Smallest Human Ascent or Descent Lander for Mars Or Earth

GW Johnson · 2016-09-07 12:12:50

What you do is scale the mass/chute area loading, and the Reynolds number. You can match chute area loading with a small experiment, there is little need to fly giant things, until the final, all-up tests. As for Reynolds number effects, your experiment usually just needs to be large enough so that you achieve turbulent flow patterns. You only need turbulent flow, not an actual parameter match.

But it is necessary to do this at the correct densities, which is why Mars chutes are tested above 100,000 feet in Earth's atmosphere. It is also necessary to do this at the correct speeds. That way you get the correct dynamic pressures and the correct object forces and dynamics (provided you have the right mass/area loading).

You have a little more latitude to achieve these things in flight tests. Wind tunnels can be very limiting, ranging to completely unsuitable. It varies from one facility to the next. They're all different. Flight tests are more expensive, though.

GW

Last edited by GW Johnson (2016-09-07 12:14:24)

SpaceNut · 2017-03-06 18:12:13

I typically will use comparison scaling when I look to what we have and what we want to do with it.

So lets look at the Falcon 9 heavy which can lift 53 mT to earth orbit and when I look at the same rocket being launched from mars surface we can scale it down by a factor of 2.66 as mars is only 0.375 that of Earth for gravity.

Or even think of it this way when we launch a falcon 9 v1.2 FT we can loft to orbit 22.8 mT so if this was setting on mars and scaled down by the 2.66 factor then we have a payload of ~8.6 mT which is approximately the size of a red Dragon. Normally its 70 m tall for payload and both stages. Which suggests a scaled down height of ~26.5 meters which is a long way to the ground from a capsule door but that could change as we work through the stage numbers. The Dragon Spacecraft & Trunk 8.1m (26.6 ft) height. Payload flairing is 5.2 m tall but I think we can use the dragons height once we recalculate the stages under the capsule changes.

Since Mars is a near vacuum it only makes sense to configure the lander with the correct engine (934 kN (Vacuum) Merlin Burn time of 397sec) for this iteration of design meaning we need for the second stage only ~351 kN for mars. Which means only 3 engines are required.

The normal first stage engine at sea level, for a total thrust of 6,804 kN ( Merlin 1D burn time 162sec) means we need to target ~ 2,560 kN for launch from mars surface for 1st stage. One engine is only 756 kN so giving the new first stage the 934 kN (vacuum engine) only makes sense.

Current Space x page indicates first stage Thrust At Sea Level 7,607 kN Thrust In Vacuum 8,227 kN, which appears to be 47 m tall. Mars height would be 17.7 m tall for the stage. Normally both stages plus flairing is 70 so if we remove the Pay load flairing height of 5.2 m out of the 70 m and 47 m first stage then the second stage is 17.8 m tall for the falcon 9 rocket and for mars at ~ 6.7m.

Merlin Vacuum Certification On March 7, 2009, SpaceX performed a full mission duration firing of the new Merlin Vacuum engine at McGregor. The engine fired for six minutes or 360 seconds, consumed 45.36 tonnes of propellant, and demonstrated a vacuum specific impulse of 342 seconds, highest ever for a U.S. hydrocarbon rocket engine.

The two-stage, 313 tonne, kerosene/LOX rocket so proportional for mars means ~118 t of fuel and oxidizer.
The second stage normally fires for 397 sec. so probably close to 50 t for the stage and for mars its 18.8 t.
The first stage then calculates to 263 t of which for reusuability is about 70 % leaving the rest for landing. So if we were on orbit and coming down then we would need ~79 t so for a retropropulsion we need ~ 30 t to land not count stage and capsule mass plus on orbit.

So mars sized falcon 9 is 32.5 m tall for a dragon plus truck of 8.1 m with a first stage 17.7 m with fuel/oxidizer of 30 t with a second stage that is 6.7 m tall and fueled with 18.8 t of which we would need to land with the second stage empty and a bit large fuel fill for the first stage for th combined landing attempt.

