Forty 40 Ton Mars Delivery Mechanism

GW Johnson · 2024-02-22 10:29:14

What I did was size a 1-shot lander and transfer stages (reusable or nonreusable) to send it to Mars. The 1-shot lander has to drop its 1-shot thin ablative heat shield to fire up its landing engines and unfold its legs.

If instead equipped with a segmented low-density ceramic heat shield that folded out of the way, it could be refueled with propellants made on Mars (I did my design with storables NTO-MMH!!!!) and flown again. There is no concept for how to make those propellants on Mars, though. And the loaded-with-cargo dV capability was quite low at 1.05 to 1.4 km/s, in order to put 40 tons of cargo into a vehicle massing about 80 tons at Mars entry. With that low a dV capability, I do not know what it could be reused for. What's the point of flying unloaded?

I did it with storables because I am definitely NOT a believer in long term storage of cryogens without serious or even fatal evaporation losses, unless substantial cooling power is available, which is also added mass. LCH4 is not as bad at evaporating and leaking as LH2, and it is worse than LOX, but they all have evaporation loss rates, even from a Dewar.

The reusable transfer stage design that used LOX-LH2 used the header tank construction within main tanks, to make the vented main tank into essentially a Dewar outer vessel around the inner header vessels. That buys you more time and a lower refrigeration requirement for months or years in space with cryogens. The reusable transfer stage would spend about 3.5 years in space before it could enter LEO for recovery. It is one hell of a risk to take, to get a reusable design.

To my knowledge there are no ready-to-use long-term re-liquifaction technologies available, to make years in space possible with LH2 or even LCH4. There are only lab benchtop demo toys. There's light-years' of development effort between a bench-top toy and a real, usable technology. I know that is an unpopular thing to say, but it is VERY TRUE!

Probably the most ready-to-build-and-test portion of my study is the lander and the non-reusable transfer stage. No long-term cryo storage was involved. Only departure from LEO was with LOX-LH2. The rest was all storables.

The lander would be sitting there to salvage. It would contain some amount of unused storable propellant that could be recovered and used elsewhere. The transfer stage would crash after suffering severe entry heating damage. It would not be salvageable in any sense.

GW

Last edited by GW Johnson (2024-02-22 10:33:51)

Calliban · 2024-02-22 11:35:07

Some potential long-term storable liquid propellants that we can talk about making on Mars.

Fuels: Propane, Butane, methanol, dimethyl ether. Methanol can be synthesised using using the same chemical reactor technology used to make methane. The catalysts are different and the CO and H2 are fed into the reactor in different proportions. Dimethyl ether has a vapour pressure similar to propane at room temperature. It is made from a condensation reaction between two methanol molecules over a catalyst. Any methanol producing sabateur reaction will also produce DME. But higher DME selectivity can be designed for.

Oxidants: LoX, HNO3, N2O, NO2, H2O2, F2. LOX and F2 give the highest exhaust velocity. F2 is so toxic that it just isn't a serious contender. Concentrated nitric acid is a metastable liquid and is 40% denser than water. But is very corrosive. Even 316SS has a limited life exposed to it. Under standard conditions, it will evolve NO2, which will form a gas over its surface. H2O2 has stability issues that can be reduced via chilling. I have heard of N2O (laughing gas) used as oxidiser. Energy density is reduced and from memory, there were problems with ignitability. But it is storable as a saturated liquid at room temperature and this is its selling point. I don't know much about NO2, aside from its toxicity.

Last edited by Calliban (2024-02-22 11:39:00)

SpaceNut · 2024-02-23 20:05:07

Here is the Marco polo topic first link for fuel manufacturing lander for mars.

https://kiss.caltech.edu/workshops/isru … anders.pdf

SpaceNut · 2024-02-25 16:55:26

cargo 1 way can go to orbit if we wanted, to the surface to support and for cyclical orbital if that's the thing we are looking for, but we need to know what we are sending as each items has a different density to volume.

With the ship characteristic nailed down we can then determine what goes, how much and how long it can support for any of the three uses.

tahanson43206 · 2024-02-25 20:50:45

For SpaceNut re #54

You don't ** have ** to say something just to fill a post.

This topic is about plans for landing 40 tons on Mars. Your post #53 with the link to just such a plan is a good fit for this topic.

Your hand waving in Post #54 has nothing to do with the topic, and it is not needed. No one who reads this topic to learn about landing a 40 ton payload on Mars will get anything out of Post #54.

