2012''
|[[NASA]]
|Gale Crater
|Active
|-
|'''[[Mars Orbiter Mission]]'''
|[[:category:orbiters|Orbiter]]
|5 November 2013
|''24 Septemer 2014''
|[[ISAC]]
|Highly elliptical non-polar orbit
|Active
|-
|'''[[MAVEN]]'''
|[[:category:orbiters|Orbiter]]
|18 November 2013
|''22 September 2014''
|[[NASA]]
|Highly elliptical non-polar orbit
|Currently in science mission, planned orbit change in 2019 to serve as comm relay
|-
|'''[[ExoMars Trace Gas Orbiter]]'''
|[[:category:orbiters|Orbiter]]
|14 March 2016
|''19 October 2016''
|[[ESA/Roscosmos]]
|Circular 74 deg science orbit
|Active
|-
|'''[[InSight]]'''
|[[:category:landers|Lander]]
|5 May 2018
|''November 2018''
|[[NASA]]
|En-route
|
|}

See also [[List_of_Planned_Missions|List of Planned Missions]]

[[category:Missions]]

Landing on Mars

2018-10-03T03:23:32Z

Pb: link

A typical mission to land on Mars consists of atmospheric entry phase using an aeroshell, a parachute descent phase, a terminal descent phase, and ends with the spacecraft touching down somewhere within a predetermined [[Landing Ellipse|landing ellipse]]. The spacecraft can enter the atmosphere either directly from the [[Earth-Mars_Transfer_Trajectory|Earth-Mars transfer orbit]] or it can descend after first being captured into Mars orbit. The parachute descent phase begins in the supersonic flight regime and normally utilizes ring-sail or disk-gap-band supersonic parachute. The terminal descent phase has been achieved with retrorockets and lanyards/skycranes, and spacecraft have been cushioned on touchdown with foam and airbags.

To date there have been 9 successful soft landings: Mars 3, Vikings 1 and 2, Pathfinder, Beagle 2, MER A and B (Spirit and Opportunity), Phoenix, and MSL (Curiosity). Mars 3 and Beagle 2 have been declared as successful soft landings though the spacecraft did not operate correctly once on the surface. There have been 4 failed landing attempts: Mars 2, Mars 6, Mars Polar Lander, and the Schiaparelli EDM lander.

Wind tunnels, rocket sleds, and drop tests have been used to test equipment for entry, descent, and landing. Work continues on new technology such as supersonic retropropulsion, ballutes, low-density-supersonic-decelerators (and other expandable entry shields), biconic heat shields (and other entry vehicle shapes), and new ablative or non-ablative heat shield materials.

==Concepts==

Marspedia contributors have supplied the following concepts for EDL technology.

If we need a four hundred foot diameter parachute manufactured in space out of aluminum oxide fiber and sent to Mars in stiff deployed condition instead of being packed, we will not learn about it unless we see a need to experiment. Such a parachute might merit investigation. It would avoid opening shock and might be sufficiently heat resistant to maintain structural integrity during the entire descent in Mars' low gravity well. The larger the diameter of the parachute, the less the max g loading. So let us be honest with ourselves about all necessary colonization technology.

The expected max temperature for ballistic entry into Mars atmosphere is expected to be a thousand or more Kelvin degrees above the melting point of aluminum oxide so coating course aluminum oxide fibers with potassium oxide which decomposes at 490 Centigrade might protect the fibers through atmospheric entry by ablative cooling or it might not. A mixture of potassium and sodium oxides as a coating or Teflon as a coating are things that are conceivable. Engineers in this specialty would have a better idea.

=== High Lift Vertical Landing Vehicle ===

Another alternative with a greater probability of working, but possibly high cost, is a delta winged entry vehicle or lifting body with insulation like that on the space shuttle. The insulation would be somewhat cheaper because Mars atmospheric entry is less demanding than Earth reentry. After losing most of its orbital velocity to the atmosphere by heating the atmosphere in passing, this vehicle would fly supersonic close to the ground then ignite its rockets for landing. Then it would perform a Pugachev's Cobra<ref>http://en.wikipedia.org/wiki/Pugachev%27s_Cobra</ref> maneuver losing horizontal velocity by drag and by rocket thrust. It would then touch down on its tail. Rocket thrust directly into the supersonic slipstream of Mars' atmosphere will not work to safely land on Mars because the supersonic slipstream that the lander flies into would carry the noise of the rocket exhaust right back to the lander. The potential for the chaotic forces of this rocket noise to destabilize the lander's orientation and damage its structure rule out this technique. In the Pugachev's Cobra maneuver, rocket thrust is never directed directly into the supersonic slipstream. The rocket thrust always has a vertical component while the slipstream moves horizontally until the slip stream velocity is reduced to a negligible value.

This sort of vehicle might approach the point of entering a Pugachev's Cobra maneuver by flying horizontally near Mars' surface while increasing angle of attack to maintain lift while killing velocity. At a pitch attitude of 45 degrees there is little lift left to be gained by increasing angle of attack. This should occur at about Mach 2.5, which is about 600 meters per second on Mars. Then the rockets are ignited generating two Mars gravities of acceleration and the angle of attack is further increased past 90 degrees to generate negative lift and keep the vehicle in horizontal flight. As the speed decreases and negative lift generated by the wings decreases, the pitch angle is increased to reduce the component of rocket thrust in the vertical direction and increase the component of rocket thrust directed to braking. As the vehicle eventually slows to a stop in horizontal motion, a combination of throttling and thrust deflection reduces thrust to about 1 Mars gravity, the vehicle moves to a 90 degree pitch angle and settles on its tail. A guesstimate of the required rocket delta V for killing the last 600 meters per second and landing in this way is about 850 meters per second. This includes the amount of speed lost to atmospheric drag and very substantial gravity losses.

