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#126 Interplanetary transportation » Miniature ITV for Mars Flyby and Exploration Missions » 2025-05-25 17:55:14
- kbd512
- Replies: 22
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.
#127 Re: Not So Free Chat » Politics » 2025-05-24 21:54:19
Calliban and Terraformer,
What the heck is going on over there?
I'm worried about you guys.
#128 Re: Not So Free Chat » US Space Force to Deploy Lasers and Interceptor Missiles in Space » 2025-05-21 21:49:21
Calliban,
The Chinese and Russians have been building lots of new nuclear weapons, but I doubt many people have been paying attention. The Chinese have recently built far more nuclear weapons than they ever stockpiled historically, which started well before any missile defense system was announced. If that's what you're afraid of, Russia and China are already doing that, and they're not going to stop because someone there thinks a nuclear war is somehow winnable when you have enough warheads.
Given the sheer quantity of satellites already in space, if the Kessler Syndrome outcome was a probable outcome, then it should've happened already. There are over 28,000 satellites in space, 55,000 pieces of debris that we can track, and possibly as many as 170 million smaller pieces considered dangerous to spacecraft that are too small to track using current methods. We wouldn't necessarily have to wait decades because we could use lasers to vaporize little bits and pieces of satellites.
The US is moving towards near-satellites, high altitude drones, and alternative methods of providing battlefield communications and reconnaissance. That means there will be fewer opportunities to create large debris fields when the satellites are low cost commercial commodity hardware skimming through the upper atmosphere for a few months at a time, at most, before upgraded satellites take their place.
Doing nothing in response is one potential course of action, but not one likely to lead to desirable results for all us civilians who would lose our lives if Russia or China felt emboldened to launch their nuclear weapons at us. Is President Trump's plan "the best plan"? Probably not, but thus far it's the only solution on offer from our military industrial apparatus. This plan most certainly did not originate with President Trump, either. Presidents from both political parties, most notably Presidents Reagan, Bush Sr, Bush Jr, Obama, Biden, and now Trump have all spent large sums of money on missile defense capabilities. That means there never was any other realistic plan. President Clinton operated in a world where Russia and China were functionally crippled by Russia and China, so he was able to capitalize on the "peace dividend" at the end of the Cold War. Sadly, we no longer live in that world.
#129 Re: Not So Free Chat » US Space Force to Deploy Lasers and Interceptor Missiles in Space » 2025-05-21 09:05:24
I quite liked this one:
#130 Not So Free Chat » US Space Force to Deploy Lasers and Interceptor Missiles in Space » 2025-05-20 16:59:39
- kbd512
- Replies: 4
As the topic suggests, as of today President Trump has allocated $25B in defense spending to create a "Golden Dome" Homeland Missile Defense System against incoming cruise missiles, hypersonic missiles, ballistic missiles, and drones. The bill is anticipated to spend approximately $175B on ground and space-based missile defense systems, with initial rollout and construction taking place over the next three years.
President Trump's response follows on the heels of the apparent weaponization of space-based assets by Russia and China. Russia, specifically, has undertaken development of anti-satellite weapons carrying nuclear warheads. Both Russia and China have already deployed electronic warfare satellites intended to disrupt communications satellites and to blind reconnaissance satellites. Since 2015, China has deployed some 490 military satellites that would ostensibly be used in conjunction with their long range hypersonic and ballistic weapons to target their adversaries.
This article linked below from 2024 details US diplomatic efforts to attempt to deescalate the weaponization of space, which Russia vetoed and ignored:
Is There a Path to Counter Russia’s Space Weapons?
There are a number of other articles from 2024 outlining various changes to Russian and Chinese use of space-based military assets.
#131 Re: Human missions » Mars Ascent Vehicle - NTO/MMH/HAN vs LOX/LCH4/LH2 » 2025-05-20 12:44:22
An analysis of storable chemical propulsion stages for lunar and Mars missions from DLR:
High-Thrust in-Space Liquid Propulsion Stage: Storable Propellants
It makes use of the ATV and Aestus-II (Aerojet-Rocketdyne RS-72, which is based upon the XLR-132) since it's of European origin.
Edit:
Abstract
In the frame of a project funded by ESA, a consortium led by Avio in cooperation with Snecma, Cira, and DLR is performing the preliminary design of a High-Thrust in-Space Liquid Propulsion Stage for two different types of manned missions beyond Earth orbit. For these missions, one or two 100 ton stages are to be used to propel a manned vehicle. Three different propellant combinations; LOx/LH2, LOx/CH4 and MON-3/MMH are being compared.
The preliminary design of the storable variant (MON-3/MMH) has been performed by DLR. The Aestus II engine with a large nozzle expansion ratio has been chosen as baseline. A first iteration has demonstrated, that it indeed provides the best performance for the storable propellant combination, when considering all engines available today or which may be available in a short- to medium term. The RD-861 K engine has been proposed as alternative to reduce the development duration of the high-thrust stage. Structure analyses and optimisations have converged towards a common bulkhead architecture with a Whipple shield, similar to the one used on the ATV, to protect the main propellant tanks against perforations caused by meteoroids and space debris. The propulsion system has been built around six Aestus II engines equipped with TVC and placed on a circular engine thrust frame. The RCS, the thermal system, and the power system have also been included in the preliminary design, and they have been sized for the most demanding mission. The performance of the high-thrust stage, resulting from the preliminary design, has been assessed for both missions taken into consideration.
#132 Re: Meta New Mars » kbd512 Postings » 2025-05-19 12:48:58
I found these gravity and drag losses posted here:
Drag: Loss in Ascent, Gain in Descent, and What It Means for Scalability
Ariane A-44L: Gravity Loss: 1576 m/s Drag Loss: 135 m/s
Atlas I: Gravity Loss: 1395 m/s Drag Loss: 110 m/s
Delta 7925: Gravity Loss: 1150 m/s Drag Loss: 136 m/s
Shuttle: Gravity Loss: 1222 m/s Drag Loss: 107 m/s
Saturn V: Gravity Loss: 1534 m/s Drag Loss: 40 m/s (!!)
Titan IV/Centaur: Gravity Loss: 1442 m/s Drag Loss: 156 m/s
As you can see, gravity losses tend to dominate.
