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#51 2024-02-22 10:29:14
- GW Johnson
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Re: Forty 40 Ton Mars Delivery Mechanism
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)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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#52 2024-02-22 11:35:07
- Calliban
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Re: Forty 40 Ton Mars Delivery Mechanism
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)
"Plan and prepare for every possibility, and you will never act. It is nobler to have courage as we stumble into half the things we fear than to analyse every possible obstacle and begin nothing. Great things are achieved by embracing great dangers."
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#53 2024-02-23 20:05:07
- SpaceNut
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Re: Forty 40 Ton Mars Delivery Mechanism
Here is the Marco polo topic first link for fuel manufacturing lander for mars.
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#54 2024-02-25 16:55:26
- SpaceNut
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Re: Forty 40 Ton Mars Delivery Mechanism
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.
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#55 2024-02-25 20:50:45
- tahanson43206
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Re: Forty 40 Ton Mars Delivery Mechanism
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)
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#56 2024-02-26 19:34:41
- SpaceNut
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Re: Forty 40 Ton Mars Delivery Mechanism
Hand waving for shuttle was ignored and that leads to failure.
volume metrics at work
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: 15tSanta 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
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#57 2024-02-27 09:23:44
- GW Johnson
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Re: Forty 40 Ton Mars Delivery Mechanism
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)
GW Johnson
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#58 2025-09-22 06:29:03
- SpaceNut
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Re: Forty 40 Ton Mars Delivery Mechanism
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?
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/
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#59 2025-09-22 14:28:37
- SpaceNut
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Re: Forty 40 Ton Mars Delivery Mechanism
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
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#60 2025-09-22 17:13:54
- SpaceNut
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Re: Forty 40 Ton Mars Delivery Mechanism
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
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#61 2025-09-23 15:47:03
- SpaceNut
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Re: Forty 40 Ton Mars Delivery Mechanism
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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#62 2026-02-16 09:10:24
- SpaceNut
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Re: Forty 40 Ton Mars Delivery Mechanism
bumping for all to find for delivery to payload of construction items.
The rockets landing on mars require the EDL systems but to change the equation for payload mass to not require more fuel for retro propulsion means changing the amount of atmospheric drag. These are just the small changes of the HIAD inflatable, Expandable heat shield, ADEPT umbrella structured fabric, and any other ballistic ballute type which can add just minimal mass but gain mass to the surface.
It does not change the core of the rocket or its landing legs or other structures used to make it set stable on the Martian surface. It is just how we cheat to get more mass to the surface with the same rocket.
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#63 2026-02-16 14:36:51
- SpaceNut
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Re: Forty 40 Ton Mars Delivery Mechanism
Here is a clean sheet design no specified fuels, engines ect.
delivery of 40 T payload to mars surface launched from earth rocket fresh design
See my thinkingYou’re speaking my language here—let’s sketch a clean, end‑to‑end architecture for “40 t to Mars surface” as a design target, not just a slogan.
High-level architecture
Segment Function Rough mass scale Key design choices
Earth launch Put full Mars stack in LEO ~250–350 t to LEO New heavy launcher or multi‑launch assembly
In-space stage LEO → TMI, maybe MOI ~80–120 t propellant Chemical or NTR, with or without aerocapture
Mars lander EDL + 40 t payload ~70–110 t at Mars entry Supersonic retroprop + aeroshell + maybe chute
Payload Delivered to surface 40 t Cargo, habitat, ISRU, etc.I’ll assume chemical propulsion and aerocapture/entry (no propulsive MOI), because that’s the most “buildable soon” path.
Step 1: Mars entry mass and lander sizing
For 40 t delivered to the surface, a realistic landed-mass fraction (payload / entry mass) for a big, powered‑descent lander is around 0.35–0.45 if you’re aggressive but not magical.Let’s pick 0.4 as a design target: ? entry ≈ 40 t 0.4 = 100 t
So:
Entry mass: ~100 t
Payload: 40 t
EDL hardware + structure + residuals: ~60 t
EDL concept:
Aeroshell: Rigid, lifting body or blunt cone, diameter ~10–12 m.
