Smallest Human Ascent or Descent Lander for Mars Or Earth

SpaceNut · 2018-12-09 21:43:33

Numbers from the Insight lander

Overall
Total mass in cruise: 694 kg (1,530 lb)[3]
Lander: 358 kg (789 lb)[3]
Aeroshell: 189 kg (417 lb)[3]
Cruise stage: 79 kg (174 lb)[3]
Propellant and pressurant: 67 kg (148 lb)[3]
Relay probes flew separately but they weighed 13.5 kg (30 lb) each (there were 2)[3]
Lander specifications
Lander mass (Earth weight): 358 kg (789 lb)[3]
Mars weight (0.376 of Earth's)[55] : 134.608 kg (296.76 lb)
About 6.0 m (19.7 ft) wide with solar panels deployed.[3]
The science deck is about 1.56 m (5.1 ft) wide and between 0.83 and 1.08 m (2.7 and 3.5 ft) high (depending on leg compression after landing).[3]
The length of the robotic arm is 2.4 m (7.9 ft)[3]
Tilt of lander at landing on Mars: 4 degrees[56]
Power
Power is generated by two round solar panels, each 2.15 m (7.1 ft) in diameter and consisting of SolAero ZTJ triple-junction solar cells made of InGaP/InGaAs/Ge arranged on Orbital ATK UltraFlex arrays. After touchdown on the Martian surface, the arrays are deployed by opening like a folding fan.[57]
Rechargeable batteries[58]
Solar panels yielded 4.6 kilowatt-hours on Sol 1[59]
Each panel is 7 feet in diameter (2.2 meters) unfurled

elderflower · 2018-12-10 03:50:07

Lander mass is 358 kg. Earth weight of that is 3512 Newtons. At mars it still has mass 358 kg, but its weight is approx. 1335 N.
I do wish people would stop mixing up two different concepts!

RobertDyck · 2018-12-10 14:00:00

SpaceNut wrote:

ADEPT is a foldable device that opens to make a round, rigid heat shield, called an aeroshell.

GW Johnson wrote:

I'm glad folks are really working on this. It will be a game-changer at Mars, and perhaps elsewhere. But, those folks have a long way to go before this technology is read-to-apply.

Technology does not develop itself. Robert Zubrin and his partner David Baker included this with the original Mars Direct, developed in the last quarter of 1989 and first half of 1990, presented to NASA in June 1990. It was ready for testing at that time, but unfortunately looks like no progress since then. This is technology at least 28½ years old, probably older. It won't develop itself, it has to be tested by an unmanned explorer to Mars before committing human lives to it. Do it, and do it now! That test is already 28½ years overdue.

SpaceNut · 2018-12-14 18:00:31

I got wonder how to beef up the insight platform to be able to place a beam inflatable with internal stucture and controls that would expand inside when the unit was inflated. I would place a payload shroud around the outside to pertect the inflatable for landing and then on the first space walk split that shell and then after inflation expand the inner control deck putting in drop in panels to take up the open area.

SpaceNut · 2018-12-15 21:00:47

http://exrocketman.blogspot.com/2018/09 … -mars.html

Retro-propulsion is necessary from the moment hypersonics end at about 0.7 km/s speeds. Altitudes are on the order of only 5 km, so that impact is but seconds away, unless retropropulsion is on the order of 3-4 gees.

Mars%2Blanding.png

retro-propulsion lander that is one stage and not reusable, can have at most a payload fraction of about 50%, using long-term storable propellants, such as MMH-NTO.

Of course clean sheet would take longer and even thou the Red Dragon is out as well.
The crewed Dragon we have not looked at yet but it may have promise and the cargo Dragon is run here http://exrocketman.blogspot.com/2017/03 … -data.html

kbd512 · 2018-12-16 13:30:21

SpaceNut,

After watching "The Everyday Astronaut's" (a YouTube personality who is dedicated to space exploration technology) interview Rocket Lab's CEO, Peter Beck, something finally crystallized in my mind about how our lander and ascent vehicle should be powered. The descent / ascent engines should use electric turbo-pumps, or e-turbos, as I like to call them.