=== Another Alternative is the Sky Crane ===

<blockquote>the 2009 Mars Science Laboratory (MSL) rover, weighing 775 kilograms (versus MER at 175.4 kilograms each) requires an entirely new landing architecture. Too massive for airbags, the small-car sized rover will use a landing system dubbed the Sky Crane. "Even though some people laugh when they first see it, my personal view is that the Sky Crane is actually the most elegant system we've come up with yet, and the simplest," said Manning. MSL will use a combination of a rocket-guided entry with a heat shield, a parachute, then thrusters to slow the vehicle even more, followed by a crane-like system that lowers the rover on a cable for a soft landing directly on its wheels. Depending on the success of the Sky Crane with MSL, it's likely that this system can be scaled for larger payloads, but probably not the size needed to land humans on Mars. (See Ref #1) </blockquote>

=== A Sure Way to Land on Mars ===
A sure but expensive way to land on Mars with a ten metric ton vehicle is to build a heat shield in orbit around Earth and send it to Mars as part of the spacecraft. After the heat shield slows the spacecraft, rockets bring it to a safe stop on Mars. Since Mars' atmosphere at the surface is one hundredth the density of Earth's atmosphere at the surface, make the heat shield proportionally bigger. Considering that the 12,250 pound Apollo command module was 12.8 feet in diameter, a ten metric ton Mars lander should have a 52 meter diameter heat shield. Assembled from 127 roughly hexagonal pieces about 4 meters in diameter, this would be a hexagonal heat shield instead of a round one. That should do. Each hexagonal piece would have a layer of ablative material on one side of a hexagon of aircraft grade aluminum. Aluminum t cross section extrusions would be fastened to the Aluminum sheet as stiffeners. In orbit, two hexagon sections would have their ablative sections butted against each other, protrusions fitting into cavities. A small gap would remain between the aluminum sheets. A 2 inch strip along the edge of each aluminum sheet would be pre-coated with brazing material. A 4 inch wide strip of aluminum to join them would likewise be coated with brazing material on one side. A ridge on the joining strip would fit in the gap between the aluminum hexagons. Then an iron heated to the right temperature would be placed on the joining strip and left for the right time. When the iron is removed and the piece cools the two hexagons make one piece with brazing material partially filling the gap between the two hexagons and rounding out the corner where the hexagons meet the joining strip. Likewise, a trusswork joining the stiffeners of all of the hexagons would be assembled by the same brazing technique and make the whole heat shield one strong rigid light weight piece. Some work has already been done considering robotic truss assembly on orbit. [http://www.ri.cmu.edu/publication_view.html?pub_id=1691 Light-Weight Mobile Robot for Space Station Trusswork] A cone section for the sphere-cone reentry vehicle would also need to be built or a somewhat different shaped section serving the same aerodynamic function, avoiding excessive, uncontrolled and chaotic side slip.

The advantage of sending up a ten ton vehicle, many pieces of heat shield and a robotic assembly station two make a big heat shield as compared to sending up a vehicle with heat shield and parachutes on an Ares V is that the big assembled-on-orbit heat shield would allow a 10 ton vehicle to land cargo safely on Mars while the Ares V scheme would not land cargo or people safely on Mars. Mars direct would do no better. See [http://www.universetoday.com/2007/07/17/the-mars-landing-approach-getting-large-payloads-to-the-surface-of-the-red-planet/ The mars landing Approach: Getting Large Payloads to the Surface of the Red Planet]

Just as all economic activity in orbit so far has been done by robots, assembling a spacecraft to go to Mars should be done by robots and setting up the infrastructure for people to survive on Mars should be done by robots. There are some technical difficulties with this approach that must be addressed, but they seem likely to be amenable to solution.

=== Terminal Velocity of the Big Heat Shield Vehicle ===
For a rough estimate take as a starting point the estimated terminal velocity for the Apollo reentry vehicle. Estimate that the square of the terminal velocity is directly proportional to mass and the force of gravity and inversely proportional to air density and cross sectional area. The 10 metric ton vehicle with a 52 meter diameter heat shield should then have a terminal velocity less than 56 meters per second (125 miles per hour) at Mars' surface. If near the end of this descent a hole is burst through the bottom of the heat shield in the center right under the payload and hinges swing that portion of the heat shield out of the way, then retro rockets can fire at an altitude of 160 meters for 5.1 seconds with an acceleration of 11 meters per second squared and bring the payload to a stop about 17 meters above the heat shield that crashes into Mars. A few more seconds for horizontal maneuvering and throttling the rocket motor brings the payload safely to Mars on landing legs. A guesstimate of the required rocket delta V for this maneuver is about 68 meters per second.

== References ==
<references/>

[[Category:Missions]]

Earth-Mars Transfer Trajectory

2018-10-03T03:22:12Z

Pb: starting transfer trajectory page. this will take some work.

An Earth-Mars transfer trajectory is an orbital path which a spacecraft follows to travel between Earth and Mars. Several types of trajectories have been studied, but all must satisfy the following conditions:

* The starting point must intersect the Earth in its orbit around the sun
* The ending point must intersect Mars in its orbit around the sun
* The intervening trajectory must be heliocentric, though one or more gravitational swing-bys of other bodies are allowed

There are many variations on this theme, such as whether the spacecraft ends in a Mars-centered orbit or if the spacecraft directly enters the atmosphere from the heliocentric transfer orbit. Other less common variations include the starting orbit at Earth, which could be a low Earth orbit or could instead the mission could begin from geostationary transfer orbit (GTO) as in the MEGA proposals of the late 90's, or one could vary the heliocentric portion of the flight to include a low-thrust trajectory.

List of Planned Missions

2018-10-03T03:05:07Z

Pb: adding section for planned secondary payloads, and a section for proposed missions

==Planned Missions==

These missions have a high likelihood of making it to the launch pad based on funding levels, program history, and the track record of the selected launch vehicle.