Seems like we could make this a single stage to orbit with the 9 vacuum engines.

Of course as meantioned the RP-1 and Lox is not the correct fuel for mars if we are making insitu fuels so now we would redo these calculations for the new fuel type and engines that we could use.

Something else is to change the 3.7 m diameter to the more practical 10 m to make the stages shorter as well.

With this long post we can see that for mars we are only landing a stackup of no more than the hieght of the current stage Falcon 9 for the entire mars lander. Which is a small 40 m tall at the 3.7 m diameter but if we go to the 10 m diameter then we are looking at something under 15 m tall. That said 4 engines for a complete single stage to orbit would seem to be correct.

So now to covert for the methane-lox and engines for the same scaling once engines numbers are looked up.

While we could land with NTO-hydrazine we do not have an insitu ability developed yet for manufacturing it for refueling.

The same holds true of NTO / MMH even with the super Draco engines.

We really need the ability to make other insitu fuels so as to be able to use what we have rather than changing plans for what we need developement for.

https://en.wikipedia.org/wiki/Dinitrogen_tetroxide

Nitrogen tetroxide is used as an oxidizer in one of the more important rocket propellants because it can be stored as a liquid at room temperature. Nitrogen tetroxide is made by the catalytic oxidation of ammonia: steam is used as a diluent to reduce the combustion temperature. In the first step, the ammonia is oxidized into nitric oxide:
4NH3 + 5O2 → 4NO + 6H2O
Most of the water is condensed out, and the gases are further cooled; the nitric oxide that was produced is oxidized to nitrogen dioxide, which is then dimerized into nitrogen tetroxide:
2NO + O2 → 2NO2
2NO2 ⇌ N2O4
and the remainder of the water is removed as nitric acid. The gas is essentially pure nitrogen dioxide, which is condensed into dinitrogen tetroxide in a brine-cooled liquefier.

https://en.wikipedia.org/wiki/Monomethylhydrazine
Monomethylhydrazine (MMH) is a volatile hydrazine chemical with the chemical formula CH3(NH)NH2.

SpaceNut · 2018-07-25 17:32:17

Redirect from Getting to Mars with REAL technology, & what's currently missing

Current launchers but they can not without subassembly on orbit get anything reasonable for mars to the surface of mars.
Current capsules for earth reentry for humans are getting closer by the day but none at the moment from Orion, Dragon, Starliner and Dreamchaser.
Current habitats for transit to mars are iss modules, lunar gateway from simular base of build, ISS Beam but little else has been built and
lest we for get no artifical gravity and minimal radiation protection.

Current mars capable landers none for humans but cargo a modified skycrane for 1.5 mT.
Current surface habitats none and even though a Beam weighs with in the skycrane landing capability.

https://en.wikipedia.org/wiki/Bigelow_E … ity_Module
https://www.nasa.gov/mission_pages/stat … /1804.html
https://en.wikipedia.org/wiki/SpaceX_CRS-8

The BEAM weighs approximately 3,000 lbs (1,360 kg) and travels within the unpressurized cargo hold of a Dragon capsule.
BEAM is expanded from its packed dimensions of 5.7 feet long and just under 7.75 feet in diameter to its pressurized dimensions of 13 feet long and 10.5 feet in diameter and has 560 cubic feet of pressurized volume.

https://en.wikipedia.org/wiki/Mars_Science_Laboratory

NASA has looked into a skycrane system to land larger craft, but the numbers are tough. Curiosity had a total mass of two tons, but a manned lander would probably clock in at 10 or 15 tons. It’s unclear if it would be possible to land something like that on Mars with our current technology.

1. use a dragon for 1 and a cygnus for going to mars orbit and back with.

Mars lander with modified skycrane mass 2mT and maybe a skosh more if possible.
All landers would mean about 15mT on orbit mass before attempting to land.

