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I'm starting this topic to discuss the technology for minimal risk / minimal complexity crewed Mars flyby and exploration missions using a miniaturized version of the 500 passenger Interplanetary Transport Vehicle (ITV), transporting an exploration crew of up to 4 human crew per ship plus 4 robotic crew members to assist with maintaining the ship and potentially dangerous tasks such as EVAs.
The same basic design principles would apply, albeit to a much smaller and simplified ship design that uses an appropriate mix of electric and storable chemical propulsion systems to complete exploration missions. A colonization class (c-ITV) requires advanced Solar Thermal or Nuclear Thermal propulsion to contain propellant / launch costs. An exploration class (e-ITV) can use simpler near-term propulsion solutions that won't work at the scale required for colonization. The e-ITV program will feature many of the programmatic elements of a colonization campaign, but at a smaller scale.
To the extent feasible, vehicle fabrication methods, propulsion systems, computer control systems, and mission architecture will only feature mature or highly refined systems. This means the major mission elements and associated systems as a whole must constitute an engineering exercise, rather than a clean sheet developmental program. As an example, for a propulsion system to be used, it must have already flown in space at least once. Novel applications of existing tech are fine, but if you want to develop a brand new "clean sheet" engine design, that's a separate development program that won't become part of this program until it flies in space.
ITV Design Characteristics:
Counter-Rotation Artificial Gravity
Counter-rotation minimizes gyroscopic precession effects. A motor of some kind can spin the two habitation spaces in opposite directions, likely varying speeds just a little, using accelerometer sensors to provide feedback, to help minimize the destabilizing effect of small mass imbalances between both sets of rotors. The destabilizing effect of rotating only in one direction complicates the operation of propulsion systems or requires deployment of tethers to connect a habitation module to an upper stage mass. A rigid "baton" design could and probably would also work, but batons typically become heavier the longer they are, to resist deformation caused by the artificial gravity induced by spinning them. If the baton is much stronger than it needs to be, because it was originally designed as a rocket vehicle's upper stage, then it's also much heavier than it needs to be, so more propulsive power and propellant are required.
CFRP Primary Structures
We can use automated tape-winding machines to deposit very thin (70g/m^2) layers of Carbon Fiber pre-preg onto a mold using a laser to secure and partially cure the tape in place. Since Boeing's autoclaves are large enough for all the major hull parts to fit, a complete curing process can be utilized for a full density / full hardness / minimum porosity hull. Composite parts produced this way have a density on-par with Magnesium, at 1.7g/cm^3. A 1m^2 panel of CFRP material 10mm thick is therefore 17kg/m^2, or 42.5kg/m^2 at 25mm thick. Excluding end domes, a 4.5m OD 10m long module with 25mm thick walls weighs 11,883kg. The 4.3m OD 8.5m long Destiny module, which was primarily Al-2219-T87, had a mass of 14,515kg for comparison purposes.
We'll use the same IM7 fiber and Cycom 5320-1 resin system which has proven so resistant to O2 and H2 permeation in NASA's Composite Cryogenic Tank Demonstrator program. CFRP LH2 tanks were constructed using automated tape winding laying fiber onto collapsible molds with alternating 70g/m^2 and 140g/m^2 IM7 fiber layers, to speed up the tape winding process. Those tanks were not vacuum bagged and autoclaved, yet they held pressure and resisted H2 / LH2 permeation, at pressures up to 58psi, and were less than 10mm thick. That particular fiber and resin system combination is CTE-matched such that temperature extremes had little effect on durability.
For fire resistance, the hull's interior can be lined with Carbon-X fabric, affixed in place using CFRP rods woven into the fabric, so no adhesives or metallic fasteners are required. The exterior MMOD protection layers should be a combination of Kevlar and Nextel. A Vectran outermost layer will reflect sunlight and resist atomic Oxygen attack in LEO.