{| class="wikitable"
|-
! Mission
! Category
! Planned Launch
! Launch Vehicle
! Organization
|-
| ExoMars 2020
| [[:category:landers|Lander]], [[:category:rovers|Rover]]
| 2020
| Proton
| ESA
|-
| Mars 2020
| [[:category:rovers|Rover]]
| 2020
| Atlas V 541
| NASA
|-
| Mars Hope
| [[:category:orbiters|Orbiter]]
| 2020
| H-IIA
| UAE
|-
| Mars Global Remote Sensing Orbiter and Small Rover
| [[:category:orbiters|Orbiter]], [[:category:rovers|Rover]]
| 2020
| Long March 5
| CNSA
|-
| Mars Orbiter Mission 2
| [[:category:orbiters|Orbiter]]
| 2022
| GSLV Mark III
| ISRO
|-
| Martian Moons Exploration Phobos Lander with Sample Return
| [[:category:landers|Lander]]
| 2024
| H3-24L
| JAXA
|}

==Secondary Payloads on Planned Missions==

As of October 2018, there have been no formal announcements of any secondary payloads launching with any of the above primary missions. There has been word of a [https://en.wikipedia.org/wiki/Mars_Terahertz_Microsatellite Japanese microsatellite] launching as a secondary payload in 2020, but no details have emerged.

==Proposed Missions==

These missions have been proposed but lack in either certainty concerning the launch date, or certainty of launching at all. They are arranged in no particular order (i.e. the first is not necessarily the most likely to launch first).

* SpaceX first BFR cargo mission in 2022, followed by crew and cargo in 2024 (as stated on their website in October 2018).
* NASA's "Next Mars Orbiter" (NeMO) in the late 2020's. This was originally planned for 2022 but has taken a back seat in favor of an "accelerated" sample return budgeting scheme
* NASA's Mars sample return mission. The details of this mission are to be decided in 2019.
* Mars One still exists, though support and confidence have waned
* NASA still presents Mars as a "horizon goal," though according to a report delivered to the United States Congress in September of 2018, no decision will be made concerning the architecture of this mission until 2024. Certain members of Congress, notably Rep. Perlmutter of Colorado, advocate strongly for a 2033 date.

List of Planned Missions

2018-10-03T02:31:14Z

Pb: adding launch vehicles to planned missions

{| class="wikitable"
|-
! Mission
! Category
! Planned Launch
! Launch Vehicle
! Organization
|-
| ExoMars 2020
| [[:category:landers|Lander]], [[:category:rovers|Rover]]
| 2020
| Proton
| ESA
|-
| Mars 2020
| [[:category:rovers|Rover]]
| 2020
| Atlas V 541
| NASA
|-
| Mars Hope
| [[:category:orbiters|Orbiter]]
| 2020
| H-IIA
| UAE
|-
| Mars Global Remote Sensing Orbiter and Small Rover
| [[:category:orbiters|Orbiter]], [[:category:rovers|Rover]]
| 2020
| Long March 5
| CNSA
|-
| Mars Orbiter Mission 2
| [[:category:orbiters|Orbiter]]
| 2022
| GSLV Mark III
| ISRO
|-
| Martian Moons Exploration Phobos Lander with Sample Return
| [[:category:landers|Lander]]
| 2024
| H3-24L
| JAXA
|}

Landing on Mars

2018-09-16T18:30:49Z

Pb: new intro paragraph

A typical mission to land on Mars consists of atmospheric entry phase using an aeroshell, a parachute descent phase, a terminal descent phase, and ends with the spacecraft touching down somewhere within a predetermined [[Landing Ellipse|landing ellipse]]. The spacecraft can enter the atmosphere either directly from the [[Earth-Mars transfer orbit]] or it can descend after first being captured into Mars orbit. The parachute descent phase begins in the supersonic flight regime and normally utilizes ring-sail or disk-gap-band supersonic parachute. The terminal descent phase has been achieved with retrorockets and lanyards/skycranes, and spacecraft have been cushioned on touchdown with foam and airbags.

To date there have been 9 successful soft landings: Mars 3, Vikings 1 and 2, Pathfinder, Beagle 2, MER A and B (Spirit and Opportunity), Phoenix, and MSL (Curiosity). Mars 3 and Beagle 2 have been declared as successful soft landings though the spacecraft did not operate correctly once on the surface. There have been 4 failed landing attempts: Mars 2, Mars 6, Mars Polar Lander, and the Schiaparelli EDM lander.

Wind tunnels, rocket sleds, and drop tests have been used to test equipment for entry, descent, and landing. Work continues on new technology such as supersonic retropropulsion, ballutes, low-density-supersonic-decelerators (and other expandable entry shields), biconic heat shields (and other entry vehicle shapes), and new ablative or non-ablative heat shield materials.

==Concepts==

Marspedia contributors have supplied the following concepts for EDL technology.

If we need a four hundred foot diameter parachute manufactured in space out of aluminum oxide fiber and sent to Mars in stiff deployed condition instead of being packed, we will not learn about it unless we see a need to experiment. Such a parachute might merit investigation. It would avoid opening shock and might be sufficiently heat resistant to maintain structural integrity during the entire descent in Mars' low gravity well. The larger the diameter of the parachute, the less the max g loading. So let us be honest with ourselves about all necessary colonization technology.

The expected max temperature for ballistic entry into Mars atmosphere is expected to be a thousand or more Kelvin degrees above the melting point of aluminum oxide so coating course aluminum oxide fibers with potassium oxide which decomposes at 490 Centigrade might protect the fibers through atmospheric entry by ablative cooling or it might not. A mixture of potassium and sodium oxides as a coating or Teflon as a coating are things that are conceivable. Engineers in this specialty would have a better idea.