Entry: Aerocapture directly from hyperbolic arrival into a steep entry corridor.
Supersonic phase: Possibly a small ring‑parachute or none; rely mostly on lift + retroprop.
Powered descent: Methalox or hypergolic engines, throttling to soft land with 10–20% propellant margin.
Step 2: Trans-Mars injection mass
Assume we inject the 100 t entry vehicle directly onto a Mars transfer trajectory (no Mars orbit insertion burn—pure aerocapture/entry).For a Hohmann‑like transfer:
Δv (LEO → Mars transfer): ~3.6–3.8 km/s (including some margin)
Chemical stage Isp: ~360 s (hydrolox) or ~340 s (methalox)
Use the rocket equation with Isp = 360 s: Δ ? = ??? ?0 ln (?0 ??) 3,800 = 360 ⋅ 9.81 ⋅ ln (?0 100) ln (?0 100) ≈ 3,800 3,531.6 ≈ 1.08 ⇒
?0 100 ≈ ? 1.08 ≈ 2.94 ⇒ ?0 ≈ 294 tSo: Mass in LEO (lander + TMI stage + prop): ~300 t
Dry mass of TMI stage: say 15–20 t
Propellant: ~180–200 t
That’s a single 300 t‑to‑LEO stack or multiple launches assembled in orbit.
Step 3: Earth launch system
You now have two basic architectural choices:Single super‑heavy launcher
Capability: ~300 t to LEO (expendable or partially reusable).
Configuration: 2.5‑stage or 3‑stage, LOX/CH4 or LOX/LH2 core, strap‑on boosters.
Advantage: No on‑orbit assembly; simpler operations.
Disadvantage: Very large, expensive, and politically heavy.
Multi‑launch with orbital assembly/refueling
Example: 3–4 launches of ~80–120 t to LEO each.
Ops:
Launch lander dry + TMI stage structure.
Launch 2–3 tanker flights to fill the TMI stage.
Advantage: More incremental, uses a “medium‑super‑heavy” class instead of a monster.
Disadvantage: On‑orbit docking, prop transfer, schedule complexity.
For a “fresh design,” I’d lean toward option 2: a modular heavy launcher (~100 t to LEO) plus a standardized Mars cargo stack.
Step 4: A concrete “fresh” concept
4.1. Earth launcher
Name: “MHL‑100” (Modular Heavy Lifter, 100 t to LEO)Performance: 100 t to 400 km, 28° LEO, expendable; 70 t reusable.
Stages:
First stage: LOX/CH4, ~4,000–5,000 t liftoff mass, cluster of ~9–15 engines.
Second stage: LOX/LH2 or LOX/CH4, ~150–200 t propellant, 1–2 vacuum engines.
Reusability: First stage downrange landing; second stage expendable for Mars missions.
4.2. Mars cargo stack (per 40 t delivery)
Element A (lander):100 t Mars entry mass.
Integrated aeroshell + tanks + engines.
Payload bay sized for 40 t, ~6–8 m internal diameter.
Element B (TMI stage):
Dry mass: ~20 t.
Propellant: ~180 t LOX/LH2 or LOX/CH4.
Docking ports + prop transfer interfaces.
4.3. Launch sequence (3–4 launches)
Launch 1:Put lander (100 t) into LEO.
Launch 2:
Put TMI stage (dry, 20 t) + some propellant into LEO.
Dock to lander.
Launch 3 & 4:
Tanker flights delivering propellant to fill TMI stage to ~200 t.
Departure:
TMI burn sends the 100 t lander to Mars.
Stage is discarded or sent to disposal trajectory.
Arrival and EDL:
Direct entry, aerothermal braking, then powered descent.
40 t payload on the surface.
Step 5: Where this gets interesting for you
This 40 t delivery architecture becomes a standard “unit”:Every campaign window, you can send N × 40 t to the surface by repeating the stack.
The lander geometry can be standardized, and you can swap payloads: habitats, ISRU plants, bulldozers, tank farms, etc.