Extensive use of fabrics and composites, which contain a lot of Carbon and Hydrogen, will reduce the secondary radiation dose associated with Galactic Cosmic Rays (relativistic ionized particles) striking thin-walled metallic structures. Some protection from the intense proton storms produced by Solar Particle Events (SPEs) and Coronal Mass Ejections (CMEs) will be provided as well. Additional protection using food and water crew provisions will still be required for adequate protection from the most powerful solar storms.
Storable Chemical and Solar Electric Propulsion (SEP) Systems
After the initial Trans-Mars Injection (TMI) burn performed with cryogenic liquid propellants, Mars Orbit Capture, Mars Orbit Transfer, Low Mars Orbit station-keeping, Trans-Earth Injection, Earth Orbit Capture, Earth Orbit Transfer, and Low Earth Orbit insertion burns will be performed using an appropriate combination of storable chemicals and electric thrusters. The crew need not remain aboard the ITV during the spiral into LMO, from TMI, nor to LEO, on the way back to Earth from Mars. Direct entry with appropriate vehicles is acceptable. SEP using Argon ($35/kg) or Iodine ($61/kg) propellants can deliver 2,000s Isp. Both Argon and Iodine are dramatically less expensive and more plentiful than Xenon ($3,000/kg) and Krypton ($2,100kg to $4,800/kg). Iodine is very dense and can be stored as a liquid with modest heating applied. Storable chemicals like Hydrazine ($78.50/kg) or HAN ($200/kg to $350/kg) or Ammonia ($0.73/kg) can deliver up to 600s Isp using ArcJet thrusters. Thrust levels for all propellants mentioned are very modest, on the order of 10s of mN/kW of input electric power, but Isp ranges from 580s to 2,000s.
Conventional NTO/MMH or NTO/HAN can provide high thrust for impulsive maneuvers, with Isp ranging between 340s and 350s when used by pump-fed engines. Small electric pumps can generate extreme pressures for brief periods of time, such as the Rutherford engine used by the RocketLabs Electron small satellite launch vehicle. Better still, these pumps can suck the tanks dry without exploding, unlike conventional turbopumps, so there is no residual / unusable propellant left. In normal turbopump-fed chemical rocket engines, 2% to 3% of the propellant mass is unusable because attempting to use the pump to extract that last bit of propellant would cause pump cavitation, swiftly followed by rapid unscheduled disassembly.
Whenever Zero Boil-Off (ZBO) technology is truly ready for space flight applications, then LOX/RP1 with electric pumps provides the best Density Impulse and Total Impulse characteristics for a given propellant tank and engine hardware mass, at the expensive of propulsive efficiency. LOX/LCH4 or LOX/LH2 are highly desirable for in-space propulsion if sufficient tank insulation and cryocooler power is available, but the greater dry mass fraction of LCH4 or LH2 fueled stages vs RP1 stages for equivalent Delta-V (ΔV), up to about 4.5km/s per increment.
At 465.5s, using LH2 fueled RL-10B-2 engines, possibly 470s using greater nozzle expansion ratios, we might get a 15% to 20% payload performance advantage over RP1 fueled engines, for equal in-space propulsion stage mass. Unfortunately, LH2 ZBO tech is still in active development and nowhere near ready for crewed missions.
The last page of that document seems to indicate that we still have another 5 to 10 years of developmental work before ZBO systems for LH2 is ready to go. It's a complex bit of tech. Every listed piece of tech in the chart on the last page needs to be at TRL7 or higher, preferably TRL8, before we stake human lives on it. That takes time, money, and a coherent development effort with parallel development of stages designed to capitalize on it.
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I think we should pursue a modified minimum risk mission that uses existing air and water recycling technologies that we already paid through the nose to develop:
Minimum Risk Deep Space Habitat and Life Support
I think the technological risk associated with using regenerative life support is greatly over-played. ISS wouldn't have been habitable for 20+ years if the tech didn't work.