If you want to help the topic, please do as you did with the link to the article you showed us in Post #53.

(th)

SpaceNut · 2024-02-26 19:34:41

Hand waving for shuttle was ignored and that leads to failure.

volume metrics at work

img.jpg?width=1200&height=628

A softball measures between 11.88 and 12.13 inches in circumference and weighs between 6.25 and 7.00 ounces;
a baseball measures between 9.00 and 9.25 inches in circumference and weighs between 5.00 and 5.25 ounces.

Both are under 10 ounce for payload not broken but a ship designed with small diameter means just a baseball, but we need softballs which cannot ship which means altering the design.

Design is more than engines, fuel, oxidizer, tanks, landing support to surface items as this requires stated inside cargo area dimensions. As you need other Craft support items in communications, telemetry, sensors, power sources to keep craft functioning ant that is not nuclear. Then again, a nuclear option for the fuel making plant is something that can be swapped in later.

My posts give these in those documents

The cargo bay of Starship is approximately 650 cubic meters in volume. Sure it can have 100 to 150 mT at one time of things we know we need for the crewed landing to follow. Pioneer Astronautics demonstrated a reactor capable of producing 1 Kg a day of methalox fuel from hydrogen and carbon dioxide while consuming a power of 700W. For 710 tons in 400 days that is 1.89 MW. (Zubrin et al., 2013)
Assuming 400 days to produce the 710 tons of fuel needed, 352 tons of water (for electrolysis) and 1.89 MW of power would be needed. Using the methods and assumptions detailed in section 4.3 (including a 20% margin for safety), the solar infrastructure would be:
• 229.2 tons in mass.
• 3437.4 cubic meters in volume.
• 57290.1 square meters in area.
The deployment would require 5 to 6 Starships (volume constrained) and significant deployment operations and maintenance.

We all thought that this many were to many and yet it takes that many per each ship to refuel before leaving earth to go to mars. My topic for getting the fuel from the ground gave those same quantities of ships to land on mars. Of course, crew support required 2 more ships to land with them to be able to do a 500-day surface mission for 100 people.

. Power remains one of the most significant challenges of a Mars mission architecture that accounts for the return of the astronauts. As with issue 1, failure in this area would result in loss of crew.

Of course, nuclear can go out as cargo that is not active until deployed in any of the ships that have room for it once on the surface with crew setup.

Then there are the needs for landing area conditions to be just right as we have seen with the moon landing failures.

You can see that the second ship is less capable of mass, and it does not have the volume to be able to do so.

The much smaller ship is more like the 40 mT design to mars but even that is a much smaller crew size and with other goals. Caravel class naval ships on Christopher Columbus’ first voyage to America

Pinta. Cargo on board (Total: 26.5t):
• Water tanker rover. Mass budget: 3.5t
• Food, water and supplies. Mass budget: 6t
• 20 KW of solar panels: Mass budget: 2t
• Scientific equipment, batteries, carbon dioxide electrolyzers and other. Mass budget: 15t
Santa María (uncrewed) Cargo on board (Total: 120t):
• Cranes, batteries and all operating equipment. Mass budget: 8t
First Martian habitat, including crew quarters and a common area. Mass budget: 34t
• Pressurized Rover. Mass budget: 10t
• Water extraction/ice mining machinery. Mass budget: 20t
• Extra water and supplies: Mass budget: 12t
• Additional solar panels/fission reactors. Mass budget: 36t

GW Johnson · 2024-02-27 09:23:44

Try looking at it this way. Starting from LEO, it takes about 3.7-3.8 km/s worth of delivered dV to cover the departure onto a min-energy Hohmann transfer to Mars (closer to 4+ if faster) plus two course corrections along the way, which dV has to come from something.

Once you are there, you have two simple options, and you have to come up with more complicated schemes if you dislike both simple options. Option 1: decelerate into LMO where the surface is within easy reach. That cost is variable due to Mars's orbital eccentricity, but is on the order of 2 km/s. Option 2: enter Mars's atmosphere directly for as much aerobraking as you can achieve, and then do a rocket-powered landing because the air is too thin, and your end-of-hypersonics altitude with a big object is too low, for getting any help out of chutes. That dV is on the order of 1 km/s, which makes this the cheaper option in terms of dV-to-ship.