The MHL‑100 launcher and TMI stage are reusable design primitives for your broader Mars base architecture.
If you want, next step we can:
Lock in exact lander geometry (diameter, deck layout, tank placement, engine count).
Do a more precise mass budget for the EDL propellant and structure.
Or design a variant that can also bring a small ascent stage or sample return canister back to Mars orbit.
add retro propulsion, inflatable HIAD for mars EDL with landing legs
See my thinking
You’re dialing this exactly where it gets fun—let’s wrap the 40 t lander in an inflatable HIAD, then hang a proper retro‑prop/leg system off it.1. Target masses and fractions
Payload to surface: 40 tLanded mass (payload + structure + residuals): ~70 t
Entry mass (with HIAD + full descent prop): ~100 t (same design target as before)
So we’ll work with: ? entry ≈ 100 t ? landed ≈ 70 t
Descent propellant: ~25–30 t (depending on Isp and ignition altitude)
2. HIAD geometry and role
Concept: A rigid central body (lander) with an inflatable, stacked‑torus HIAD that deploys before entry.Central lander body:
Diameter: 6–7 m cylindrical core.
Height: 8–10 m including tanks and payload bay.
HIAD:
Inflated diameter: 12–14 m (gives you a big ballistic coefficient reduction).
Shape: 60–70° half‑angle, slightly lifting (offset CG for cross‑range and corridor control).
Mass: ~5–8 t (structure, TPS, inflation system).
Function:
Cuts peak decel and heating.
Reduces terminal velocity at retro‑prop ignition altitude.
Keeps the retro‑prop burn shorter and prop mass reasonable.
3. EDL sequence with HIAD + retroprop
Cruise & approach:Lander arrives on hyperbolic trajectory.
Attitude control aligns for entry; HIAD is stowed.
HIAD deployment:
At a few minutes before entry interface, HIAD inflates and rigidizes.
Central body is now “buried” behind the HIAD, engines and legs tucked inside the wake.
Atmospheric entry:
HIAD takes the heating and loads.
Modest lift via CG offset for corridor control.
Peak decel ~5–7 g, depending on entry speed and flight path angle.
Subsonic/supersonic transition:
You can skip parachutes entirely or use a small drogue; assume no main chute for simplicity.
At Mach ~2–3, you start preparing for retro‑prop.
HIAD jettison + retro‑prop ignition:
At a pre‑set dynamic pressure/altitude (e.g., 8–12 km), HIAD is cut loose on frangible joints and pushed away with springs.
Lander flips to engines‑down attitude.
Retro‑prop engines ignite and throttle to arrest vertical velocity.
Terminal descent and landing:
Guidance shapes a near‑vertical final approach.
Landing legs deploy during the last ~1–2 km.
Touchdown with ~1–2 m/s vertical velocity.
4. Retro‑propulsion sizing
Assume:Mass at ignition: ~80–90 t (you’ve burned off some TPS/consumables, but still have full descent prop).
Mass at touchdown: ~70 t.
Δv needed for powered phase: ~600–800 m/s (depends on how deep you ride the atmosphere).
Take Δv = 700 m/s, Isp = 340 s (methalox): Δ? = ??? ?0 ln (?0 ??) 700 = 340 ⋅ 9.81 ⋅ ln (?0 70) ⇒ ln (?0 70) ≈ 0.21 ⇒ ?0 70 ≈ 1.23 ⇒
?0 ≈ 86 tSo:
Descent propellant: ~16 t (plus margin → design for 20–25 t).
Thrust requirement:
Mars gravity:
?? ≈ 3.71 m/s2.For 90 t: weight ≈ 333 kN.
With 2× margin for decel: ~700 kN total thrust.
Engine layout:
4 engines × 200 kN each (800 kN total) around the base ring.
Deep throttling to ~20–30% for final hover/flare.
Gimbaled for attitude control; RCS for fine trim.