=== High Lift Vertical Landing Vehicle ===

Another alternative with a greater probability of working, but possibly high cost, is a delta winged entry vehicle or lifting body with insulation like that on the space shuttle. The insulation would be somewhat cheaper because Mars atmospheric entry is less demanding than Earth reentry. After losing most of its orbital velocity to the atmosphere by heating the atmosphere in passing, this vehicle would fly supersonic close to the ground then ignite its rockets for landing. Then it would perform a Pugachev's Cobra<ref>http://en.wikipedia.org/wiki/Pugachev%27s_Cobra</ref> maneuver losing horizontal velocity by drag and by rocket thrust. It would then touch down on its tail. Rocket thrust directly into the supersonic slipstream of Mars' atmosphere will not work to safely land on Mars because the supersonic slipstream that the lander flies into would carry the noise of the rocket exhaust right back to the lander. The potential for the chaotic forces of this rocket noise to destabilize the lander's orientation and damage its structure rule out this technique. In the Pugachev's Cobra maneuver, rocket thrust is never directed directly into the supersonic slipstream. The rocket thrust always has a vertical component while the slipstream moves horizontally until the slip stream velocity is reduced to a negligible value.

This sort of vehicle might approach the point of entering a Pugachev's Cobra maneuver by flying horizontally near Mars' surface while increasing angle of attack to maintain lift while killing velocity. At a pitch attitude of 45 degrees there is little lift left to be gained by increasing angle of attack. This should occur at about Mach 2.5, which is about 600 meters per second on Mars. Then the rockets are ignited generating two Mars gravities of acceleration and the angle of attack is further increased past 90 degrees to generate negative lift and keep the vehicle in horizontal flight. As the speed decreases and negative lift generated by the wings decreases, the pitch angle is increased to reduce the component of rocket thrust in the vertical direction and increase the component of rocket thrust directed to braking. As the vehicle eventually slows to a stop in horizontal motion, a combination of throttling and thrust deflection reduces thrust to about 1 Mars gravity, the vehicle moves to a 90 degree pitch angle and settles on its tail. A guesstimate of the required rocket delta V for killing the last 600 meters per second and landing in this way is about 850 meters per second. This includes the amount of speed lost to atmospheric drag and very substantial gravity losses.

=== Another Alternative is the Sky Crane ===

<blockquote>the 2009 Mars Science Laboratory (MSL) rover, weighing 775 kilograms (versus MER at 175.4 kilograms each) requires an entirely new landing architecture. Too massive for airbags, the small-car sized rover will use a landing system dubbed the Sky Crane. "Even though some people laugh when they first see it, my personal view is that the Sky Crane is actually the most elegant system we've come up with yet, and the simplest," said Manning. MSL will use a combination of a rocket-guided entry with a heat shield, a parachute, then thrusters to slow the vehicle even more, followed by a crane-like system that lowers the rover on a cable for a soft landing directly on its wheels. Depending on the success of the Sky Crane with MSL, it's likely that this system can be scaled for larger payloads, but probably not the size needed to land humans on Mars. (See Ref #1) </blockquote>

=== A Sure Way to Land on Mars ===
A sure but expensive way to land on Mars with a ten metric ton vehicle is to build a heat shield in orbit around Earth and send it to Mars as part of the spacecraft. After the heat shield slows the spacecraft, rockets bring it to a safe stop on Mars. Since Mars' atmosphere at the surface is one hundredth the density of Earth's atmosphere at the surface, make the heat shield proportionally bigger. Considering that the 12,250 pound Apollo command module was 12.8 feet in diameter, a ten metric ton Mars lander should have a 52 meter diameter heat shield. Assembled from 127 roughly hexagonal pieces about 4 meters in diameter, this would be a hexagonal heat shield instead of a round one. That should do. Each hexagonal piece would have a layer of ablative material on one side of a hexagon of aircraft grade aluminum. Aluminum t cross section extrusions would be fastened to the Aluminum sheet as stiffeners. In orbit, two hexagon sections would have their ablative sections butted against each other, protrusions fitting into cavities. A small gap would remain between the aluminum sheets. A 2 inch strip along the edge of each aluminum sheet would be pre-coated with brazing material. A 4 inch wide strip of aluminum to join them would likewise be coated with brazing material on one side. A ridge on the joining strip would fit in the gap between the aluminum hexagons. Then an iron heated to the right temperature would be placed on the joining strip and left for the right time. When the iron is removed and the piece cools the two hexagons make one piece with brazing material partially filling the gap between the two hexagons and rounding out the corner where the hexagons meet the joining strip. Likewise, a trusswork joining the stiffeners of all of the hexagons would be assembled by the same brazing technique and make the whole heat shield one strong rigid light weight piece. Some work has already been done considering robotic truss assembly on orbit. [http://www.ri.cmu.edu/publication_view.html?pub_id=1691 Light-Weight Mobile Robot for Space Station Trusswork] A cone section for the sphere-cone reentry vehicle would also need to be built or a somewhat different shaped section serving the same aerodynamic function, avoiding excessive, uncontrolled and chaotic side slip.

The advantage of sending up a ten ton vehicle, many pieces of heat shield and a robotic assembly station two make a big heat shield as compared to sending up a vehicle with heat shield and parachutes on an Ares V is that the big assembled-on-orbit heat shield would allow a 10 ton vehicle to land cargo safely on Mars while the Ares V scheme would not land cargo or people safely on Mars. Mars direct would do no better. See [http://www.universetoday.com/2007/07/17/the-mars-landing-approach-getting-large-payloads-to-the-surface-of-the-red-planet/ The mars landing Approach: Getting Large Payloads to the Surface of the Red Planet]

Just as all economic activity in orbit so far has been done by robots, assembling a spacecraft to go to Mars should be done by robots and setting up the infrastructure for people to survive on Mars should be done by robots. There are some technical difficulties with this approach that must be addressed, but they seem likely to be amenable to solution.