Now, logistics:

With option 1, if you want to take the cargo down to the surface, then you have to go and get it somehow. It's complicated, but crudely speaking, your orbit taxi will need around 5 km/s unladen dV capability to cover launch into orbit from the surface, plus a maneuver kitty to rendezvous. Then it will need about 1.1-1.2 km/s worth of dV capability fully laden with the shipped cargo, to cover deorbit, and the final rocket-powered landing. That's 3.7-3.8 km/s to put it on course for Mars, about 2 km/s to enter LMO, then about 6.1-6.2 km/s more to go and get it with an orbit taxi to the surface. You cannot add these dV's, because the weight statements are all different, but the clear implication is that you will use a lot of propellant doing this, just to get your cargo onto the surface of Mars, because the incorrect summing gets you 11.8-12.0 km/s. That's true whether some of the propellant is made at Mars, or not.

With option 2, the cargo is already on the surface! It took 3.7-3.8 km/s to put it on course for Mars, and about 1 km/s to land it there. Again, you cannot add those dV's because the weight statements are different, but the clear implication is that takes a lot less propellant, to put your cargo onto the surface of Mars, because the incorrect summing gets you only 4.7-4.8 km/s.

Exactly how you go about doing either option can make the actual propellant quantities vary, but only by percentages, not factors of 2+! And THAT is exactly why SpaceX decided it could use Starship with big payloads to go to Mars with direct entry and landing, but did not really consider going to LMO instead. But, in so choosing, they gave up any surface scouting for the best place to land, instead just having to pick a spot and go there blind to any real ground truth.

Those are the kinds of tradeoffs you have to make to choose between the two options, or considering any more complicated alternatives (each of which is likely to have an even higher incorrectly-summed dV). Those are the kinds of tradeoffs that can get crews killed, because the disparity between remote sensing and ground truth is still non-zero, despite what everyone so desperately wants to believe.

The kind of "ground truth" I refer to is subsurface: what resources are buried there, and how best do we recover them, and how best can we make use of them? Wrong answers to any of those questions can kill. None of the landers and rovers we have ever sent to Mars can answer questions like that!

Myself, I prefer option 2 (direct entry and landing) for bulk cargo and big hab items, for a real base. But I want to see some surface/subsurface scouting done, before I have to pick a final site for building a base. That kind of scouting is better done as multiple short landings at multiple sites, from a single mission based in LMO. The bad news: you cannot really do both in the same mission!

And if you just mount a big base-building mission to a site picked from remote sensing, you are betting your crew's lives that it is the right site, that your remote sensing was 100% accurate. The history of real ground truth vs remote sensing refutes that assumption.

GW

Last edited by GW Johnson (2024-02-27 15:48:24)

SpaceNut · 2025-09-22 06:29:03

All start with the assumption that we have limited lift from earths gravity well. Also that we have trouble with landing tonnage as well.

Mars Direct 3 is a Mars mission architecture developed by Miguel Gurre

Mars Direct; Mars Semidirect; Design Reference Mission. Need Updating?

Starboat

AI Overview

For a 26-metric-ton (mt) spacecraft, the estimated fuel mass required for a propulsive landing on Mars is approximately 40 to 60 mt, bringing the total landing vehicle mass to 66–86 mt. This relies on significant aerodynamic deceleration in the Martian atmosphere before the final propulsive braking maneuver. The final mass varies based on the engine's efficiency and the exact landing trajectory. Fuel requirements for Mars landing Several factors influence the fuel mass needed to land a 26 mt spacecraft on Mars: Deceleration strategy: A Mars landing is a complex process known as Entry, Descent, and Landing (EDL). Due to Mars's thin atmosphere, a propulsive-only landing is inefficient. Instead, spacecraft typically use a combination of methods, including:A protective aeroshell and heat shield to withstand atmospheric entry at high speed.A parachute to provide further slowing.A final rocket-powered braking phase for the precision touchdown.The 26 mt figure would refer to the mass of the final lander after shedding the heat shield and parachute system.Engine specific impulse (\(I_{sp}\)): The efficiency of the rocket engine is a critical factor, described by the specific impulse (\(I_{sp}\)).Higher \(I_{sp}\) engines, like those using liquid hydrogen and oxygen, provide more thrust per unit of fuel, but hydrogen is difficult to store.Methane and oxygen (\(\text{CH}_{4}/\text{O}_{2}\)) offer a lower \(I_{sp}\) but are easier to store and can be manufactured on Mars using in-situ resource utilization (ISRU). This trade-off is central to Mars mission architecture.Delta-V (\(\Delta v\)): The amount of total change in velocity required for the propulsive landing phase is roughly 3.8 km/s from orbit to the surface if parachutes aren't used, but is less when combined with aerodynamic braking. A higher \(I_{sp}\) reduces the propellant mass needed to achieve this \(\Delta v\). Architectural approaches Planetary mission planners have developed different architectures to manage the challenge of large-scale Mars landings: Heavy landers: An analysis of Mars landing vehicles for future human missions found that a total initial mass of 73.0 mt was needed to land a 10 mt payload, while a 25 mt payload (closer to your scenario) required an even larger vehicle. A significant portion of this mass would be propellant for the final descent phase.SpaceX Starship: SpaceX's Starship is designed to land payloads of 100 mt or more using a methane/oxygen engine system. In this architecture, the Starship tanker refuels the Mars-bound ship in Earth orbit, making propellant for landing part of a larger, refueled system.In-situ resource utilization (ISRU): Some mission architectures propose landing an initial vehicle with a fuel-manufacturing plant. This plant would use Martian resources (water ice and atmospheric \(\text{CO}_{2}\)) to produce methane and oxygen propellant for a later landing or for the return trip, significantly reducing the mass that needs to be transported from Earth/