The Saturn V has a 10m diameter core, but it also has a very different length-to-diameter ratio.
This blurb from that "GravityLoss.com" blog post seems to lend some credibility to my assumption that a "stubbier" rocket would require less descent propellant:
Meaning that the terminal velocity is inversely proportional to the square root of the length of the rocket l and the density of the almost empty rocket, rho_ter. This means that if density is held constant (a reasonable first assumption), as well as shape (similar C_d), the terminal velocity is doubled when the rocket size (length) is quadrupled. And this means more delta vee needed for landing. It doesn’t matter if the rocket is lengthened and widened or only lengthened. If the constant density is held, the increased length (or depth) just means more mass per frontal area.
Edit:
The website seems to be run by a fellow space enthusiast named Valtteri Maja. The data / analysis comes from Jon Goff (co-founder of Masten Space Systems, President and CEO of Altius Space Machines), from a blog post over on Selenian Boondocks, which is another space / Thorium power enthusiast website run by former NASA employee Kirk Sorensen. That means it has some credibility behind it.
#133 Re: Meta New Mars » kbd512 Postings » 2025-05-19 08:33:17
tahanson43206,
My proposal WILL produce more aerodynamic drag, especially in the lower atmosphere. There can be no doubt about that. The question I have is, if we deleted over 15,000kg in mass from the stage by using a mass and volume optimized stage, and improve the engines to use a staged combustion vs gas generator cycle, do we still gain additional payload performance by making the Falcon's booster stage drastically lighter?
Aero drag matters a lot to airliners that continuously operate at high subsonic speeds lower in the atmosphere.
Does it matter nearly so much to rockets / rocket boosters that are only briefly within the lower atmosphere?
For example, the real world Falcon 9 booster attains an altitude of around 60km / 196,850ft at burnout, and it only burns for 162 seconds. There's almost no atmosphere up there, and average ascent rate is 72,907ft per minute. The vehicle is traveling slower in the lower atmosphere due to TWR, then much faster as its altitude increases and propellant supply is depleted.
When the booster comes back through the lower atmosphere, my contention is that more aero drag is better, because less propellant is required to slow the booster for landing.
Do we end up gaining back some of what we lost to drag during the ascent, in terms of reduced propellant consumption for the landing burns?
Maybe we do or maybe we don't, but a 15,000kg dry mass reduction using optimized stage geometry is a very significant amount of mass to remove from the Falcon 9 rocket's dry mass. The landing gear mass would also be significantly reduced because the landing gear track is so much wider, relative to the height of the booster. There's very little risk of toppling over. What might otherwise be a failure at a landing attempt with a 3.66m diameter / 42.6m height booster core would be perfectly acceptable with a 10m diameter / 11m height booster core. The required stroke length of the landing gear would be reduced with that much mass removed. The grid fins should become proportionally more effective at maneuvering the core as well. The cost of spinning fiber around a mold is much less than the cost of machining Al-2195 gores, precisely bending them, and welding them. Recall how long that process takes ULA when Tory Bruno gave a guided tour of ULA's rocket factory, because ULA is doing the exact same thing for the Atlas V core stage.
Composite cryotank development was intended to reduce tank mass, fabrication cost, and fabrication timeline over Al-2195, and all program objectives were met. I would think that the "less extreme" LOX and RP1 propellants, would permit further mass reductions since it doesn't have to survive LH2 internal pressurization and Hydrogen permeation.
Peak heating of the Falcon 9 booster's sidewalls is up to 20kW/m^2 (12.9W/in^2), and peak temperature is 250C (482F), caused not by aero heating but impingement of the rocket exhaust during the retro-propulsive burn. That said, the composites surrounding the F-35's engine are continuously subjected to 1,000F exhaust bypass from its F-135 engine, with peaks to 1,100F. Since those things don't fail, we should not expect a composite propellant tank to fail, either, when appropriate resins are used. However, the F-35's high temp skin panels also have their high temp resin cured in a high temp autoclave, which removes porosity from the composite and improves strength and weight. This would significantly add to the cost of the booster, as compared to NASA's non-autoclave process, but it might be worth the cost because of how much durability and temp resistance is increased. SpaceX uses Falcon 9 / Falcon Heavy as a high op-tempo workhorse rocket, much like the US military uses its F-35s. If the airframe / propellant tanks are now good for a thousand flights, then you may only need a handful of them to continue providing launch services to the US government.
The bottom line is that there are tradeoffs here. I'm very curious to know if that whopping stage mass reduction is "worth it" in terms of payload performance and other secondary effects such as landing propellant reduction, reduction of landing gear mass, landing stability on the deck of a heaving drone ship, etc.
#134 Re: Single Stage To Orbit » SSTO Engine Technology » 2025-05-18 17:40:30
Kbd512:
I knew there was an extendible bell at least tested on some engine. From what you said, it was a variant of the RL-10. Did that ever fly on anything?
GW
This Day in Aviation - 5 December 2014, 12:05 UTC
From the article:
The Delta IV Heavy’s second stage is 42.8 feet (13.05 meters) long, and is also 16 feet, 10.0 inches in diameter. It uses an Aerojet Rocketdyne RL-10B-2 engine, producing 24,750 pounds of thrust (110.09 kilonewtons) of thrust. The RL-10B-2 is 13.6 feet (4.15 meters) long, 7.0 feet (2.13 meters) in diameter, and weighs 611 pounds (277 kilograms).
It looks like it flew to me, 11 years ago now:
It's going to fly again, or actually, "again again" (since it's already flown once on SLS as well), with the SLS ICPS stage for Artemis II.
Edit:
RL-10B-2
According to this sales brochure from Pratt & Whitney, the RL-10B-2 engine first flew in space 1999 aboard a Delta-III. That was about 26 years ago, and it was used by the Delta-IV Heavy as an upper stage engine. I'm going to assume this is very well established technology at this point.
#135 Re: Fully Reusable Two Stage to Orbit (FR-TSTO) » FR-TSTO policy » 2025-05-18 17:31:44
GW,
Such designs will have loaded stage inert weight fractions in the 15-20+% range, or more (it's 40-50% in jet aircraft, and they still break up if they go broadside at full speed).