=== Terminal Velocity of the Big Heat Shield Vehicle ===
For a rough estimate take as a starting point the estimated terminal velocity for the Apollo reentry vehicle. Estimate that the square of the terminal velocity is directly proportional to mass and the force of gravity and inversely proportional to air density and cross sectional area. The 10 metric ton vehicle with a 52 meter diameter heat shield should then have a terminal velocity less than 56 meters per second (125 miles per hour) at Mars' surface. If near the end of this descent a hole is burst through the bottom of the heat shield in the center right under the payload and hinges swing that portion of the heat shield out of the way, then retro rockets can fire at an altitude of 160 meters for 5.1 seconds with an acceleration of 11 meters per second squared and bring the payload to a stop about 17 meters above the heat shield that crashes into Mars. A few more seconds for horizontal maneuvering and throttling the rocket motor brings the payload safely to Mars on landing legs. A guesstimate of the required rocket delta V for this maneuver is about 68 meters per second.

== References ==
<references/>

[[Category:Missions]]

Low energy trajectories

2018-09-13T05:06:24Z

Pb: fix link

Cargo could be inserted into '''low-energy trajectories'''. These trajectories will take longer then current [[Hohmann transfer orbit|Hohmann transfer orbits]], but will save [[propellant]] and thus allow more payload.
Low cost of transport will be vital in [[Unmanned setup of a whole settlement|setting up a colony]] and in resupplying a [[semi-autonomous colony]].

== Methods ==
*[[Bi-elliptic transfer|Bi-elliptic transfers]]
*The [https://en.wikipedia.org/wiki/Interplanetary_Transport_Network Interplanetary Transport Network]
*Using [[ion thrusters]]/[[solar sails]]/[[plasma sails]]

== Disadvantages ==
*Such trajectories will take much longer then Hohmann transfers
*This means cargo will be exposed to the harsh conditions of space for a longer period of time. New and ingenious methods will be needed to solve this.

[[category:Technology]]
[[category:Orbital Mechanics]]

List of Maps

2018-09-03T15:33:48Z

Pb: creating lists

==Accepted Norms==
The currently accepted prime meridian of Mars is centered on the crater [https://en.wikipedia.org/wiki/Airy-0 Airy-0]. From there you can measure longitude using any system, degrees/minutes/seconds, decimal degrees, east/west, positive/negative, [lon, lat], [lat, lon], whatever. The prime meridian's location is fixed, but after that all bets are off.

MOLA data uses planetocentric latitude (as opposed to planetographic latitude).<ref name="mola">https://astrogeology.usgs.gov/search/details/Mars/GlobalSurveyor/MOLA/Mars_MGS_MOLA_DEM_mosaic_global_463m/cub</ref>

There is a system of [https://en.wikipedia.org/wiki/List_of_quadrangles_on_Mars quadrangles] invented by the USGS which are occasionally used.

The zero-altitude datum is not officially standardized, though most maps I've seen use MOLA elevation data, which is defined to be relative to a datum defined by the gravitational field of Mars.<ref name="mola" />

==Web-based Maps==

* [https://marstrek.jpl.nasa.gov JPL "Mars Trek"]
* [https://webgis2.wr.usgs.gov/Mars_Global_GIS/ USGS Mars Global GIS Mapping Application]
* [https://www.google.com/mars/ Google Mars]
* [http://www.arcgis.com/apps/Viewer/index.html?appid=ee4fd19d7d514bb192359534f27169b8 Life On Mars?] - Visualize "filling up" to different sea levels.
* [http://chrisherwig.org/planets/map/curiosity-rover/ Interactive Curiosity path through sol 351]
* [http://maps.stamen.com/mars/#6/10.034/-62.161/0/-600/400/0.785 3D contours by Stamen]

==Downloadable Maps==

* [https://pubs.usgs.gov/sim/3292/ USGS geologic map]
* [https://pubs.usgs.gov/imap/i2782/ USGS topographic map]

==Data Resources==

* [https://astrogeology.usgs.gov/search/details/Mars/GlobalSurveyor/MOLA/Mars_MGS_MOLA_DEM_mosaic_global_463m/cub MOLA data]
* [https://github.com/openplanetary/opm/wiki/OPM-Data-Sets-Repository OpenPlanetaryMap datasets]
* [https://github.com/openplanetary/opm/wiki/OPM-Basemaps OpenPlanetaryMap basemaps]
* [https://webgis.wr.usgs.gov/pigwad/maps/mars.htm USGS available vector layers and basemaps]
* [https://planetarynames.wr.usgs.gov/Page/MARS/target Feature names]

==Software Tools==

* [https://leafletjs.com Leaflet]
* [https://github.com/jupyterlab/jupyter-renderers/tree/master/packages/geojson-extension Jupyter geojson extension]

==References==

User:Pb

2018-09-03T13:37:04Z

Pb: Created page with "I'm just another Mars enthusiast. Peter Brandt. Feel free to call me Pete. github/pbrandt1 peterbrandt.space"

I'm just another Mars enthusiast. Peter Brandt. Feel free to call me Pete.

github/pbrandt1
peterbrandt.space

Areostationary orbit

2018-09-01T03:34:24Z

Pb: /* Satellite Coverage */

A satellite in areostationary orbit is a satellite in a circular, [[synchronous orbit]] at the equator, making it appear stationary in the sky above the ground.

As shown on the [[synchronous orbit]] page, a satellite must have a semi-major axis of 20,427.7 km to have its orbital period equal the rotational period of Mars. Since an aerostationary satellite is in a circular orbit, the orbital radius equals the semi-major axis, 20,427.7 km, placing it at an altitude of 17,038.2 km above the Martian surface. This puts Mars stationary orbit above Phobos (9,376 km) and below Deimos (23,463 km).