SpaceNut · 2025-09-22 14:28:37

so the mars ship plus payload would create a propulsive landing

To land a 66 metric ton (mt) craft on Mars using propulsive methods, a significant amount of methane (\(CH_{4}\)) and oxygen (\(O_{2}\)) propellant is needed to perform the required braking maneuvers. A full propulsive landing is not necessary due to the thin Martian atmosphere, which can provide initial braking through aerobraking. A combination of aerobraking and retro-propulsion is the most efficient method for large spacecraft. Here is a breakdown of the fuel estimate and the assumptions in BBCode. [center][size=5]Propellant for a 66 mt Mars Propulsive Landing[/size][/center] SummaryLanding a 66 mt craft on Mars is most efficiently accomplished by using a combination of aerobraking and a final propulsive burn. For a 66 mt craft, estimates suggest around 20-30 mt of methalox propellant would be required for the terminal landing phase. Mission Parameters
[] Craft Mass (Dry): 66 mt (Metric Tons)[] Propellant Type: Methane (\(CH_{4}\)) and Liquid Oxygen (\(LOX\) or \(O_{2}\)), commonly known as "methalox".[] Landing Strategy: Aerobraking followed by supersonic retro-propulsion.[] Specific Impulse (\(I_{sp}\)): Approximately 350-370 seconds for a methalox engine in vacuum, but lower during a landing burn in the atmosphere.
Delta-V (\(\Delta v\)) Required: Estimated to be in the range of 400-500 m/s for the final, propulsive landing burn, after aerobraking has significantly slowed the craft.
Propellant Calculation (\(M_{p}\))The mass of propellant required is determined using the Tsiolkovsky Rocket Equation. \(M_{p}=M_{wet}-M_{dry}\) \(M_{wet}=M_{dry}\cdot e^{\frac{\Delta v}{I_{sp}\cdot g_{0}}}\) Where:
[] \(M_{wet}\) is the initial wet mass of the craft (including propellant).[] \(M_{dry}\) is the final dry mass of the craft (payload + empty tanks).[] \(\Delta v\) is the change in velocity.[] \(I_{sp}\) is the specific impulse.[] \(g_{0}\) is standard gravity (9.81 \(m/s^{2}\)).[] \(e\) is the mathematical constant (approximately 2.718).
Assuming a \(\Delta v\) of 480 m/s and an average effective \(I_{sp}\) of 300 seconds for the atmospheric landing burn: \(M_{wet}=66mt\cdot e^{\frac{480}{300\cdot 9.81}}\) \(M_{wet}\approx 66mt\cdot e^{0.163}\) \(M_{wet}\approx 66mt\cdot 1.177\) \(M_{wet}\approx 77.68mt\) \(M_{p}=77.68mt-66mt\) \(M_{p}\approx 11.68mt\) Key Considerations
[] Mixture Ratio: Methalox engines use a mixture ratio (oxidizer to fuel) of around 3.5. For 11.68 mt of propellant, this equates to roughly 9.17 mt of \(LOX\) and 2.51 mt of \(CH_{4}\).[] Margin and Boil-off: Space missions require significant margins for unexpected events. Propellant boil-off during the long transit to Mars must also be accounted for by loading extra fuel.
Atmospheric Conditions: The actual performance of the engine will vary with atmospheric pressure, which affects the effective \(I_{sp}\) during the landing burn.
Estimated Propellant Breakdown
[] Total Propellant Mass: 11.68 mt (minimum calculated for a 480 m/s \(\Delta v\)).[] Total Methalox Mass (with Margin): An operational mission would carry more, likely in the range of 20-30 mt, to be safe.[] Methane (\(CH_{4}\)) Required: ~ 5-7.5 mt[] Liquid Oxygen (\(LOX\)) Required: ~ 15-22.5 mt