The Falcon booster weighs 25,600kg empty and contains 395,700kg of propellants. The 9 Merlin engines weigh 4,230kg, so 21,370kg is mostly Al-2195 alloy propellant tank structure. It's a fully reusable vehicle, as proven hundreds of times now. It's dry mass fraction is 6.08%. Suppose we make its structure 30% lighter using composites, so 14,959kg (structures) + 4,230kg (engines), 19,189kg in total, for a 4.63% dry mass fraction. If it's engines had a 200:1 TWR like Raptor 3, then it's engines weigh 3,878kg, so 18,837kg total dry mass, and 4.54% of wet mass.
The Falcon 9 is not mass optimized, so much as it's truck transportability optimized.
Falcon 9's total LOX load is 312,200kg (273.62m^3) and total RP1 load is 186,006kg (229.637kg), or 503.257m^3 in total. That 10m diameter composite cryotank holds 634m^3 of propellant and only weighs 3,037kg, and would only be 80% full. A tank structure for this mass optimized Falcon 9 booster would therefore weigh 7,267kg for propellant tanks (3,037kg) and engines (4,230kg). If we have 200:1 staged combustion engines, then we get an Isp bump and mass reduction. That cube-square law is really helping us out here. Aero drag at low altitude will be much higher, obviously, but we don't spend much time in the lower atmosphere. Our dry mass fraction is only 1.8%, though, or 1.7% using improved engines. Landing gear mass will obviously drive the inert mass fraction up, but now we have a much more stable vehicle when it comes time to land. There's almost no chance of tipping over.
All that said, we've cut our dry mass fraction by more than half by making those changes.
#136 Human missions » Mars Ascent Vehicle - NTO/MMH/HAN vs LOX/LCH4/LH2 » 2025-05-18 02:14:20
- kbd512
- Replies: 3
I'm starting this topic for a Mars Ascent Vehicle (MAV) using NTO (N2O4) oxidizer and MMH (Mono-Methyl Hydrazine) fuel vs NTO oxidizer and HAN (Hydroxyl Ammonium Nitrate) storable liquid propellants as alternatives to cryogenic liquids such as LOX/LCH4, LOX/LH2, and LOX/LCO. I think it's highly probable that the first exploration missions to Mars use legacy storable liquid propellants because the mission mass increase and technical challenges associated with ISPP to produce cryogenic liquids from scratch and then store them for many months will likely require some experimentation and refinement on Mars to ensure that everything works properly. To date, no serious funding has been allocated to ISPP. If ISPP tech is placed on "the critical path" to future Mars missions, then ISPP will be treated as another excuse as to why we still haven't sent humans to Mars.
Arguments For Storable Propellants
1. Respectable Isp in both pressure-fed (up to 330s) pump-fed (340s) engines
2. Ability to be stored almost indefinitely in chemically compatible tanks
3. Very high propellant bulk density, greater than liquid water, so propellant tank volume for Total Impulse is much smaller than cryogenic fuels
HAN Fuel Notes:
HAN fuel, when used as a substitute for MMH, eliminates most of the undesirable qualities of MMH.
1. HAN is mildly acidic and will corrode certain metals over extended storage duration, especially ordinary Carbon steel, but otherwise benign enough that humans can and do handle the fuel without extreme precautions taken to avoid skin, eye, or respiratory tract contact. A pair of goggles, Nitrile gloves, and ordinary lab clothing is considered to be sufficient protection. If any HAN fuel is spilled, cleanup is similar to diesel fuel cleanup. Handling of Hydrazine-based fuels requires a fully sealed hazmat suit with a positive pressure respirator.
2. HAN fuel provides a modest 7s Isp improvement over MMH when combusted with NTO. However, HAN's Density Impulse is a 50% improvement over MMH. HAN fuel tank volume for equivalent Total Impulse is half that of MMH.
3. HAN fuel can be frozen solid and then reheated in the tank prior to use without any risk of decomposition and thermal runaway, unlike MMH. This means no electrical power needs to be provided to keep the fuel warm during extended storage periods.
4. HAN fuel does not require double check valves typically used to prevent any accidental release of Hydrazine-based fuels, so fuel handling hardware is lighter and simpler. Research and testing, including testing in space by NASA and the US Air Force, has shown that HAN is broadly compatible with existing hardware commonly used to store and inject Hydrazine-based fuels, with the noted exceptions about metals that corrode easily.
5. It's conceivable that a staged combustion engine burning NTO/HAN, operating in a vacuum or near-vacuum and using a suitable nozzle expansion ratio, could achieve 360s of Isp. That would place overall engine performance in the same range as LOX/RP1, but with significantly greater Density Impulse. A lot of hay is made over the fantastic Isp of LCH4 and LH2, but not enough consideration is given to how greater propellant volume affects dry mass and cost. Using more materials to enclose a greater volume of propellant, even when that propellant is lower mass than more readily storable lower-Isp propellants for a given payload performance, especially for fully reusable spacecraft that must perform reentries, greatly affects monetary cost and developmental timeline. If a vehicle using cryogenic liquids must literally be twice as large as one using storable liquids, then the larger vehicle is likely to cost more and have a higher dry mass fraction as a result. There are no LH2 powered rockets that cost less to develop and operate than comparable RP1 powered rockets for that reason. Unfortunately, materials don't get any stronger as overall size increases. The cube-square law partially compensates for this, but any vehicle twice as large as an alternative design is almost certain to have greater dry mass or require stronger but more expensive lighter materials to compensate. Mindless fixation on Isp, as important as Isp is to rocketry, is driving up the monetary cost and developmental timelines for both rockets and landers.