2. mars inflateable Beam habitat lander
3. the small crew lander
4. power lander
5. return ascent vehicle
6. however many cargo landers we need for a crew of one to survive.
7. science lander

SpaceNut · 2018-07-26 16:57:05

Repost of multiple topics to consolidate information for a 1 manned mission to mars using technology we have today.

GW Johnson wrote:

As near as I can tell, Red Dragon was a modified crew Dragon. They ripped out the seats and life support, and installed cargo racks more like cargo Drago, inside the pressure shell.
They deleted the parachutes entirely. It was to use the same landing legs-through-the-heat shield that crew Dragon was originally supposed to have.
I'm just guessing it had the cargo Dragon-type unpressurized service module, with the solar wings.
Crew Dragon carries more propellant than cargo Dragon, as best I can tell. 1800 kg vs 1200 kg of MMH-NTO.
If you judiciously limit cargo weight loaded aboard to 1-2 tons, then as best I can tell, the delta-vee without the service module is about 1 km/s. There's side and top hatches for moving stuff in and out, just like crew and cargo Dragon.
That is just barely enough for a propulsive landing after aerobraking reentry direct from an interplanetary trajectory. I'd prefer something in the 1-to-1.4 km/s range, but Spacex knows its ship better than me.
Current status is Red Dragon "gone", and the legs and propulsive landing capability removed from crew Dragon at NASA's insistence.
GW

That said going with the modified Dragon for a 1 person landing what would we need to do to be able to back to orbit if we could alter the lander for that purpose on the surface of mars by using cargo brought to the surface. Would making the engines run with LH2 / Lox or some other combination of fuels that we could bring or make on mars be enough or are we in need of more engines? Could we modify the dragon to be a two stage or add solids to get the needed boost to orbit?

GW Johnson · 2018-07-27 08:11:56

Anything involving Dragon is a one-way trip to the surface. There is no chemical Isp high enough to ascend to orbit using any mod to the SuperDraco thruster system. The mass ratio simply isn't there to do any sort of refueling on the surface and expect this to work. Don't forget, in order to aerobrake for the landing, the service module must have been jettisoned. There is no place in the capsule to put extra propellant.

To ascend to Mars orbit requires as an utter theoretical minimum 3.6 km/s delta-vee as a deliverable. Something closer to 4 km/s is more practical and safer for realistic manned (or unmanned) operations. Therefore, some sort of ascent stage or vehicle is required, separate from any version of Dragon-as-descent vehicle, in order for the trip not to be "one-way, period".

I would not recommend using a version of the Dragon as the human transport capsule/payload of an ascent vehicle, unless you also intend to provide an abort-to-surface capability. It is just too heavy compared to the utter minimum needed only to ascend (in which case you are betting the crew's lives that ascent will be successful). If you intend to provide the abort-to-surface capability, you must also have something to use for rescue. Otherwise, there is no point to having the crew survive an abort during ascent.

Whatever this ascent vehicle turns out to be, first it must survive descent, second it must land, and third it must land IN THE RIGHT PLACE in order to be accessed and used. If you have done all that, you might as well forget the separate Dragon-based descent vehicle, and just use the descent of the ascent vehicle for your crew descent. Why build 2 when 1 will do?

Just sayin' ....

GW

Last edited by GW Johnson (2018-07-27 08:12:46)

GW Johnson · 2018-07-27 09:10:11

Having said what I said in the previous post, now consider the total delta-vee required of a chemically-propelled two-way descent/ascent vehicle. As I said above, you need around 3.6 to 4.0 km/s delivered delta-vee out of it to make an ascent to orbit. As best I can tell, most of the delta-vee for descent is done with aerobraking, not propulsion.