SpaceNut · 2025-09-22 17:13:54

Assuming you are referring to a Starship-class vehicle with a mass of 66 metric tons (mt) and three Raptor engines, a propulsive landing on Earth would require approximately 3–6 tons of liquid methane and liquid oxygen propellant. This is based on the following factors:
Vehicle mass and engine thrust:
The mass of 66 mt is the dry weight of the spacecraft, excluding propellant. The final mass during landing would be higher, including any remaining payload and the three Raptor engines. Each Raptor engine is capable of at least 230 tons of thrust, giving a three-engine cluster significant propulsive capability.
Delta-V for landing:
A propulsive landing on Earth requires a change in velocity (\(\Delta v\)) to transition from atmospheric braking to a final, controlled vertical descent. This terminal velocity is typically around 50–100 m/s.
Rocket equation and exhaust velocity:
You can estimate the required propellant using the Tsiolkovsky rocket equation:\(m_{fuel}=m_{final}\cdot (e^{\Delta v/v_{exhaust}}-1)\)For a Raptor engine, the exhaust velocity (\(v_{exhaust}\)) is about 3,500 m/s (from a specific impulse of 350s). The final mass (\(m_{final}\)) is the spacecraft's mass just before the final landing burn.
Propellant mass estimation:
Assuming a 70 mt final mass (including a small payload) and a 100 m/s burn:\(m_{fuel}=70\cdot (e^{100/3500}-1)\approx 2\ tons\)SpaceX's own internal analysis has produced slightly higher figures, around 6 tons, based on simulations and real-world results. This higher figure accounts for additional fuel reserves, engine gimballing, and safety margins.
Breakdown of the landing process
A propulsive landing with this type of vehicle and engine setup would include these phases:
Header tanks:
The fuel for the landing maneuver is drawn from smaller header tanks, which contain a fraction of the total propellant. This is more reliable and prevents the main tanks from sloshing. It also ensures the engines have a steady propellant flow, a key factor in successful propulsive landings.
"Belly-flop" maneuver:
During atmospheric reentry, the spacecraft enters a belly-flop orientation, using its body and control flaps to slow down. This reduces the need for propulsive braking.
"Landing flip" maneuver:
Shortly before touching down, the engines ignite and perform a flip maneuver to orient the spacecraft vertically for landing.
Precision and controls:
The final landing requires precise throttling and gimballing of the engines to counteract gravity and achieve a soft touchdown

SpaceNut · 2025-09-23 15:47:03

Desired is the ability to land a ship on mars surface in the 40mt neighborhood after using airbreaking and propulsive landing.

Of course there is some wiggle room.

Launch of ship to mars from earth orbit so how to have a starship that is 120 mt dry loss that much mass does no seem possible.

Assuming a launch profile based on SpaceX's Starship system, which uses Raptor engines and a methane CH4 propellant, a 40-tonne payload for a Mars transfer mission requires approximately 676 metric tonnes (mt) of propellant for the transit portion alone. However, the total fuel needed is significantly higher when considering all mission phases. The complex mission requires multiple propulsive burns and in-orbit refueling. A simplified estimate for a mission to transport a 40-tonne payload to Mars would involve the following phases: Launch from Earth to low-Earth orbit (LEO).On-orbit refueling in LEO.Trans-Mars Injection (TMI) burn to escape Earth's orbit.Entry, descent, and landing (EDL) at Mars.