Example Pump-Fed NTO/MMH Engine
Engine Designation: Aestus II / RS-72 (German-American collaboration)
Application: Ariane V orbital transfer stages
Propellants: NTO/MMH
Mixture Ratio: 1.9:1
Cycle: Gas Generator, Pump-Fed (XLR-132 pumps)
Thrust: 5,647kg-f
Isp: 340s
Chamber Pressure: 60bar
Nozzle Expansion Ratio: 84:1
Max Burn Time: 600s
Length: 2.29m
Engine Nozzle Exit Diameter: 1.31m
Weight: 138kg
Human Landing System Architecture Options Utilizing the XLR-132 Rocket Engine
NASA’s Artemis lunar exploration program aims to return humans to the moon over 50 years after the Apollo 17 crew lifted off from the surface. This time around, the goal is not only to return to the lunar surface, but to stay and explore, learning and preparing for later human exploration missions to Mars. In late 2019, Aerojet Rocketdyne (AR) performed an architecture trade study to evaluate several vehicle configurations and propulsion options that would satisfy the current mission requirements set forth by NASA. This study was completed under a NextSTEP-2 Appendix E: Human Landing System (HLS) Studies, Risk Reduction, Development, and Demonstration contract with NASA. The overarching observations from this study are still relevant as NASA and its HLS contractors are in the midst of designing the vehicles that will ultimately land on the lunar surface. The architecture trade study included an examination of a range of HLS configurations, launch vehicle options, concept of operations (CONOPS) options, main propulsion options, and other subsystem design options. Architectures were evaluated based on cost, schedule, reliability, extensibility, and performance attributes. Architecture configurations were scored and ranked through a utility analysis methodology. High scoring configurations that spanned the design space were further studied with alternate attribute weightings and Monte Carlo uncertainty analyses. Based on the assumptions made during the definition phase of the study, architectures using pump-fed storable propulsion received the highest overall scores across each set of weightings and uncertainty analysis. Architecture configurations that utilized the XLR-132 pump-fed storable engine scored particularly well indicating the XLR-132 to be an attractive main propulsion candidate for HLS elements. This paper provides a summary of the HLS architecture study performed by AR, and emphasizes 3-element and 2-element architecture configurations that take advantage of the XLR-132 pump-fed storable engine. Background details of the XLR-132 engine are also provided.
Arguments Against Storable Propellants
1. Cryogenic LOX/LCH4 (370-380s) and LOX/LH2 (450-470s) offers significantly higher Isp
2. Advocates of ISRU / ISPP would like to make some or all of the propellants from Martian-sourced CO2 and H2O
3. NTO and MMH are highly toxic to humans and require specialized and expensive propellant loading facilities here on Earth
LCH4/LH2 Notes:
Provided that NASA's new Zero-Boil Off (ZBO) cryogenic storage technologies are flight proven in space, cryogenic oxidizers and fuels provide clear performance advantages, which is why virtually all orbital launch vehicles use LOX and RP1, LCH4, and/or LH2. In addition to a ZBO demonstrator mission with LOX and LCH4 or LH2, further demonstrator missions must be performed to evaluate ISRU / ISPP mission hardware elements for durability and reliability in their intended operating environments. Lives cannot be staked upon the proper function of this new tech until exhaustive testing has been performed. Long term bulk storage of cryogenic liquids is a challenging mission requirement, even though there is ongoing refinement of ZBO and ISPP tech in the works. ZBO development is much further along than ISPP at the present time. ZBO and ISPP both radically increase the mission power requirements, so appropriate nuclear and solar power sources must be added to the mission hardware set. This tends to increase the mass of the mission elements, and it's undeniable that making rocket propellants from scratch, using indigenous natural materials, confers greater complexity to any mission relying upon ISPP. However, the propellant mass reduction leverage that ISPP brings to the mission is a compelling reason to pursue it anyway.
1. As fuels, the primary desirable characteristics of LCH4 and LH2 are no coking of engine internals so engine restart and reusability are greatly improved, autogenous pressurization makes it possible to remove heavy and potentially dangerous COPVs, and these fuels provide the highest Isp values amongst all practical chemical fuels.
2. Liquid cryogens are fantastic for removing waste heat from engine components prior to being fed into the combustion chamber. This property makes it possible to use high thermal conductivity metal alloys that are easy to work with, at least for modern automated 3D printing machines. Since not melting engine components is a high priority if you want to use them again, this feature is particularly appealing.
3. There's an enormous body of research, development, and implementation of LCH4 and LH2. We paid dearly for the performance benefits that cryogenic liquid fuels deliver, but all that effort ultimately made interplanetary missions feasible to begin with. The upper stages of the Saturn V used LH2 because that was the only way to deliver the demanded payload performance. If the LH2 fueled J-2 engines were not available, a significantly more powerful rocket would be required. The Soviet N-1 rocket was exclusively powered by RP1, required 5 stages, and could only throw 33t to TLI. In contrast, Saturn V only required 3 stages because both upper stages were LH2 fueled, so it could throw over 52t to TLI as a result. Vacuum LH2 fueled engines have Isp values of up to 470s. On a per mass basis, no other fuel can compete with LH2. Since Hydrogen is also the lightest element on the periodic table, LH2 is also the least dense fuel available by a very wide margin. This tends to make storing large amounts of Hydrogen difficult. However, LCH4 still has a high Isp of up to 380s and is far easier to store in a much smaller tank volume.
4. Since LCH4 and LH2 burn so cleanly, inspecting the engines for signs of damage after a landing burn, prior to a later liftoff and ascent to orbit, is a much simpler proposition. Since virtually all ISPP schemes make a little extra propellant, ground vehicles for surface exploration can also be powered by these propellants. Specialized in-space piston spark-ignited engines have already been developed that run exclusively on injected O2 and H2 or CH4, for example. One such engine has been developed as an APU for a LH2 fueled orbital transfer stage to reduce power generation and propellant pressurization mass over batteries and COPVs for longer duration LH2 storage between the time the rocket ascends to orbit and the time that the upper stage restarts to inject the payload into a higher orbit or escape trajectory. There are a variety of turbine-based APUs that very briefly run on Hydrazine fuels, but no such engine has been used to power ground vehicles as far as I'm aware, because long term engine durability is problematic. Every different kind of engine that exists has operated reliably on H2 or CH4, often for many years. Modern life as we know it would not be possible without fuels containing Hydrogen and Carbon.
5. There are no toxicity concerns with CH4 and H2. Even fairly extreme exposures to these fuels are not known to cause any lasting damage to the human body. Short of inhaling enough gaseous Methane or Hydrogen to be asphyxiated, it's not going to kill you. Virtually all storable oxidizers and fuels are capable of producing at least some long term damage from either massive exposure or chronic exposure.