If you assume a retropropulsive landing for a vehicle with a large ballistic coefficient, you come out of hypersonics at local Mach 3 around 5 km from the surface, traveling steeply downward. Theoretically, you have to kill 0.7 km/s velocity before you strike. Actually, you need to add in some serious gravity loss and some hover/terminal maneuver capability to safely land. The landing delta-vee is 1 to maybe 1.4 km/s delivered delta-vee.

Add those together, and a design with something around 5.0 km/s total delta-vee capability is capable of making the descent, and the ascent, COMPLETELY UNREFUELED!!! In other words, because of the Martian atmosphere (unlike the moon), a two-way, single-stage chemically-propelled "landing boat" is indeed quite feasible for Mars!!!

Assume something like 320 s Isp for storables like MMH-NTO. No evaporative losses!!! The exhaust velocity is something like 3.14 km/s. The velocity ratio is 5/3.14 = 1.592. Exponentiating, the mass ratio is 4.915. Slightly challenging, but doable. It corresponds to propellant mass fraction of 80%.

Now, this vehicle must have retractable landing legs, a heat shield capable of more than one use, and the volume in which bulky cargo may be carried. As a hunch, I think 15% inert fraction is more realistic than 10%. The very best stages that have none of these features are now 5% inert. Go with 15% That leaves 5% as payload.

Now, payload is maybe 2-4 people, a 1 ton rover car, and 2-3 tons of equipment and supplies, assuming the lander will serve as surface habitat. Call it 7 tons delivered. That's 5% of ignition mass, which then is 140 tons. Not too bad for a two-way unrefueled item. It looks closer to 78 tons at the higher Isp of LOX-LCH4. Using 360 s Isp, I get 76% propellant mass fraction. With the same inert, payload is 9%. At 7 tons payload, vehicle is 78 tons at ignition.

If surface propellant production were available, you could land a far bigger payload by landing with only the landing propellant in the tanks. Then fill them on the surface to near-full for the return trip, or fly full-tank with more than 7 tons payload for the return trip.

Point is, if you design this thing for operating as a baseline unrefueled two-way, that gives you great growth potential WITH THE SAME VEHICLE if rapid local propellant production becomes feasible.

It is this kind of a vehicle that provides rescue capability for a stranded crew, if you have more than one with you when you go to Mars. For that design, the crew cabin could indeed be a Dragon or something similar, for ascent abort capability to the surface. If you do this, you also have a descent abort-to-surface capability. All bases covered for crew survival, no matter what.

The only downside is a mass ratio designed around low Mars orbit. Another 1.5-2 km/s delta-vee is required for Mars escape directly to an interplanetary trajectory. That puts your payload fractions too low for practical design. Having this kind of reusable two-way single-stage landing boat DOES IMPLY that you (at least initially) base your explorations out of low Mars orbit, perfect for an orbit-to-orbit manned transport design for the trip. But it does offer the opportunity to explore multiple sites in the one trip to Mars.

See where this is headed? Designing-in both safety and practicality leads directly to a preferred mission architecture of an orbit-to-orbit manned transport, with landers and their propellant sent ahead unmanned, for rendezvous in low Mars orbit. You get the smallest orbit-to-orbit transport design if you also send its return propellant ahead to low Mars orbit unmanned.

This assumes a chemically-propelled orbit-to-orbit transport, so that it can easily be spun for artificial gravity during the coasting. The next burn's propellant is your radiation shadow shield. The unmanned stuff can be sent slow-boat spiral-out / spiral-in with electric propulsion to minimize mission cost. The men cannot be sent that way, because of long exposure time spiraling through the Van Allen Belts, outbound, and for the return.

If you plan on recovering (and reusing) the orbit-to-orbit transport in Earth orbit, then you still have the final arrival propellant as a radiation shadow shield during the return coast. Suspenders-and-belt. Every safety item has a way out.

And THAT is the Mars Mission 2016 posting I put up over at "exrocketman". If while doing this, you grade off and maybe pave a big landing field at your "best" selected base site, then you have a safe place to land BFS's and eventually start a settlement.