To determine the methane CH4 and liquid oxygen LOX fuel requirement for a 40 metric ton (mt) ship landing on Mars, several factors must be calculated. The key steps are determining the change in velocity (delta-v) needed for the landing, applying the Tsiolkovsky rocket equation, and calculating the specific masses of methane and oxygen based on the Raptor engine's characteristics.
Assumptions for this calculation Initial ship mass: 40 mt (40,000 kg).
This is the dry mass of the ship plus any payload, but before the addition of landing propellant.Propulsive landing only: The calculation assumes no aerodynamic braking or very minimal atmospheric drag assistance. However, SpaceX's actual Starship landing profile uses substantial aerodynamic braking, which significantly reduces the propellant needed.
Raptor engine specific impulse (Isp):
An average vacuum Isp of 380 seconds is assumed for the Raptor Vacuum engines, which is more representative of a landing scenario than the sea-level variants.
Raptor engine mix ratio:
The Raptor engine uses liquid methane and liquid oxygen, typically at a mass ratio of 1:3.6 (methane to oxygen)
For a fully propulsive landing without using atmospheric drag, a Mars landing requires a delta-v Delta V of approximately 4.5 to 6 km/s. If the ship uses supersonic retro-propulsion with atmospheric braking, the required propulsive delta-v is much lower, possibly as low as 75 m/s, although this is very dependent on the entry velocity.
The total propellant mass is approximately 5.5 mt. We use the Raptor engine's mix ratio of 1:3.6 for methane CH4 to oxygen LOX by mass.
Based on the assumptions, the approximate fuel requirements for a 40 mt ship using two Raptor engines for a propulsive Mars landing would be:
Total propellant: 5.5 mt
Methane CH4: 1.2 mt
Liquid oxygen LOX: 4.3 mt

Same ship returning to earth from mars surface.

For a 40-metric-ton (mt) ship returning to Earth from Mars using two Raptor engines, the estimated propellant requirement is approximately 194.2 mt of methalox (liquid methane and liquid oxygen). This calculation assumes a propellant depot is available in Mars orbit and that the engines are vacuum-optimized Raptor variants.
This is an estimate based on the Tsiolkovsky rocket equation and can be affected by factors such as mission profile and gravity losses.
This estimate relies on three key parameters:
Mass of the spacecraft: 40 mt.
This is the "dry mass" m_{f} in the rocket equation, representing the ship, cargo, and all components except for the propellant.
Specific impulse Isp of the engines:
The vacuum-optimized Raptor engines (RVac) have a specific impulse of approximately 380 seconds.
Delta-v required for the maneuver:
The delta-v needed to launch from the Martian surface to a trans-Earth injection (TEI) trajectory is approximately 4.27 km/s.
Adjusting for a 20% methane/80% oxygen mix Raptor engines use a methalox propellant mix, which consists of approximately 20% methane (fuel) and 80% liquid oxygen (oxidizer) by mass. The total propellant mass is the combination of the fuel and oxidizer.
Total propellant mass m_{p}: 86 mt
Fuel (methane) mass: 0.20 * 86 mt = 17.2 mt
Oxidizer (liquid oxygen) mass: 0.80 * 86 mt = 68.8 mt
Assumptions and other considerations The calculated fuel requirement is a theoretical minimum based on the ideal rocket equation. Several factors can increase the actual fuel mass needed:
Atmospheric drag on Mars:
While the Martian atmosphere is thin, it can cause some drag during ascent, requiring a small amount of extra propellant.
Gravity losses:
The effect of gravity pulling against the rocket during its ascent and burn means the rocket must use additional propellant to counteract this force. The 4.27 km/s figure already accounts for typical gravity losses, but actual losses can vary.
Engine inefficiencies:
The Isp value of 380s is an ideal figure, and the engine may not achieve this perfectly throughout the burn.Vehicle mass variations: A fully fueled ship is heavier and less agile than one with less fuel.

So to make this work starship is now a two or three stage rocket that expends takes are we go.
one crew type and a cargo as its not leaving.

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#51 2024-02-22 10:29:14

Re: Forty 40 Ton Mars Delivery Mechanism

#52 2024-02-22 11:35:07

Re: Forty 40 Ton Mars Delivery Mechanism

#53 2024-02-23 20:05:07

Re: Forty 40 Ton Mars Delivery Mechanism

#54 2024-02-25 16:55:26

Re: Forty 40 Ton Mars Delivery Mechanism

#55 2024-02-25 20:50:45

Re: Forty 40 Ton Mars Delivery Mechanism

#56 2024-02-26 19:34:41

Re: Forty 40 Ton Mars Delivery Mechanism

#57 2024-02-27 09:23:44

Re: Forty 40 Ton Mars Delivery Mechanism

#58 2025-09-22 06:29:03

Re: Forty 40 Ton Mars Delivery Mechanism

#59 2025-09-22 14:28:37

Re: Forty 40 Ton Mars Delivery Mechanism

#60 2025-09-22 17:13:54

Re: Forty 40 Ton Mars Delivery Mechanism

#61 2025-09-23 15:47:03

Re: Forty 40 Ton Mars Delivery Mechanism

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