The next generation life support tech, specifically the Amine Swingbed CO2 Scrubber and Ionomer Waste Water Processor have been quietly doing an even better job aboard ISS, using a lot less power and suffering fewer system casualties, for over 2 years now. They're much closer to closed-loop systems, drastically lighter / more compact, and they've had very few teething problems. If the ionic liquid CO2 scrubber experimentation pans out, the total amount of power both systems consume will be under 1kWe for a crew of 4. The total power requirement should be 2.5kWe or less. That means photovoltaic array deployment mechanisms are not required. Waste heat management radiators can be fixed / hull-integrated designs as well.
600W - Counter-Rotation Motors (CRMs)
400W - Skylab Control Moment Gyros (for 8 CMGs at normal operating rpm)
Source:
[url=https://ntrs.nasa.gov/api/citations/19790007076/downloads/19790007076.pdf]NASA TM-78212 - 25 kW POWER MODULE UPDATED
BASELINE SYSTEM - December 1978 - Page Labeled 39 in lower right-hand corner[/url]
800W - Carbon Dioxide Removal by Ionic Liquid Sorbent (CDRILS) System
Source:
Scale-up of the Carbon Dioxide Removal by Ionic Liquid Sorbent (CDRILS) System
Novel Liquid Sorbent CO2 Removal System for Microgravity Applications
60W - Ionomer Waste Water Processor (Direct, Single Cycle, or Dual Cycle)
Development of Ionomer-membrane Water Processor (IWP) technology for water recovery from urine
Demonstration of a Full Scale Integrated Membrane Aerated Bioreactor- Ionomer-Membrane Water Purification System for Recycling Early Planetary Base Wastewater
140W - Avionics / flight computers, electronics, communications, sensors (redundant mobile chips using power-over-fiber sensors)
2kWe is a very modest amount of power for 4 crew members. That can be supplied by hull-integrated thin film arrays. Additional power can be stored in Lithium-ion batteries.
Last edited by kbd512 (Yesterday 01:44:12)
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The NASA science mission portion of what NASA does has been gutted. They will not even complete the missions already out there.
The NASA manned mission stuff is going to be greatly changed. Trump wants flags-and-footprints on the moon and on Mars, but nothing else. Once he gets that, no more NASA anything. Look to see SLS and Orion killed after the Srtemis-2 flight around the moon. And Gateway is toast, too.
That future loss of NASA contracts is really where the new baby rift between Musk and Trump is coming from. Musk's NASA contracts will dry up when that happens. DOD will not make up the difference, not with the likes of that incompetent Hegseth in charge of it. Musk knows that.
It's already just barely started over DOGE, with Musk essentially complaining that Trump's "big beautiful bill" does not kill enough Americans for lack of cuts to health care, social security, etc.
GW
Last edited by GW Johnson (2025-05-28 19:22:28)
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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Getting back to the proper topic here, what is needed is vehicles assembled in circular LEO at a space station facility that does both assembly work and propellant fill/refill work. You do it in circular LEO at fairly low inclination to make it reachable from the surface at the least launch dV.
There needs to be a space tug or multiple tugs based at that same facility. The tugs take the interplanetary craft to just below escape, so that a much smaller departure dV is demanded of the interplanetary craft, while the tug itself stays in an extended ellipse, and returns unladen to the station in LEO for reuse. That leaves more dV available for the interplanetary craft to enter low Mars orbit, should that be desired. Direct landings require less dV, but are not reusable. The direct landing dV is not as much less than the enter-LMO dV as most people think. Why? Aerobraking cannot do it all!
BTW, the same basic idea also works for lunar missions.
GW
Last edited by GW Johnson (2025-05-28 19:30:52)
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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GW,
Your complaints about Elon Musk, President Trump, and DOGE are duly noted, but this topic is about ITV technologies for exploration class Mars missions. There is a political topic for politics.