Rocket Lab's Rutherford LOX/RP1 engine delivers 22kN (5,000lbf) in a vacuum with a flight-demonstrated specific impulse of 333s, as compared to Merlin 1D Vac's 348s. I presume a LOX/LCH4 variant of their technology could deliver at least 340s. Their carbon composite propellant tanks have already delivered satellites to orbit for NASA, so it's not a guess as to whether or not their composite tank technology is LOX compatible. I can't think of a reason why it wouldn't be LCH4 compatible, too.

Mr. Beck said he's done work on aerospike engines for our government and that he did, in point of fact, lower the mass of the engine to the same as an engine with a conventional nozzle, but the tradeoff was increased complexity for the same performance achievable with a conventional nozzle. That's the best explanation I've heard to date about why it is that so few companies use them. That means there's not enough of a performance benefit to justify the added complexity for no actual performance gain. Furthermore, he said that the e-turbos can replicate some of the fuel-saving / altitude-compensating effects of the aerospike by varying e-turbo output and plume interactions.

Mr. Beck said the reason the engines are produced in America and the carbon composite propellant tanks are produced in New Zealand is that the competition for the aerospace composite manufacturing talent pool in America to make the composites is extreme, whereas employs 1/3rd of New Zealand's carbon composite manufacturing talent pool at cost-competitive labor rates. However, he also said that launch cadence may dictate making some propellant tanks in America at some point in the future.

Mr. Beck said the reason they launch from New Zealand is that the Falcon Heavy launch delayed or cancelled 560 flights in America because there's so much competition for the airspace around the government-owned rocket ranges and they need to launch twice per week.

* I don't personally see hypergolic propellants as a major problem since we've successfully used them for so long, even if some others here have been vocal about this issue, but LOX/LCH4 propellants are electrically ignited and self-pressurized, so no NTO/MMH or COPV handling precautions are required if tank pressurization is provided by electrically heating the LOX/LCH4, also negating the requirement for an external piston engine tank pressurization pump as in ULA's IVF solution to the problem

* Rutherford engine is 35kg on Earth or 13.3kg on Mars and doesn't look like Rube Goldberg's worst nightmare, meaning the engines are human transportable and shockingly simple to disassemble/reassemble with hand tools- something that is or should be a requirement for what should be obvious reasons related to maintenance

* Rutherford engines are nearly entirely 3D-printed in less than 24 hours, so if someone back on Earth discovers a major design problem with an engine component, then 3D printed engines are probably the only type of engines that can be slightly modified by the astronauts while they're on Mars by printing new components- this should never happen with proper testing, but if it does then you want a way to effect design changes

* the entire engine, to include the e-turbos, is software-controlled for infinitely variable thrust, which is required to soft land without using pulsed power and the associated sub-optimal specific impulse and extra wear & tear that results from rapidly starting / stopping any other type of engine, to include NTO/MMH engines

* the engine vibration and vehicle acceleration performance is the most benign of any satellite launch vehicle, meaning it's perfect for human descent / ascent vehicles

* e-turbos suck the propellant tanks completely dry without risking an explosion (ask GW why it's not possible to do this with a gas generator cycle without exploding turbo pumps, but my guess is that it has to do with cavitation / chugging, which will also quickly damage or destroy any other type of gas turbine aviation engine that I'm more familiar with), and that is how Electron delivers its payload with so little fuel

* the mass of the Lithium-ion batteries is already lighter than the residual propellant quantities required to shut down a conventional turbo-pump without risking an explosion, but this will only improve as battery energy density improves (gas generators are 50% efficient at best, but the Rutherford engine's e-turbo system is already 98% efficient and new CNT / Graphene / Stanene wiring materials will only improve the efficiency and mass of the e-turbo motor as advanced conductors and electric motor designs approach 100% efficiency)

* upon ascent, some of the mass of Lithium-ion batteries can be ejected from the ascent vehicle to economize on weight on the way to orbit, so the lander can use the batteries for descent, use the batteries on the surface of Mars for the surface exploration mission- a conventional turbo pump assembly is just dead mass that can't be used for anything else in a practical manner, perhaps by disassembly of the e-turbos and use to power CO2 compression pumps, and then use the e-turbos and batteries a final time for ascent- therefore, the batteries and pumps (not operated at full rated output on the surface, but something more reasonable for durability) are a way in which we can get enough variety of uses from all individual components to justify taking a lot of motors and batteries to Mars