==Stable Orbital Slots==
Mars is not perfectly spherical, and so most satellites in areosynchronous orbit will tend to drift if left free-flying without station-keeping maneuvers. However, there are latitudes which have been shown to be equilibrium points in areosynchronous orbit. There are two stable equilibrium points at 17.92W and 167.83E, and there are two unstable equilibrium points at 105.55W and 75.34E.<ref name="Silvaa">{{cite journal |author=Juan J. Silvaa and Pilar Romero Silvaa |title=Optimal longitudes determination for the station keeping of areostationary satellites |journal=Planetary and Space Science |volume=87 |issue=October |year=2013}}</ref>

==Satellite Coverage==
For an areostationary satellite, coverage can be approximated as a circular region below the satellite. This coverage area could be defined as one in which all points remain in the satellite's field of view at all times, or an area in which a unit on the ground maintains contact with the areostationary satellite at all times.

[[Image:StationarySatelliteFOV.png]]

The visible disk of Mars takes up a field of view of about 23.6° as seen from the satellite in areostationary orbit. The coverage zone is a cap on surface of Mars that reaches up to approximately 78° latitude.

[[Image:AreostationaryCoverage.png]]

The coverage zones in this map approximate those of satellites in the two stable equilibrium points (yellow) and the two unstable equilibrium points (cyan). Plotted with Leaflet using tiles from [https://github.com/openplanetary/opm/wiki/OPM-Basemaps OpenPlanetaryMap].

A more realistic plot of coverage might take into account both altitude data and the non-spherical shape of Mars.

===References===

File:StationarySatelliteFOV.png

2018-09-01T03:17:19Z

Pb: The field of view of a satellite in stationary orbit above Mars.

== Summary ==
The field of view of a satellite in stationary orbit above Mars.

Areostationary orbit

2018-09-01T03:06:51Z

Pb: /* Satellite Coverage */

A satellite in areostationary orbit is a satellite in a circular, [[synchronous orbit]] at the equator, making it appear stationary in the sky above the ground.

As shown on the [[synchronous orbit]] page, a satellite must have a semi-major axis of 20,427.7 km to have its orbital period equal the rotational period of Mars. Since an aerostationary satellite is in a circular orbit, the orbital radius equals the semi-major axis, 20,427.7 km, placing it at an altitude of 17,038.2 km above the Martian surface. This puts Mars stationary orbit above Phobos (9,376 km) and below Deimos (23,463 km).

==Stable Orbital Slots==
Mars is not perfectly spherical, and so most satellites in areosynchronous orbit will tend to drift if left free-flying without station-keeping maneuvers. However, there are latitudes which have been shown to be equilibrium points in areosynchronous orbit. There are two stable equilibrium points at 17.92W and 167.83E, and there are two unstable equilibrium points at 105.55W and 75.34E.<ref name="Silvaa">{{cite journal |author=Juan J. Silvaa and Pilar Romero Silvaa |title=Optimal longitudes determination for the station keeping of areostationary satellites |journal=Planetary and Space Science |volume=87 |issue=October |year=2013}}</ref>

==Satellite Coverage==
For an areostationary satellite, coverage can be approximated as a circular region below the satellite. This coverage area could be defined as one in which all points remain in the satellite's field of view at all times, or an area in which a unit on the ground maintains contact with the areostationary satellite at all times.

In the case where the satellites coverage defines an area in which a unit on the ground maintains contact with the satellite, for instance if the satellite were used as a radio relay, one might start by assuming that contact can be established if the satellite is at least 10 degrees above the horizon. In this case, the coverage zone would be a circle with a radius of approximately 4176 km, with the top of the coverage zone reaching a latitude of approximately 78.2°.

[[Image:AreostationaryCoverage.png]]

The coverage zones in this map approximate those of satellites in the two stable equilibrium points (yellow) and the two unstable equilibrium points (cyan). Plotted with Leaflet using tiles from [https://github.com/openplanetary/opm/wiki/OPM-Basemaps OpenPlanetaryMap].

===References===

File:AreostationaryCoverage.png

2018-09-01T03:02:24Z

Pb: /* Summary */

== Summary ==
A plot of example coverage areas for satellites in stationary orbit above Mars. The four areas are the approximate coverage zones for satellites in the two stable equilibrium points (yellow) and the two unstable points (cyan).

Plotted with Leaflet using tiles from [https://github.com/openplanetary/opm/wiki/OPM-Basemaps OpenPlanetaryMap]

== Licensing ==
{{CC-by-sa-4.0}}

File:AreostationaryCoverage.png

2018-09-01T02:58:14Z

Pb: Pb uploaded a new version of File:AreostationaryCoverage.png

== Summary ==
A plot of an example computed coverage area for an areosynchronous satellite. The constraint used in this example is that the satellite must appear 10 degrees above the horizon in order to maintain coverage.

The coverage zone in this image is centered on the equator at 180 degrees latitude.

Plotted with Leaflet and Mapbox by Pete Brandt, source code available at https://github.com/pbrandt1/flight-dynamics-tutorials
== Licensing ==
{{CC-by-sa-4.0}}

Areostationary orbit

2018-08-30T04:05:15Z

Pb:

A satellite in areostationary orbit is a satellite in a circular, [[synchronous orbit]] at the equator, making it appear stationary in the sky above the ground.

As shown on the [[synchronous orbit]] page, a satellite must have a semi-major axis of 20,427.7 km to have its orbital period equal the rotational period of Mars. Since an aerostationary satellite is in a circular orbit, the orbital radius equals the semi-major axis, 20,427.7 km, placing it at an altitude of 17,038.2 km above the Martian surface. This puts Mars stationary orbit above Phobos (9,376 km) and below Deimos (23,463 km).

==Stable Orbital Slots==
Mars is not perfectly spherical, and so most satellites in areosynchronous orbit will tend to drift if left free-flying without station-keeping maneuvers. However, there are latitudes which have been shown to be equilibrium points in areosynchronous orbit. There are two stable equilibrium points at 17.92W and 167.83E, and there are two unstable equilibrium points at 105.55W and 75.34E.<ref name="Silvaa">{{cite journal |author=Juan J. Silvaa and Pilar Romero Silvaa |title=Optimal longitudes determination for the station keeping of areostationary satellites |journal=Planetary and Space Science |volume=87 |issue=October |year=2013}}</ref>

==Satellite Coverage==
For an areostationary satellite, coverage can be approximated as a circular region below the satellite. This coverage area could be defined as one in which all points remain in the satellite's field of view at all times, or an area in which a unit on the ground maintains contact with the areostationary satellite at all times.