5. Landing leg system
With a 6–7 m core and HIAD jettisoned, you can use a wide‑stance leg system:Number of legs: 4 or 6; I’d pick 6 for redundancy and softer soil.
Deployed radius: 8–9 m footprint (legs extend beyond the core diameter).
Stroke: 1–1.5 m crushable/energy‑absorbing stroke.
Design loads:
Assume 1.5 g vertical at touchdown (including margin).
For 70 t landed mass: ~1,030 kN total → ~170 kN per leg (6 legs).
Geometry:
Legs hinged at the lower tank ring.
Folded up along the side of the core during HIAD‑covered phases.
After HIAD jettison and flip, legs deploy outward and downward.
6. Putting it all together as a “standard unit”
Mars cargo lander (HIAD + retroprop + legs):Entry mass: 100 t
HIAD: 12–14 m inflatable, 5–8 t
Descent propellant: 20–25 t methalox
Engines: 4 × 200 kN, Isp ~340 s
Landed mass: 70 t
Payload: 40 t
Legs: 6‑leg system, 8–9 m footprint
This slots cleanly into the earlier architecture: ~300 t in LEO (lander + TMI stage + prop) gives you 40 t on Mars surface with a reusable design pattern.
If you want to go deeper, we can next:
Lock exact HIAD diameter and torus stack vs entry mass and β.
Sketch a plan view of the lander base: engine ring, leg attach points, and payload bay footprint.
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#64 Yesterday 15:00:50
- SpaceNut
- Administrator
- From: New Hampshire
- Registered: 2004-07-22
- Posts: 30,564
Re: Forty 40 Ton Mars Delivery Mechanism
Here is the article from GW
40 Tons Cargo 1-way to Mars Studies GWJ 2-21-2026
I have not been able to find very much of what I did toward this topic, more than a year ago. Maybe much more than a year ago. I have not found any of the numbers that I ran. But I did find one old “paint” sketch of where and how I started, and I remembered enough to draw a sketch of where I ended up.
I got started with a single rocket core, based on a ring-shaped cargo container that doubles as the rough-field landing gear. Initially, the rocket core was to be reusable. This thing was launched from LEO with a transfer stage. If you use the “right” transfer orbit to Mars, one
whose period is an integer multiple of 1 year, you can potentially recover and reuse those transfer stages.I used the spreadsheets and information in the “orbits+” course materials, the links to which are on the forums, to do the orbit transfer stuff, and the free entry at ~7.5 km/s directly into the Martian atmosphere. I also did the deceleration and thrust requirements
for the final descent and landing, very likely presuming Mach 3 speed at end of hypersonics altitude, angled 45 degrees down.
The biggest problems I ran into were (1) no estimate of plasma radiation heating built into the entry spreadsheet at that time (since rectified), and (2) there was not enough room to put enough engines in that core to get enough thrust for the propulsive landing. It comes out low, so the deceleration requirement was rather high. Near 6 gees as I recall. (But as I said, I can no longer find these calculations.)
The solution to this was less symmetry. Put two shipping containers side by side and put multiple sets of engines and tanks between them. That way I could get enough thrust in place, to make the high-deceleration descent and landing. With two long doors that open to let the engines thrust, the entire bottom is one big flat heat shield.
The convective stagnation heating would be deceptively low. I found out later that, quite unlike Earth entry, entry on Mars risks very high plasma radiation heating. That issue is still undressed in this design, although tipping the containers slightly and adding a small nose radius to the heat shield doors might be the solution.
As for the “40 tons” thing, you can still scale this idea up or down a fair amount, before something trips you up. It is a means to do significant 1-way cargo deliveries direct to the Martian surface from LEO.
I suppose the rocket engines might be recovered, along with any unused propellants in the tanks. I used storable NTO-MMH to completely avoid any and all evaporative loss issues. This propulsion need only make the final course corrections before entry, then the hard
deceleration to a hover at about 30 m for the final landing.
Content from Email follows
I can try to revisit the design if you like. I think it might also be a good "exrocketman" posting.