Cryogenic liquids are staple propellants used by space faring nations. Virtually all high utilization rate orbital launch vehicles use LOX, for example, because it's pretty close to the best oxidizer available, very dense, very cheap to produce, and easy to store as cryogens go. LOX usually constitutes the majority of the propellant mass. While Chlorine and Fluorine are even better oxidizers, they're also incredibly toxic and not generally used outside of solid propellants that bind them to another chemical.
#137 Re: Single Stage To Orbit » SSTO Engine Technology » 2025-05-17 15:25:51
GW,
Is there any reason why you cannot have multiple turbopumps feeding propellants into the same combustion chamber (so that you can use a much larger nozzle without singular gigantic turbopumps)?
#138 Re: Single Stage To Orbit » SSTO Engine Technology » 2025-05-17 15:22:26
tahanson43206,
The RL-10B-2 engine uses a RCC extendible nozzle. This nozzle extension is about the same size and weight as what a Vacuum-optimized Merlin requires. It appears as though the real payload benefit to a hypothetical Space Shuttle mass SSTO vehicle so-equipped, is no more than about 2,300kg, based upon the nozzle hardware mass required to equip 40 Merlin engines. That's more than what I thought it would be, and most of the mass appears to be the nozzle extension itself, rather than the deployment mechanism, which has almost negligible mass per engine. The extension hardware is only 10 to 20lbs per engine.
Apart from the cost of the nozzle extension, the nozzle mass increase pretty much kills this idea. You would get more useful payload by cutting the weight of each engine by using the same RCC material for all the major engine components, combined with a staged combustion cycle. The Merlin is a marvel of gas generator engine tech, but staged combustion always provides higher Isp.
That 116,120kg mass to orbit value is inflexible because 6,508,946,390N-s delivers 116,120kg to orbit. We know this because that was the Total Impulse provided by 3X RS-25 engines affixed to the Space Shuttle and 2X SRBs.
When I use Silverbird Astronautics Launch Vehicle Performance Calculator, this is what I get:
Inputs
Launch Vehicle: User Defined
Number of Stages: 1
Strap-on Boosters?: No
Dry Mass: 26,372kg
Propellant Mass: 2,185,289kg
Thrust: 33,854kN
Isp: 304.2s (90% of Vacuum Isp for the RD-180)
Default Propellant Residuals?: Yes
Restartable Upper Stage?: No
Payload Fairing Mass: 0kg
Launch Site: Cape Canaveral (USA)
Destination: Earth Orbit, Apogee 185km, Perigee 185km, Inclination: 45 degrees
Outputs
Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit: 185 x 185 km, 45 deg
Estimated Payload: 78,854kg
95% Confidence Interval: 58,892kg - 103,231kg
When the dry mass is adjusted downward to only 13,426kg (4,315kg RCC engines + same 9,111kg propellant tank mass)
Estimated Payload: 91,813kg
95% Confidence Interval: 71,844kg - 116,164kg
ISS orbit, same RCC engines:
Destination Orbit: 400 x 400km, 45deg
Estimated Payload: 86,900kg
95% Confidence Interval: 67,807kg - 110,215kg
ISS orbit, same RCC engines, 323.1s Isp (90% of RD-0124 Isp):
Estimated Payload: 108,038kg
95% Confidence Interval: 86,343kg - 134,287kg
This appears to be little better or worse than the real Space Shuttle if we added the hardware for resuability. The only measurable performance improvement achieved during the STS program came from making the External Tank lighter. These payload performance estimates confirm that a modestly better Isp from the same engines and propellant mass confers a meaningful payload performance advantage, but only when greater dry mass doesn't immediately replaces that useful payload mass.
Historical Space Shuttle GLOW and Propellant Mass: 2,032,096kg; 735,602kg LOX/LH2 plus 997,904kg APCP, 1,733,506kg total
SSTO Space Shuttle GLOW and Propellant Mass: 2,301,409kg; 2,185,289kg LOX/RP1
The historical Space Shuttle has a propellant mass reduction of 451,783kg, but both LH2 and APCP are much more expensive than RP1.
The SSTO Space Shuttle carries 587,443kg of RP1. The real Space Shuttle carried 106,261kg of LH2 and 159,665kg of solid fuel within the APCP oxidizer / fuel combo mixed into the propellant grain of the solid rocket boosters, or 265,926kg in total. The remainder of the solid propellant mass was AP oxidizer.
LOX: $0.27/kg
RP1: $2.30/kg
APCP: $5.00/kg
LH2: $6.10/kg
LCH4: $8.80/kg
Hydrazine: $75.80/kg
SSTO Space Shuttle Fuel Cost
LOX: 1,597,846kg * $0.27/kg = $431,418
RP1: 587,443kg * $2.30/kg = $1,351,119
Total Propellant Cost: $1,782,537
Historical Space Shuttle Fuel Cost
LOX: 629,341kg * $0.27/kg = $169,922
LH2: 106,261kg * $6.10/kg = $648,192
APCP: 1,733,506kg * $5.00/kg = $8,667,530
Total Propellant Cost: $9,485,644
That makes the propellant costs 5.3X cheaper for the SSTO Space Shuttle vs the real Space Shuttle, despite the fact that the SSTO Space Shuttle is burning 452t of additional propellant. Most of what the SSTO variant is burning is LOX. RP1 exhaust doesn't produce HCl, either, unlike APCP. We lacked the materials and engine tech necessary for any kind of SSTO when the real Space Shuttle was designed, so that's a moot point.
If whatever changes you're making to the vehicle lead to a greater dry mass, then you need more engine power and more propellant, period.
#139 Re: Human missions » Mars Ascent Vehicle - LOX/LCO vs LOX/LCH4 » 2025-05-17 12:13:51
RobertDyck,
We need to control cost. Using the same damn excuses that got programs cancelled will only result in getting your program cancelled.
We're already doing that to a far greater degree than ever before.