In the case where the satellites coverage defines an area in which a unit on the ground maintains contact with the satellite, for instance if the satellite were used as a radio relay, one might start by assuming that contact can be established if the satellite is at least 10 degrees above the horizon. In this case, the coverage zone would be a circle with a radius of approximately 4176 km.

[[Image:AreostationaryCoverage.png]]

The coverage zone plotted in this map is centered on the equator at 180 degrees latitude. The locations on the map, from left to right, are Jezero Crater (a NASA Mars 2020 rover candidate site), Elysium Planitia (InSight planned landing site), Gale Crater (Curiosity landing site), Columbia Hills (NASA Mars 2020 candidate), Oxia Planum (ExoMars 2020 candidate), and Mawrth Vallis (ExoMars 2020 candidate). It was plotted with Leaflet and Mapbox. Unfortunately, the Mapbox tiles are no longer available, but the source code for this map remains available here: <ref name="pbrandt">https://github.com/pbrandt1/flight-dynamics-tutorials/blob/master/Mars_Stationary_Satellite.md</ref>

===References===

Areostationary orbit

2018-08-30T03:51:25Z

Pb: fixing ref

A satellite in areostationary orbit is a satellite in a circular, [[synchronous orbit]] at the equator, making it appear stationary in the sky above the ground.

As shown on the [[synchronous orbit]] page, a satellite must have a semi-major axis of 20,427.7 km to have its orbital period equal the rotational period of Mars. Since an aerostationary satellite is in a circular orbit, the orbital radius equals the semi-major axis, 20,427.7 km, placing it at an altitude of 17,038.2 km above the Martian surface. This puts Mars stationary orbit above Phobos (9,376 km) and below Deimos (23,463 km).

==Stable Orbital Slots==
Mars is not perfectly spherical, and so most satellites in areosynchronous orbit will tend to drift if left free-flying without station-keeping maneuvers. However, there are latitudes which have been shown to be equilibrium points in areosynchronous orbit. There are two stable equilibrium points at 17.92W and 167.83E, and there are two unstable equilibrium points at 105.55W and 75.34E.<ref name="Silvaa">{{cite journal |author=Juan J. Silvaa and Pilar Romero Silvaa |title=Optimal longitudes determination for the station keeping of areostationary satellites |journal=Planetary and Space Science |volume=87 |issue=October |year=2013}}</ref>

==Satellite Coverage==
For an areostationary satellite, coverage can be approximated as a circular region below the satellite. This coverage area could be defined as one in which all points remain in the satellite's field of view at all times, or an area in which a unit on the ground maintains contact with the areostationary satellite at all times.

In the case where the satellites coverage defines an area in which a unit on the ground maintains contact with the satellite, for instance if the satellite were used as a radio relay, one might start by assuming that contact can be established if the satellite is at least 10 degrees above the horizon. In this case, the coverage zone would be a circle with a radius of approximately 4176 km. The coverage zone in this image is centered on the equator at 180 degrees latitude.

[[Image:AreostationaryCoverage.png]]

Areostationary orbit

2018-08-30T03:45:42Z

Pb: /* Satellite Coverate */

A satellite in areostationary orbit is a satellite in a circular, [[synchronous orbit]] at the equator, making it appear stationary in the sky above the ground.

As shown on the [[synchronous orbit]] page, a satellite must have a semi-major axis of 20,427.7 km to have its orbital period equal the rotational period of Mars. Since an aerostationary satellite is in a circular orbit, the orbital radius equals the semi-major axis, 20,427.7 km, placing it at an altitude of 17,038.2 km above the Martian surface. This puts Mars stationary orbit above Phobos (9,376 km) and below Deimos (23,463 km).

==Stable Orbital Slots==
Mars is not perfectly spherical, and so most satellites in areosynchronous orbit will tend to drift if left free-flying without station-keeping maneuvers. However, there are latitudes which have been shown to be equilibrium points in areosynchronous orbit. There are two stable equilibrium points at 17.92W and 167.83E, and there are two unstable equilibrium points at 105.55W and 75.34E.<ref name="Silvaa"/>

==Satellite Coverage==
For an areostationary satellite, coverage can be approximated as a circular region below the satellite. This coverage area could be defined as one in which all points remain in the satellite's field of view at all times, or an area in which a unit on the ground maintains contact with the areostationary satellite at all times.

In the case where the satellites coverage defines an area in which a unit on the ground maintains contact with the satellite, for instance if the satellite were used as a radio relay, one might start by assuming that contact can be established if the satellite is at least 10 degrees above the horizon. In this case, the coverage zone would be a circle with a radius of approximately 4176 km. The coverage zone in this image is centered on the equator at 180 degrees latitude.

[[Image:AreostationaryCoverage.png]]

<ref name="Silvaa">{{cite journal |author1=Silvaa, J. |author2=Romero, P. |title=Optimal longitudes determination for the station keeping of areostationary satellites |journal=Planetary and Space Science |volume=87 |issue=October |year=2013}}</ref>

Areostationary orbit

2018-08-30T03:44:29Z

Pb: /* Stable Orbital Slots */

A satellite in areostationary orbit is a satellite in a circular, [[synchronous orbit]] at the equator, making it appear stationary in the sky above the ground.

As shown on the [[synchronous orbit]] page, a satellite must have a semi-major axis of 20,427.7 km to have its orbital period equal the rotational period of Mars. Since an aerostationary satellite is in a circular orbit, the orbital radius equals the semi-major axis, 20,427.7 km, placing it at an altitude of 17,038.2 km above the Martian surface. This puts Mars stationary orbit above Phobos (9,376 km) and below Deimos (23,463 km).