GW

Last edited by GW Johnson (2018-07-27 09:23:51)

SpaceNut · 2018-07-27 20:25:40

Thanks GW for the postings.

I am still learning about the fuels, engines, ISP, Mass fraction ect...but will figure it out someday....

Went searching for images and mass information today and here are some of them.

What we did on the moon

Current and possibly close capsules

Something simular to what we would be using for mars that was the early images for the moon

GW Johnson · 2018-07-28 10:12:12

These images, particularly the last one, look much like the two-way one-stage lander designs I posted over at "exrocketman". I did a paper design study for what such a craft might look like, for each of 4 typical propellant combinations, surprisingly coming out all crudely the same dimensions, but with different weight statements. It is payload size that sets these dimensions and weights for each propellant combination.

That study is dated August 21, 2013, and titled "Reusable Chemical Mars Landing Boats Are Feasible". The site is http://exrocketman.blogspot.com. Note that I did this before either my 2013 or 2016 looks at Mars mission architectures. It also predates any knowledge of Spacex's BFR stuff. If you combine landers like this with my 2016 take on how to stage a Mars mission, you have a pretty good plan that addresses every hazard, has "a way out" for every phase if something fails, and provides hardware that can be reused for subsequent missions.

That sort of outcome is rather hard to beat. Spacex's approach with its BFR is different, and has some unresolved hazards, such as rough-field landings with a tall, narrow vehicle. They land at only one site. I land at multiple sites before deciding which one is really best.

GW

Last edited by GW Johnson (2018-07-28 10:13:10)

elderflower · 2018-07-29 04:09:09

GW. You could improve your mission performance, I would think, if you were to manufacture just LOX from atmospheric CO2. This is fairly straightforward given a large enough power supply. Then you only need a storable propellant such as propane or even RP1. Your landed mass can be much reduced as LOX is the major part by mass of your fuel.

GW Johnson · 2018-07-29 09:57:07

I agree. It just depends upon being able to make enough fast enough, and of sufficient quality. Nobody is yet field-testing any equipment like that yet. They need to.

GW

Last edited by GW Johnson (2018-07-29 09:57:34)

SpaceNut · 2018-07-29 10:30:50

Moxie will be riding on the 2020 rover landing as a demonstrator field test.
Posted the MSl entry to landing stuff for the skycrane in its topic.
The Dragon is as it stands would be a skycrane on steriods that keeps the backshell, heat shield a pressure vessel, and payload to the surface from the same location where the parachutes and retro rockets take over to land the unit to the surface.

That said I think a 2 person mission is more than just possible if we get the landing sites for all items within a close enough target. Thats something else we will need....

GW Johnson · 2018-07-29 11:26:27

Spacenut:

Red Dragon was never intended to use any chutes. At 6+ tons, it comes out of hypersonics too low and too fast for a chute to deploy, much less decelerate anything. About Mach 3 at about 4-5 km. Trending steeply downward. It was aerobrake, then retropropulsive to touchdown.

And as I pointed out above, any form of Dragon by itself is a one-way descent to the surface. There is not the mass ratio to make an ascent possible, even if the propellants were available on the surface. It would need about 4-5 times the propellant tankage volume to create that capability. That would take a 2-3 times bigger service module, but that module must be discarded to uncover the heat shield for the descent in the first place.

Might as well design a proper lander from a clean sheet of paper. Single stage designs will fall in the 70-140 ton ignition weight for 3 tons of payload, but can be reusable. Two-stage designs will have a much larger fraction payload/ignition, but CANNOT be re-used. Rock-and-a-hard-place choice that must be faced up front.

GW

Last edited by GW Johnson (2018-07-29 11:28:18)

kbd512 · 2018-07-29 18:21:26

GW,

Separate the part of the vehicle that does the landing and launching from the payload and the vehicle mass (propellant mass) problem becomes a lot easier to solve.