Back when we were paying $54,500/kg, launch costs were a major factor in any space program's success or failure. Starship's launch cost is projected to be $10/kg. Let's say they're off-the-mark on launch costs by a factor of 10. For every 1kg we could launch at Space Shuttle prices, we can launch 545kg at Starship prices. A 2,000,000kg mission package, at $100/kg, is only $200M. Spread over two years, NASA's total investment per year is equivalent to a single Falcon Heavy Launch. If NASA won't commit to that level of investment in launch services for exploration, then our manned space exploration program is a farce that we should dismantle so that something functional can take its place.
If Starship truly can lower launch costs to $10/kg, then I don't know why we're arguing over $20M per Mars mission. There are at least 2 million Americans would donate $10 in gas money to NASA if it meant we could sit down at dinner to watch a televised Mars landing with a crew of astronauts. The entire world was glued to their TV sets when we did the lunar landings. That event was so fundamentally different from anything else in the human experience that they could not look away.
As I said before, and will continue to repeat, ISRU / ISPP will not make or break a Mars exploration mission. In any case, it's not ready to implement because no serious money or effort has been devoted to using it. ISRU / ISPP has become a toll booth, sort of like the Lunar Gateway.
We have life support, as well as air and water recycling equipment, that's good enough to do the mission.
We have storable chemical propellants and electric propulsion for return to Earth, good enough to do the mission.
We have heavy lift launch vehicles, such as Falcon Heavy, that are powerful enough to do the mission.
We have capsules and habitat modules that are large but light enough to do the mission.
Let's argue over what we could have done better or differently... After we do the mission.
#140 Re: Single Stage To Orbit » SSTO Engine Technology » 2025-05-17 02:21:25
Dr Clark,
I decided to evaluate what an extendible / vacuum nozzle might net in terms of improved payload performance:
The Russian RD-0124 engine, a non-developmental LOX/RP1 engine in active service, has a vacuum Isp of 359s.
https://en.wikipedia.org/wiki/RD-0124
323.1s is 90% of that 359s Vacuum Isp
Mass Flow Rate (mdot) = Thrust / (Isp * g0)
3,452,113.5kg-f = 33,853,669N
mdot = 33,853,669N / (323.1 * 9.80665)
mdot = 33,853,669N / 3,168.528615
mdot = 10,684.35kg/s
6,508,946,390N-s / 33,853,669N = 192.267s
192.267s * 10,684.35kg/s = 2,054,248kg
Propellant Mass savings is 131,041kg, which nets an additional 7,407kg of payload to orbit. That amount of payload performance improvement would more than cover the mass allocation for the extendible nozzles. We need all the payload performance we can get for SSTOs, so I'll take it.
#141 Re: Human missions » Mars Ascent Vehicle - LOX/LCO vs LOX/LCH4 » 2025-05-16 23:17:11
RobertDyck,
I get why change is hard, but after as much time as has already passed us by, I think it's time to update plans to sync-up with actual hardware. A lot of tech development has taken place since 1989. The Falcon Heavy / Dragon update was interesting and has some potential, but right at the edge of technological feasibility. The mass margins are a bit too tight.
If SpaceX can successfully land robotic cargo-bearing Starships on Mars, then we're much further along than we've ever been in the past, even when the Saturn V and NERVA were on-offer. Is the SpaceX plan any better than Mars Direct? Perhaps it is to Elon Musk, but not to my way of thinking. Any vehicle sent to the surface of Mars, beyond a small reusable MAV to return people to LMO, needs to stay there. The energy cost of returning it to Earth is far too high.
If Starship ever straightens up and flies right, then we can do 8-person crews fairly easily. I wouldn't fret too much over the launch mass if we can send 200t to LEO with full reusability. If NASA will only pay for 10 launches per mission, then each mission has 2,000t of mass to work with. We can do a pair of 20t 4-seat MAVs, a pair of 50t surface rovers for real surface exploration, and a 100t ITV.
Using a combination of minimalist MAVs, chemical, and electric propulsion, we're not straining our propulsion capabilities at all. ISPP is a worthwhile program objective, but pursuing it should be a second priority to the primary objective of surface exploration. We can't do that if all the money, intellectual effort, and irreplaceable time is spent pursuing these side quests. We already have storable chemical propellants and suitable engines that only need to be integrated into a minimalist ascent vehicle.
If we're only interested in getting the job done, then we're going to use the simplest and most highly developed means of accomplishing major mission tasks, such as return to orbit, and forego better options, such as ISRU / ISPP, as program features that are worked into the exploration campaign as the technology matures to the point that it can be incorporated into future missions. That's how the ISS technologies were developed. We should go to Mars first using the propulsion hardware we already have and enhance our capabilities with ISRU / ISPP over time.
The first lunar missions did not have a rover because the tech wasn't ready. If we had to wait additional years because we mandated lunar rovers for each mission, then President Kennedy's timeline for lunar exploration would not have been possible.
#142 Re: Human missions » Mars Ascent Vehicle - LOX/LCO vs LOX/LCH4 » 2025-05-16 21:06:30
RobertDyck,
Dr Zubrin and The Mars Society have had decades to refine their concept of operations, subsequent to delivering their report to NASA, into something approaching an engineering solution using hardware NASA actually has on-hand. Dr Zubrin is an aerospace engineer and this is his pet project. Why is he still making the exact same sales pitch using concepts and hardware that are over 35 years old?
Take a look at the strength knock-down factors once you start welding 2195:
The Evolution of Constellium Al-Li Alloys for Space Launch and Crew Module Applications
NASA began exploring automated fiber tape laying machines and out-of-autoclave cured composite cryotanks to cut down on the incredible cost associated with machining and welding 2195. The switch back to 2219 for the SLS propellant tanks was for that reason. Try to imagine how expensive 2195 fabrication must be that using composites and automated fiber tape laying machines was 25-30% cheaper. Part of the cost could be explained by the fact that fabricating the composite tank takes about a week and 1 to 2 people are present babysitting the machine. Boeing, ULA, and others have an entire team of machinists and welders to fabricate Aluminum pressure vessels, and it takes them several months.
#143 Re: Single Stage To Orbit » SSTO Engine Technology » 2025-05-16 16:35:22
Dr Clark,
What’s the dry mass of this stage?
Robert Clark
It depends upon what you include.