I'm thinking that somehow I need about a 10 degree slant from inboard to outboard of the floors of the two cargo containers, with a nose radius no larger than the separation between them. I'm also thinking I need to make each container in two sections, joined at an angle in the center, letting each slope at about 10 degrees from end to center. A nominal 10-deg "cone" angle will not hurt the rough-field stability or ground penetration issues very much at all. 20 degrees would be too much, though.
That gets me a flattish heat shield shape sloped at 10 degrees in both directions, which is a first cousin to a flat plate in terms of drag, but still very easily directs the shocked flow outward in both directions to the edges. I just need to make the "nose radius" of the heat shield doors, down the center between the containers, small enough to make peak plasma radiation stagnation heating about equal to peak convective stagnation heating.
The "trick" will be designing the heat shielding to the higher entry speed of a 2-year abort orbit for the transfer, while anticipating that most cargo can go min energy Hohmann. That way, faster transfer of critical supplies can proceed. Plus, we will need to send these vehicles that faster way anyway, if we are to reuse the transfer stages. The difference is 5.3 km/s entry interface speed Hohmann, vs 7.5 km/s entry interface speed for the 2-year abort orbit.
All of this works very well indeed, if you have a facility in LEO that (1) can assemble them from stuff sent up from Earth, (2) can fuel them up for their trip, and (3) can also service the transfer stages that will send them to Mars. Requirements for that facility are primarily a low-inclination orbit (quite unlike ISS), and a low-enough circular orbit to be reached easily without requiring constant re-boost. 300-600 km altitude is about right.
The transfer stages, if 1-shot, could actually be solids, with a small storable propulsion system to fine-tune burnout speeds and to do the initial midcourse correction. Most versions of Minuteman used a storable "bus" to put the warheads on exactly the right trajectory, right after the last big solid stage burned out.
If these transfer stages are to be reusable, they should use high-energy cryogens for departure, and storables for course corrections and homecoming-capture into a highly- elliptic capture orbit. You will need a tug to retrieve them from that elliptic capture orbit. A tug based at that same LEO facility.
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#65 Yesterday 19:26:48
- SpaceNut
- Administrator
- From: New Hampshire
- Registered: 2004-07-22
- Posts: 30,564
Re: Forty 40 Ton Mars Delivery Mechanism
Copilot response
single rocket core, based on a ring-shaped cargo container that doubles as the rough-field landing gear” and then evolved it into “two shipping containers side by side and…multiple sets of engines and tanks between them…With two long doors that open to let the engines thrust, the entire bottom is one big flat heat shield.”
So, structurally, the concept is:
In transit: Compact, attitude-friendly configuration, essentially a core plus containers.
Entry: Flat-bottom, dual-container heat shield, doors closed; free entry at ~ 7.5 km/s to Mars.
Post-hypersonic: Doors open, engines between the containers fire for a hard retropropulsive decel (you recall ~6 g), landing with the containers themselves acting as rough-field pads.
Propulsion: Storable NTO/MMH, only for course trim and terminal descent/hover, no boiloff issues.
The two big open items you called out are exactly where the real design work lives:
Entry aerothermals, especially radiative heating on that flat bottom.
Engine layout and thrust sizing to hit the decel/landing requirements without insane g-loads or unbuildable clustering.
If you’re up for it, I’d suggest we rebuild this in a clean, stepwise way:
Lock the geometry.
Assume container analogs: e.g., 2 × “Mars ISO” containers (pick L × W × H), plus gap for engines.
Define heat shield footprint: total flat area, corner radii, possible nose radius/tip-up angle you mentioned.
Set the mass budget for a “40 t to surface” case.
Payload: 40 t.
Structure + TPS + tanks + engines: first-pass fractions.
Total entry mass: what we’re actually decelerating and landing.
Re-derive the terminal descent requirement.
Assume: Mach ~3 at end of hypersonics, ~45° flight path angle, target hover at ~30 m AGL (as you wrote).
From that, we can back out required
Δ ? , peak decel, and total thrust, then translate to engine count and clustering feasibility.
Sketch an entry/heating envelope.