* we really need the ability to use both the electric pumps and batteries for a variety of uses from power tools to LOX/LCH4 propellant plants to LOX/LCH4 hoppers that transport people or cargo to remote work sites

* a mobile Mars surface habitat with empty descent propellant tanks could transport LOX/LCH4 from a propellant plant to an ascent vehicle so as to physically separate the ascent vehicle, chemical plant, and surface habitat during propellant manufacturing and loading operations

* far less propellant production and electrical power required than if you sent Starship all the way to the surface of Mars and back to Earth

* the 16 KiloPower reactors I intended to ship could manufacture all return propellant required to achieve orbit in less than a month, so a true "load and go" ascent vehicle is a possibility, and Starship could be supported in future colonization efforts

* ability to explore multiple sites to confirm where sufficient water ice reserves are buried to support Starship and Mars base operations

Anyway, I'm just brainstorming here.

SpaceNut · 2018-12-16 16:26:43

Try number 2 for this clunker.....

A lot depends on are we going with recycling of the lander as insitu for the moon is questionable for all of the fuels for insitu creating and is leaning to LH2 for Insitu. For Mars insitu RP1 would be a long ways off and so would hydrogels for both to be created insitu.

RP1 and LCH4 are the same as for use as they are both atomized in the chamber with only the Lox ration being different.

Going with either for a descent stage is goo with the e-Pump but would stick with the hydrogels for ascent.

Solar batteries have been the choices for quite some time for landers for either location of destination.

kbd512 · 2018-12-18 16:12:24

SpaceNut,

Are you saying that we should use LOX/LCH4 or NTO/MMH for the descent stage and LOX/LH2 for the ascent stage?

Also, are you talking about using Methane to gel Hydrogen since you talked about using "hydrogel"?

The Methane product from the Sabatier reactors never comes out pure and contains significant amounts of Hydrogen contamination, so it might actually be easier to make "hydrogel", or gelled LH2, than pure LCH4 or pure LH2.

At 70% Methane gel weight, the gelled LH2 goes from a "very fluffy" 70kg/m^3 to a significantly thicker 177kg/m^3, providing almost 3 times the storage density. Isp of the resultant fuel is still over 410s (vac) using the standard 1,000psi chamber pressure figure and 94% of theoretical Isp for the actual Isp figure. Even if it mandates a propellant tank around the same size as the LOX/LCH4 vehicle, a substantially lower total tonnage of propellant products are required, which lower power requirements or cuts production time.

I'm not sure exactly how long RP-1 can be stored, but I wouldn't trust it to last more than a year. I thought everyone here was against using Hydrogen and that Methane was all the rage. If all you have is lots of water and enough electricity, then you can make rocket propellant. The "hydrogel" from the addition of Methane would be a nice performance enhancer. I like the storage temperature of the LCH4 since it matches up so well with LOX, but it's no better than RP-1 for Isp and the bulk density is still lower than RP-1 with densification.

Either way, using e-turbos provide a nice little power pack and battery for propellant production purposes. I'm not sure if I mentioned it or not, but there's no mixing of the oxidizer and fuel before the combustion chamber with the e-turbos, unlike a gas generator or staged combustion cycle. Why not use the e-turbo for ascent as well? At just over 1/3rd the gravity and almost no atmosphere, it's not going to take as long to achieve orbit because acceleration will be higher.