116,120kg can be sent to orbit using the RD-180's Isp and mass flow rate.
Required propellant volume is 1,875.496m^3 if your engine has RD-180 Isp and mass flow rate.
Composite 8.4m Space Shuttle External Tank equivalent (larger than required propellant volume):
16,186kg
If you use 3X complete 10m diameter 634m^3 volume composite cryotanks of the sort that NASA designed about 10 years ago:
9,111kg
200:1 TWR LOX/RP1 engines for a 1.5:1 liftoff TWR (Space Shuttle liftoff TWR was ~1.2:1):
17,261kg (regeneratively-cooled metal alloys with de Leval nozzles; LOX/RP1 engines with Raptor 3 TWR)
116,120kg (ttl mass) - 16,186kg (8.4m Composite tank mass) - 17,261kg (200:1 TWR engine mass) = 82,673kg (useful payload)
116,120kg (ttl mass) - 9,111kg (10m Composite tank mass) - 17,261kg (200:1 TWR engine mass) = 89,748kg (useful payload)
You will have to subtract additional mass from the useful payload mass for:
1. engine thrust structures and propellant feed lines
2. electrical power subsystem
3. flight control avionics
4. payload fairing, if an expendable launch vehicle intended to launch a payload into orbit
5. pressurized accommodations for crew, if any
6. heat shield materials for full reusability
7. landing gear for full resuability
#144 Re: Single Stage To Orbit » SSTO Engine Technology » 2025-05-16 15:58:32
COMPOSITE CRYOTANK TECHNOLOGIES AND DEVELOPMENT 2.4 AND 5.5M OUT OF AUTOCLAVE TANK TEST RESULTS
Abstract
The Composite Cryotank Technologies and Demonstration (CCTD) project substantially matured composite, cryogenic propellant tank technology. The project involved the design, analysis, fabrication, and testing of large-scale (2.4-m-diameter precursor and 5.5-m-diameter) composite cryotanks. Design features included a one-piece wall design that minimized tank weight, a Y-joint that incorporated an engineered material to alleviate stress concentration under combined loading, and a fluted core cylindrical section that inherently allows for venting and purging. The tanks used out-of-autoclave (OoA) cured graphite/epoxy material and processes to enable large (up to 10-m-diameter) cryotank fabrication, and thin-ply prepreg to minimize hydrogen permeation through tank walls.
Both tanks were fabricated at Boeing using automated fiber placement on breakdown tooling. A fluted core skirt that efficiently carried axial loads and enabled hydrogen purging was included on the 5.5-m-diameter tank. Ultrasonic inspection was performed, and a structural health monitoring system was installed to identify any impact damage during ground processing. The precursor and 5.5-m-diameter tanks were tested in custom test fixtures at the National Aeronautics and Space Administration Marshall Space Flight Center. The testing, which consisted of a sequence of pressure and thermal cycles using liquid hydrogen, was successfully concluded and obtained valuable structural, thermal, and permeation performance data. This technology can be applied to a variety of aircraft and spacecraft applications that would benefit from 30 to 40% weight savings and substantial cost savings compared to aluminum lithium tanks.
If you read through the rest of NASA's documentation about their composite cryotank demonstrator program, which featured entries from Boeing, Lockheed-Martin, and Northrop-Grumman, what you will note is that the tank's minimum mass was dictated by the internal pressurization or hoop stress applied by the higher pressure LH2 tank. Despite being much heavier, the LOX tank is less highly stressed than the LH2 tank because the internal pressure and strain limit of 5,000 micro-inches per inch is driving the minimum required propellant tank mass to fabricate LOX and LH2 tanks. The strain limit apparently prevents microcracking in the presence of cryogens, or at least LH2, over repeated pressurization cycles. LH2 was the "challenge cryogen" of choice because it represents the most aggressive propellant in common use for space flight applications.
The results of this project are already more than 10 years old. We now have improved materials like Toray Composites America's T1200 fiber. We have improved resins from Cycom and other providers that supply materials to NASA, Boeing, Lockheed-Martin, Northrop-Grumman, and SpaceX.
#145 Re: Single Stage To Orbit » SSTO Engine Technology » 2025-05-16 15:28:44
NASA Game Changing Development Program - Composite Cryotank Technologies and Demonstration Project
Page 16: Orthogrid Stiffened 33 Foot Diameter Tank Design
Material: Al-2195 Aluminum-Lithium alloy
Internal Volume: 22,396ft^3 (634m^3)
Height: 413in (10.49m)
Weight: 10,925lbs (4,956kg)
Page 19: 10m Composite Cryotank
Material: Hexcel IM-7 and Cycom 5320-1 epoxy resin
Internal Volume: 22,396ft^3 (634m^3)
Height: 417.6in (10.6m)
Weight: 6,696lbs (3,037kg)
Operating Presure: 42psi (290kPa)
LH2 Weight: 99,072lbs (44,938kg)
Full Tank Weight: 105,768lbs (47,976kg)
#146 Re: Single Stage To Orbit » SSTO Engine Technology » 2025-05-16 15:12:01
Total Impulse to Send a Space Shuttle to Orbit
6,508,946,390N-s represents the Total Impulse generated by the historical Space Transportation System, aka "Space Shuttle", to deliver a 116,120kg wet mass orbiter vehicle, which includes the crew mass and the useful payload mass in the cargo bay, to orbit.
LOX/RP1 to Deliver 6,508,946,390N-s
Real World Engine Proxy: RD-180 engine's Isp (90% of Vacuum Isp) / thrust / mass flow performance figures?:
LOX: 1,597,846kg (1,192.422m^3)
RP1: 587,443kg (683.074m^3)
Total: 2,185,289kg (1,875.496m^3)
2,301,409kg LOX/RP1 + 116,120kg (vehicle and useful payload) = 2,301,409kg GLOW
116,120kg = 5.05% of GLOW
Composite vs Al-2195 Space Shuttle External Tank
LOX Tank: 19,744 cu ft (559.1 m3) at 22 psi (150kPa)
LH2 Tank: 53,488 cubic feet (1,514.6 m3) at 29.3 psi (202 kPa)
Total Volume: 2,073.699m^3 (198.203m^3 in excess of what's required for 1,875.496m^3 of LOX/RP1)
Space Shuttle Super-Light Weight Tank (Al-2195 alloy): 58,500lbs (26,535kg)
Composite Tank (IM7 CFRP using Cycom 5320-1 epoxy, out-of-autoclave curing): 35,685lbs (16,186kg)
Hexcel IM7 Carbon Fiber Tensile Strength: 820ksi
Toray Composites America T1200 Carbon Fiber Tensile Strength: 1,160ksi (preferred fiber for added tensile strength)
Composite tank mass reduction will come from autoclave curing and vacuum bagging to siphon off excess epoxy resin.