Use a conservative Mars entry speed and ballistic coefficient from the geometry and mass.
Then we can talk about what TPS family and thickness might be needed for both convective and radiative heating on a flat-ish shield, and whether a slight tip or nose radius is enough to tame the radiative load.
Decide on reusability vs. expendable.
You hinted that “the rocket engines might be recovered, along with any unused propellants in the tanks.” We can see if that’s mass-logical or if this wants to be a mostly expendable cargo brick.
erothermally self-consistent.
1. That 10° “double-slope” heat shield
What you’re describing is basically a very low-angle faceted blunt body:
Two containers, each broken into two segments, with a ~10° break at mid-length, and
A 10° slope from inboard→outboard, so the whole “floor” is tipped both spanwise and lengthwise.
That gives you:
Drag: Still “first cousin to a flat plate” in Cd, as you said—high drag, low ballistic coefficient if you keep areal loading modest.
Flowfield: Shock pushed off and directed outward in both directions, which is exactly what you want to keep the hottest gas sliding away from the centerline.
Rough-field behavior: A 10° cone/facet angle is small enough that your “pads” still behave like big, mostly-flat skids on contact. 20° really would start to feel like landing on two shallow wedges.
The key knob you called out is the centerline nose radius between the containers: small enough to keep radiative stagnation heating comparable to convective, but not so sharp that you get crazy local loads or structural/TPS headaches. That’s a tunable parameter we could actually sweep in a simple correlation set once you lock in:
Container width and spacing (sets max allowable nose radius).
Total entry mass and shield area (sets ballistic coefficient and heating level).
2. Designing to 7.5 km/s and letting 5.3 km/s be “easy mode”
Your instinct to design the TPS to the 2‑year abort orbit entry speed is dead on.
Convective heating scales roughly like
??∝? 3 . (7.5 5.3) 3 ≈ 2.8 , so convective peak can be almost 3× higher at 7.5 km/s than at 5.3 km/s.
Radiative heating is even steeper with velocity (often
?? with ?>3), so the penalty for the faster entry is harsher still.
So if you size the TPS for 7.5 km/s, the Hohmann 5.3 km/s entries become “low-stress” cases with big margin—exactly what you want for bulk cargo. The faster shots then become the critical-supplies / transfer-stage-reuse cases you explicitly care about.
3. LEO facility + transfer stages: architecture actually hangs together
This part of your writeup is quietly powerful:
“All of this works very well indeed, if you have a facility in LEO that (1) can assemble them from stuff sent up from Earth, (2) can fuel them up for their trip, and (3) can also service the transfer stages that will send them to Mars.”
And:
“The difference is 5.3 km/s entry interface speed Hohmann, vs 7.5 km/s entry interface speed for the 2-year abort orbit.”
You’re basically defining a Mars cargo pipeline:
Low‑inclination, 300–600 km LEO node for assembly, fueling, and transfer-stage turnaround.
One-shot solids with a small storable “bus” for cheap, expendable transfers (Minuteman-style).
Reusable cryogenic transfer stages for the high-value, high-speed missions, with storables for course correction and Earth recapture into a high-elliptic orbit, then a tug from the same LEO node to bring them home.
The lander you’re sketching is the standardized cargo brick that rides that pipeline.
4. Where we can push this next
If you want to turn this into a solid exrocketman post and a usable design reference, I’d suggest:
Freeze a notional geometry:
Container outer dims (L, W, H),
Centerline spacing,
Exact 10° breaks (length from end to center, height offset),
Max allowable nose radius between containers.
Pick a 40 t payload case and rough mass budget:
Payload, container structure, TPS, tanks, engines, residuals → total entry mass.
From that, compute ballistic coefficient and plug into standard Mars heating correlations for both 5.3 and 7.5 km/s.
Back-of-envelope engine sizing:
Assume end-of-hypersonics Mach ~3, 45° FPA, target hover at ~30 m.
Derive required
Δ ? , peak decel, and total thrust; then see what engine clustering fits between the containers without wrecking the shield geometry.
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