* with cryogenic propellants, no COPV's necessary, nor all the attendant hardware and failure modes that those create

* no ignition sources outside of the combustion chamber, unlike all gas generator or staged combustion cycles

* variable thrust using computer control of the e-turbo's electric motor speed to prevent cavitation inside the impeller assembly

* e-turbos suck the propellant tanks dry, so only actual loaded propellant weight, rather than variable residual propellant quantity needs to be computed to determine your dV (you either have enough propellant weight to fly or you don't, so no guess work is required here)

* ability to receive software and component updates from Earth, if absolutely necessary, and reload software or reprint major engine components (the only hardware that isn't printed consists of the electronics, engine fasteners, and electric motor shafts and windings)

* all components, including a completely assembled engine, are hand portable and hand-swappable using hand tools, meaning no cranes of any kind are required for engine maintenance (it may still be useful to have a hand crane for other reasons such as battery swap, but special engine removal jigs are not required)

* ability to construct an in-situ engine test stand to test individual engines for thrust performance profile

* ability to use magnetic particle inspection to re-qualify critical engine components, such as the thrust chamber and nozzle, for re-flight

* ability to measure electric motor performance on a dynamometer to assure that motor performance is nominal

That's a pretty long list of advantages for this specific application, mostly in the reliability department where we need all the advantages we can get. The only disadvantage is the weight of the battery, which is already less than the weight of the residual propellant quantity required to prevent a mechanical turbo pump from exploding upon engine shutdown. I'm not sure how much better we can make Lithium-ion batteries, but their storage capacity only continues to improve by a few percent per year. Maybe GW could talk about any mechanical turbo pump improvements that have come about, but I think that cycle is tapped out because thermodynamic efficiency limits it to a maximum of 50%.

SpaceNut · 2019-01-04 20:00:46

Land large cargo but go small for the crew to get us there so that we can reuse the lander multiple times from the landed resources.

SpaceNut · 2019-12-30 21:22:13

The only vehicle calculated to be able to land in its current state is a Dragon but its not able to return to orbit and would not be large enough for a crew of 6 to use but maybe for 1 it might be large enough when preloading more vehicles at the same site.

SpaceNut · 2019-12-30 21:59:19

Solving the expendable lander and MAV trap

A miniature fuel production plant would have over two years to make the fuel needed, and would only need to make a single metric ton of fuel per month, or about 30 kilograms a day.

The mass of the plant would be a fraction of the mass of the 30 metric tons of propellant it would create, which could then be replaced with additional supplies and equipment for the mission.

http://sicsa.egr.uh.edu/sites/sicsa/fil … ptions.pdf

While we would want to make fuel on the moon the lander has a different requirement
http://www.tsgc.utexas.edu/archive/design/lander/

https://www.ulalaunch.com/docs/default- … 6-7284.pdf

http://strategic.mit.edu/docs/Dipl_hofsetter_2005.pdf

SpaceNut · 2020-04-04 18:00:09

Emergencies happen and escape under that senerio is seldom thought of for moon or for mars. Of course its only going to work if they have orbital platforms to escape to other wise you might be an endangered species...
https://en.wikipedia.org/wiki/Lunar_escape_systems
At a minimum we need space suits, surplus oxygen and water to make a rendevous possible to that orbital platform. I would think we would dawn a cone like structure to shield the crew men toward orbit on such a desperation effort to survive. Its just a signle fuel and go stage that has seats to strap into and not much else....

kbd512 · 2020-04-04 18:47:52

SpaceNut,

Does that mean putting a miniature space station in orbit around each of these planets we want to visit isn't such a dumb idea after all?

SpaceNut · 2020-04-04 19:19:07

Thats 2 cygnus extended length with solar panels with a docking airlock to enter into.

SpaceNut · 2024-02-25 17:37:18

I wonder if 40 mT is small?

New Mars Forums

Announcement

#226 2018-12-09 21:43:33

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#227 2018-12-10 03:50:07

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#228 2018-12-10 14:00:00

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#229 2018-12-14 18:00:31

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#230 2018-12-15 21:00:47

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#231 2018-12-16 13:30:21

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#232 2018-12-16 16:26:43

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#233 2018-12-18 16:12:24

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#234 2019-01-04 20:00:46

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#235 2019-12-30 21:22:13

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#236 2019-12-30 21:59:19

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#237 2020-04-04 18:00:09

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#238 2020-04-04 18:47:52

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#239 2020-04-04 19:19:07

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

#240 2024-02-25 17:37:18

Re: Smallest Human Ascent or Descent Lander for Mars Or Earth

Board footer