Why do we need to send a complete capsule system back into orbit?

If something goes wrong during reentry or launch, there's only one realistic outcome and we both know what that is.

What's so wrong with a beefed up SkyCrane system that stays attached to the payload until it lifts off again, carrying only the astronauts back into orbit?

Why do we need a massive pressurized capsule system to do this?

To be clear, this is what I want to build:

1. SkyCrane stays attached to a HIAD-equipped habitat module slung below for EDL, so the heavy landing gear stays attached to the payload and maybe the HIAD functions as a sort of air bag / landing gear to eliminate purpose-built landing gear for small payloads
2. SkyCrane has a small pod / back shell for suited astronauts to strap into atop the habitat module
3. When it's time to leave, the astronauts strap themselves in and ride the SkyCrane all the way to orbit
4. SkyCrane gets refueled and reattached to another payload in orbit after rendezvous with the orbiting ITV

That should eliminate the immediate need for giant vehicles with giant fuel tanks to obtain some measure of reusability.

GW Johnson · 2018-07-29 18:46:26

Spacenut:

The skycrane idea works for an aeroshell / chute combination to get you all the way up to 1 ton landed. NO ONE ANYWHERE can say this would also work with 2 tons, much less 10 or 20 tons. And I say it won't, because you come out too fast and too low to use a chute at all, at weights like that.

Skycrane is just the final retropropulsion for the landing of a nonpropulsive item. If you already have ascent engines, use them to land. What the hell do you need a skycrane for?

GW

Last edited by GW Johnson (2018-07-29 18:47:23)

kbd512 · 2018-07-29 19:48:53

GW,

I don't want to use any parachutes. The entire landing will be propulsive with a HIAD aerobrake. Regarding the ascent engines, that's exactly what I want to do. The SkyCrane sits atop the payload, the payload provides the landing gear / surface. My "SkyCrane" idea (not my idea, obviously) is an enlarged version of the original, without parachutes, and with more powerful engines and more fuel. I just want a device that attaches to the top of the habitat that pulls double duty as a propulsive lander and minimal capability ascent vehicle.

SpaceNut · 2018-07-29 21:11:29

MSL EDL 5,293-pound (2,401-kilogram) entry, descent and landing system (aeroshell plus fueled descent stage)

Here is an important piece to the puzzle Comparing Heat Shields: Mars Science Lab vs. SpaceX Dragon

Measuring nearly 15 ft (4.5 m) in diameter, the MSL heat shield was the largest to ever travel to another planet.

The heat shield on the Dragon is slightly smaller than the MSL measuring 13 ft (4 m) in diameter.

SpaceNut · 2018-07-29 21:44:14

The MSL descent stage uses eight rockets, called Mars lander engines (MLE), positioned around the perimeter in four pairs. These are the first throttleable engines for a Mars landing since the Mars Viking landings in 1976. They were built by Aerojet, in Richmond, Wash., withthrottle valve assemblies from Moog Inc., East Aurora, N.Y. Each can provide an adjustable amount of thrust up to about 742 pounds (3,300 newtons). The propulsion system of the descent stage uses pressurized propellant. Three spherical fuel tanks provide a usable propellant load of about 853 pounds (about 387 kilograms) of hydrazine, a propellant that does not require an oxygen source. Two spherical tanks of pressurized helium provide pressure for propellant delivery, moderated by a mechanical regulator.

Engines: eight side-mounted SuperDraco engines, clustered in redundant pairs in four engine pods, with each engine able to produce 71 kilonewtons (16,000 lbf) of thrust Each pod—called a "quad" by SpaceX—contains two SuperDraco engines plus four Draco thrusters. "Nominally, only two quads are used for on-orbit propellant with the Dracos and two quads are reserved for propulsive landing using the SuperDracos. 1900 kg of propellant would provide the Δv required for soft landing at 2.4 m/s.