==Stable Orbital Slots==
Mars is not perfectly spherical, and so most satellites in areosynchronous orbit will tend to drift if left free-flying without station-keeping maneuvers. However, there are latitudes which have been shown to be equilibrium points in areosynchronous orbit. There are two stable equilibrium points at 17.92W and 167.83E, and there are two unstable equilibrium points at 105.55W and 75.34E.<ref name="Silvaa"/>

==Satellite Coverate==
For an areostationary satellite, coverage can be approximated as a circular region below the satellite. This coverage area could be defined as one in which all points remain in the satellite's field of view at all times, or an area in which a unit on the ground maintains contact with the areostationary satellite at all times.

In the case where the satellites coverage defines an area in which a unit on the ground maintains contact with the satellite, for instance if the satellite were used as a radio relay, one might start by assuming that contact can be established if the satellite is at least 10 degrees above the horizon. In this case, the coverage zone would be a circle with a radius of approximately 4176 km. The coverage zone in this image is centered on the equator at 180 degrees latitude.

[[Image:AreostationaryCoverage.png]]

<ref name="Silvaa">{{cite journal |author1=Silvaa, J. |author2=Romero, P. |title=Optimal longitudes determination for the station keeping of areostationary satellites |journal=Planetary and Space Science |volume=87 |issue=October |year=2013}}</ref>

Areostationary orbit

2018-08-30T03:43:02Z

Pb: creating aerostationary orbit

A satellite in areostationary orbit is a satellite in a circular, [[synchronous orbit]] at the equator, making it appear stationary in the sky above the ground.

As shown on the [[synchronous orbit]] page, a satellite must have a semi-major axis of 20,427.7 km to have its orbital period equal the rotational period of Mars. Since an aerostationary satellite is in a circular orbit, the orbital radius equals the semi-major axis, 20,427.7 km, placing it at an altitude of 17,038.2 km above the Martian surface. This puts Mars stationary orbit above Phobos (9,376 km) and below Deimos (23,463 km).

==Stable Orbital Slots==
Mars is not perfectly spherical, and so most satellites in areosynchronous orbit will tend to drift if left free-flying without station-keeping maneuvers. However, there are latitudes which have been shown to be equilibrium points in areosynchronous orbit. The are two stable equilibrium points at 17.92W and 167.83E, and there are two unstable equilibrium points at 105.55W and 75.34E.<ref name="Silvaa"/>

==Satellite Coverate==
For an areostationary satellite, coverage can be approximated as a circular region below the satellite. This coverage area could be defined as one in which all points remain in the satellite's field of view at all times, or an area in which a unit on the ground maintains contact with the areostationary satellite at all times.

In the case where the satellites coverage defines an area in which a unit on the ground maintains contact with the satellite, for instance if the satellite were used as a radio relay, one might start by assuming that contact can be established if the satellite is at least 10 degrees above the horizon. In this case, the coverage zone would be a circle with a radius of approximately 4176 km. The coverage zone in this image is centered on the equator at 180 degrees latitude.

[[Image:AreostationaryCoverage.png]]

<ref name="Silvaa">{{cite journal |author1=Silvaa, J. |author2=Romero, P. |title=Optimal longitudes determination for the station keeping of areostationary satellites |journal=Planetary and Space Science |volume=87 |issue=October |year=2013}}</ref>

File:AreostationaryCoverage.png

2018-08-30T03:41:04Z

Pb: A plot of an example computed coverage area for an areosynchronous satellite. The constraint used in this example is that the satellite must appear 10 degrees above the horizon in order to maintain coverage. The coverage zone in this image is centered...

Synchronous orbit

2018-08-30T02:53:40Z

Pb: creating synchronous orbit page

A satellite in synchronous orbit takes one day to complete one orbit, making the satellite return to the same point above the ground every day. If a satellite's inclination is 0 and the orbit is perfectly circular, it will appear stationary in the sky. If the inclination is nonzero and the orbit is non-circular, then the satellite will follow the path of a figure 8 about a central point on the equator.

For a satellite to maintain a synchronous orbit around Mars, set the orbital period of a satellite equal to one sidereal day.

T = 2 pi sqrt(a^3 / mu)

where:

* T = the orbital period, which is set to one sidereal day on Mars, 88642.66 s
* mu = GM = the standard gravitational parameter, which is 4.282837e4 km3 s−2 for Mars
* a is the semi-major axis of the orbit

Solving for a, we get

a = 20,427.7 km

For an [[Areostationary orbit]], the radius will be 20,427.7 km and the inclination will be zero.

List of Planned Missions

2018-08-28T19:40:25Z

Pb: Created page with "{| class="wikitable" |- ! Mission ! Category ! Planned Launch ! Organization |- | ExoMars 2020 | Lander, Rover | 2020 | ESA |- | Mar..."

{| class="wikitable"
|-
! Mission
! Category
! Planned Launch
! Organization
|-
| ExoMars 2020
| [[:category:landers|Lander]], [[:category:rovers|Rover]]
| 2020
| ESA
|-
| Mars 2020
| [[:category:rovers|Rover]]
| 2020
| NASA
|-
| Mars Hope
| [[:category:orbiters|Orbiter]]
| 2020
| UAE
|-
| Mars Global Remote Sensing Orbiter and Small Rover
| [[:category:orbiters|Orbiter]], [[:category:rovers|Rover]]
| 2020
| CNSA
|-
| Mars Orbiter Mission 2
| [[:category:orbiters|Orbiter]]
| 2022
| ISRO
|-
| Martian Moons Exploration Phobos Lander with Sample Return
| [[:category:landers|Lander]]
| 2024
| JAXA
|}

List of current missions

2018-08-28T19:18:04Z

Pb: adding location, status columns. removing phoenix (not current mission). adding insight