I think space tugs and propellant depots are a splendid idea, but again, this topic is about ITV technologies. I think we're already in agreement that all real exploration missions thus far have begun in LEO. At this point, I would be pleased with any crewed mission to another planet, including a fly-by mission, irrespective of what orbit it starts in or what propulsion tech it uses. A fly-by mission would at least demonstrate that our ITV is mission capable and that the crew is no worse for wear after the mission.
Propulsion is obviously a very important part of making all interplanetary missions possible, and I understand your affinity for those systems, which I share, but do you have any specific ideas on how we can create a more practical long-duration in-space vehicle to carry exploration crews to Mars?
I've already mentioned some of my ideas:
1. Counter-rotation sidesteps the precession and loss-of-control issues with artificial gravity.
2. Composites fabricated by robotic tape winding machines can create durable yet light vehicle hulls with excellent volume-to-mass ratios.
3. SEP, storable chemicals, and CMGs can provide adequate in-space propulsion and maneuvering after achieving escape velocity using already available cryogenic chemical propulsion upper stages.
Maybe we'll have ZBO tech perfected in another 5 to 10 years so that long duration on-orbit storage of cryogens becomes a more practical proposition. ZBO tech development began over 25 years ago. I think the required insulation tech is ready to apply, but not much else is. My understanding is that insulation alone, and perhaps some modest cryocooler hardware to re-liquefy small amounts of boil off as compared to LH2, is sufficient for LOX and LCH4. Unfortunately, NASA is dead set on using LH2. If the daily loss rates weren't so high I understand the potential benefits from using LH2. Thus far, long-term LH2 storage has proven to be a very tough nut to crack.
4. Rather than trying to store absolutely everything because we're afraid of equipment failures, capitalize on the excellent progress being made on low-temperature / low-power / high reliability CO2 scrubbers and water filtration systems. I see these technologies as "de-risking" the mission, rather than adding risk. Highly efficient air and water recycling is a mission enabler. We already spent the money to develop this tech to a very high degree of reliability and readiness, so we may as well use it. The mass reduction from water recycling alone is key. IWPs provide a very consistent 98% water recycling rate. CO2 recycling would provide another significant mass reduction.
There are near room temperature processes for CO2 splitting that leaves pure elemental Carbon dust floating on top of Gallium metal as the Gallium strips the O2 from CO2. An electrostatic attraction process could remove the Carbon dust for storage, although even mechanical separation methods appear to work. Heating the metal to 500C to 700C in a vacuum is sufficient to release 2 of the 3 Oxygen atoms. Hydrogen gas will strip the remaining Oxygen, producing water and Gallium metal. By the end of the mission we have 546kg of elemental Carbon from the respiration of 4 crew members. Human feces is 40% to 55% Carbon, so an additional 200kg to 275kg of Carbon could be recovered. This recovered Carbon is one of the important precursor materials for synthetic fibers, plastics, and rubbers.
5. Use the most modern avionics suite and sensors we have, because the reduction in power consumption and heat rejection are incredible. Power-over-fiber combined with optical chips is a radiation hardened by design computing environment. Some researchers here have already built a fully optical computer that performs all the necessary logic functions of legacy Silicon-based CPUs, to include RAM, but it does this with a 100GHz clock cycle. Some simpler designs are operating at up to 800GHz. That's 200 to 300 times faster than Silicon-based CPUs. Even a very modestly capable CPU with a clock cycle that fast can still perform the same functions as more sophisticated Silicon-based chips. The result is orders of magnitude power reduction.
Photonic crystal displays can render painting-on-paper quality colors. In the tech world, there's a recently established "fetish" for near-infinite color range and quality that looks as if the image was inked onto a piece of paper, even though it's been rendered onto a clear piece of glass.
Optical computers have been tested using specialized lasers and particle accelerators to withstand radiation doses in the range of 10s of megarads- a near-instant lethal dose for the crew. There's even been some talk of putting optical computers inside the cores of fission reactors, or at least ones operated at steam generating temperatures, as integrated monitoring and control systems. That makes them highly resistant to both damage and upset from normal radiation fluence ranges found in space.
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