SpaceNut · 2018-07-29 22:14:19

GW Johnson wrote:

Spacenut:
The skycrane idea works for an aeroshell / chute combination to get you all the way up to 1 ton landed. NO ONE ANYWHERE can say this would also work with 2 tons, much less 10 or 20 tons. And I say it won't, because you come out too fast and too low to use a chute at all, at weights like that.
Skycrane is just the final retropropulsion for the landing of a nonpropulsive item. If you already have ascent engines, use them to land. What the hell do you need a skycrane for?
GW

That is why you need to get to a 10 m heatshield size to allow for the higher mass and to allow for the slowing without a parachute. As to the use of HIAD aerobraking we will need some more testing.

elderflower · 2018-07-30 04:31:42

For an exploration/pioneering mission, perhaps the descent capsule could serve as a hopper to allow a wider area exploration, returning to an ascent vehicle with a separate orbital capsule. The ascent vehicle booster could be discarded, or if it can be refuelled on the surface, it could be landed back to its origin. In the disposable case it might be well to use a solid booster to keep it simple. This requires 3 launches from Earth.
Direct entry of the ascent equipment plus crew survival stores for say 26 months.
Orbit to surface of the descent/ hopper/hab unit with a second batch of stores and fuel for hopping exploration.
Orbit to Orbit transfer vehicle. Ascent vehicle docks with this and forms the earth return capsule for personnel and payload after fitting a heat shield in Mars orbit. I think Orbit to Orbit transfer vehicle will have to be discarded on return to Earth as there is no way that fuel would be available to get it into an Earth orbit.
The 1st and 3rd units need to be in place before the second departs from LEO.
Failure of any of these either terminates the crew, or they stay on Mars, or maybe in Mars orbit, until a recovery mission arrives.

SpaceNut · 2018-07-30 19:03:20

For hopper use of the craft that we would land in you would need lots ofpower to make fuel or a number of fuel cargo landings. A minimal landing with some alterations of what we have would put that on a back burner until someone wants to put up the bucks.

Being minimal puts a mars mission just above that of a sortie mission but at least we can prove lots of stuff out in this level of missions still the same.

What would it take to make a better lift skycrane to get towards something a bit better?

What the skycrane does is dump mass on the way down while we want it to land on the ground.

So in that sense the Dragon which lands proves that better fuels, engines would allow for the small crew to land and go about exploring mars with the issue of how to get back to orbit. So how do we change the odds?

Make a small crew ascent vehicle to bring them back to orbit after refueling form fuel sent to the surface as cargo.
Assemble a new stage from parts and pieces from the other landers to make it back to orbit.

Are there really other options not building a clean sheet to large to go with on current launchers....

SpaceNut · 2018-07-31 17:05:46

So in the end I am looking at the hardware and fuels that have flown to mars from departure to landing in order to look at a common core landing platform on which all mass of what we want at the surface to ride down on. Featuring legs, engines, pumping, fuel tanks and power storage to make it possible to deploy that item on the surface. With the plausible ability to refuel from what is brought to the surface as a means to get back to orbit..

Mars Exploration Current Missions

Image of all that have tried

https://en.wikipedia.org/wiki/List_of_missions_to_Mars

comparison of where nasa has sent missions and how much they cost

Successful rocket families that have been used to get to mars
https://en.wikipedia.org/wiki/Delta_II
https://en.wikipedia.org/wiki/Atlas_V

Here is the orbital paths taken

SpaceNut · 2018-08-03 21:28:41

https://www.universetoday.com/7024/the- … ed-planet/

Apollo lunar lander weighed approximately 10 metric tons, a human mission to Mars will require three to six times that mass, given the restraints of staying on the planet for a year.

Cargo only not the actual ship mass which puts it way up when we run the numbers.

http://marspedia.org/index.php?title=Landing_on_Mars

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.

I do not believe that number is correct.

http://sites.nationalacademies.org/cs/g … 083264.pdf

slide 6 shows the progression of the aeroshell that has landed mars crafts so far.

slide 21 shows where we are with fuels and engines

https://www.compositesworld.com/article … rs-landing
lots of material details as to making the landing a sucess

Aerothermodynamic Design of the Mars Science Laboratory Heatshield
https://ntrs.nasa.gov/archive/nasa/casi … 024218.pdf

Slide 5 heatshield data

SpaceNut · 2018-08-04 19:46:58

space.nss.org/media/Access-To-Mars.pdf

109 pages that covers quite a bit of information....

links on page 38

DEPOTS
• The Case for Orbital Propellant Depots: http://www.slideshare.net/jongoff/sa08-prop-depotpanel-
jon-goff
• Space Gas Station Would Blast Huge Payloads to the Moon:
http://www.popularmechanics.com/science … ws/4224660
• On-Orbit Propellant Resupply Options for Mars Exploration Architectures:
http://www.ssdl.gatech.edu/papers/confe … 1.1.01.pdf
MARS EDL
• High Mass Mars Entry, Descent, and Landing Architecture Assessment:
http://www.ssdl.gatech.edu/papers/confe … 9-6684.pdf
• Development of Supersonic Retro-Propulsion for Future Mars Entry, Descent, and Landing
Systems: http://www.ssdl.gatech.edu/papers/confe … 0-5046.pdf
• Fully-Propulsive Mars Atmospheric Transit Strategies for High-Mass Payload Missions:
http://www.ssdl.gatech.edu/papers/confe … 9-1219.pdf
• A Concept For The Entry, Descent, And Landing Of High-Mass Payloads At Mars:
http://www.ssdl.gatech.edu/papers/confe … D2.3.9.pdf
• Mars Exploration Entry, Descent and Landing Challenges:
http://www.ssdl.gatech.edu/papers/confe … 6-0076.pdf
• Sizing of an Entry, Descent, and Landing System for Human Mars Exploration
http://www.ssdl.gatech.edu/papers/confe … 6-7427.pdf
• Atkinson, Nancy: http://www.universetoday.com/2007/07/17 … ttinglarge-
payloads-to-the-surface-of-the-red-planet/
• Design of an Entry System for Cargo Delivery to Mars, Thompson, Robert, et al,
http://www.ssdl.gatech.edu/papers/confe … ompson.pdf

pg 66 suggest drop tanks.

pg 67 memories of the DC-X

pg 69 goes in to engine count and placement

pg 74 is the layout of 125 t ssto mars ascent 70 t decent landers
pg 98 mass breakdowns

SpaceNut · 2018-08-05 18:23:23

another thing that protects those in the capsule will be the protective coating.
https://www.nasa.gov/mission_pages/stat … ating.html

This protective thermal control coating, developed by Alion Science and Technology Corp., based in McLean, Va., made its bright appearance again with the March 1 launch of SpaceX's second commercial resupply mission. Named Z-93C55, the coating was applied to the cargo portion of the Dragon to protect it from the rigors of space.
Z-93C55 is a two-part system consisting of a pigment and a binder solution. Special additives enhance electrical conductivity without affecting thermal control properties, so the cured coating can handle high temperatures and survive the stresses of launching.
"The coating is actually an improved version of our Z-93P coating, which has had a long history in the aerospace industry," said Kenny. "It was used on Apollo missions, the station's radiators and many other missions. Z-93C55 is a thoroughly tested and qualified material, having gone through extensive testing in space simulation chambers and experimental missions."

SpaceNut · 2018-08-05 22:36:15

http://www.spacex.com/sites/spacex/file … tSheet.pdf

normal truck is 2.3 m and the extended version is 4.3 m tall

14 m3 (490 cu ft) payload volume and payload volume increases to 34 m3 (1,200 cu ft).

Its solar arrays produce a peak power of 4 kW

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