Engine Tech Requirements
We need 200:1 Thrust-to-Weight Ratio (TWR) RP1 staged combustion engines. The Cold War era RD-180 doesn't have a TWR that high, but Merlin-1D is much more representative of what modern materials and manufacturing methods can deliver. Raptor 3 has already achieved 200:1 TWR in a full-flow staged combustion engine design, so using a denser fuel than Methane should make the fuel turbo machinery smaller / lighter for the same mass flow rate, which strongly implies that 200:1 TWR LOX/RP1 engine is technologically achievable.
Rotating Detonation Wave Engines don't appear to add much practical Isp benefit, which is what everyone absolutely fixates on, but they literally cut the engine mass in half for a given level of thrust, which strongly implies that 300:1 to 400:1 TWR is achievable using the same engine materials.
Using non-regeneratively cooled Reinforced-Carbon-Carbon engine materials for all major components drastically cuts engine mass for any kind of engine to 1/4 of that using Iron-based or Nickel-Copper-based alloys, which strongly implies 800:1 TWR.
Liftoff TWR and Engine Tech TWR
2,301,409kg GLOW * 1.5 liftoff TWR = 3,452,113.5kg-f (liftoff thrust required)
3,452,113.5kg-f / 200 = 17,261kg (regeneratively-cooled metal alloys with de Leval nozzles)
3,452,113.5kg-f / 800 = 4,315kg using (non-regeneratively-cooled ZrC / NbC / HfC internally coated RCC with de Laval nozzles)
3,452,113.5kg-f / 1,600 = 2,158kg (ZrC / NbC / HfC internally coated RCC RDWE)
Useful Payload Mass / % of GLOW
IM7 prop. tanks + metal de Laval nozzle engines: 82,673kg / 3.59%
IM7 prop. tanks + RCC de Laval nozzle engines: 95,619kg / 4.15%
IM7 prop. tanks + RCC RDWE engines: 97,776kg / 4.25%
#147 Re: Human missions » Mars Ascent Vehicle - LOX/LCO vs LOX/LCH4 » 2025-05-16 13:14:29
RobertDyck,
The RCS weighs 500kg for both the ERV and tuna can, but unlike the ERV, the tuna can is going to rotate, de-rotate, and both vehicles will perform mid-course correction burns, yet the RCS mass estimate is exactly the same for both, at least according to that table, despite the ERV and tuna can having different masses and Delta-V requirements.
100kg for ERV comms / avionics for a space capsule moving through deep space for 6 months, but 200kg for the tuna can?
NASA estimates 241kg for the avionics and electrical system of a MAV that will be occupied for 3 days at most, using 2020s vs 1980s electronics. Therefore, these electronics mass estimates for vehicles that will be occupied for months to years seem naively optimistic. Don't get me wrong, I abhor piling on pointlessly heavy and expensive electronic gadgetry when it's not needed, but this just seems like someone tossed out a number backed by nothing.
The ERV cabin total mass is only 2X what it was for the Apollo Command Module (5,703kg), and fabricated from the same alloys, but that thing is going to be in deep space for 6 months with 4 vs 3 astronauts?
No. Just... No.
Where are the mass estimates for the additional insulation and cryocoolers and power subystems to keep 6,300kg of LH2 cold on its way to Mars?
The spares mass margin weighs more than the mass of the electrical subsystem?
The ADEPT heat shield is not going to protect the upper structure of the ERV or the tuna can, so I presume both have beefed up TPS mass added somewhere since "nothingness" is not going to protect against leeward side reentry heating.
What is the mass figure of the retro-propulsion system to land those things and why is that not included in the table?
They clearly aren't landing by parachute, because I don't see any parachute system mass included anywhere.
All the probes we've sent to Mars have back-shell TPS on them because the reentry heating on the leeward side, that would otherwise be imparted to the probe, despite being in the "shadow" of a considerable heat shield facing the windward side, would still melt them. GW has explained this several times, but I don't think anyone else was paying attention to him. The various capsules that reenter here at Earth all have TPS on their sides for the very same reason. The top of the Space Shuttle orbiter also had considerable TPS covering its cargo bay for the very same reason.
Reentry physics doesn't work any differently for tuna can habitats behind deployable heat shields vs robotic probes vs cargo bay doors attached to the backside of the orbiters, merely because we want to show very low mass figures.
If this is what Dr Zubrin and his team came up with for Martin-Marietta after a year or two, then I no longer wonder why NASA just kept after development of space station hardware. There are enough glaring omissions here that I'm not going to waste more of my time on this. I like his concept because I want to send humans to Mars, but no EDL hardware mass estimate included for the ERV or tuna can is telling me that this is a concept-of-operations, rather than an engineering exercise to figure out what's needed, how much everything will weigh, and thus how much propulsion hardware is required to deliver it to Mars.
If all of that wasn't enough, the ERV mass shown still weighs more than the mass of a fully fueled MAV using storable chemical propellants.
Why is it not beyond obvious that a spacecraft designed to stay in space with humans onboard for 26 weeks vs 1 week will obviously weigh more when it has to land on the surface of Mars, make its own cryogenic fuels from scratch, and then return to Earth?
Absent full reusability, using indigenous materials to make propellants for a super-sized Apollo capsule is a side quest that has nothing to do with exploration. It's an interesting idea, but one we should pursue after we prove that we can go to Mars and return to Earth.
#148 Re: Single Stage To Orbit » SSTO Engine Technology » 2025-05-15 20:06:00
Dr Clark,
Look at how long a high vs low TWR vehicle takes to climb out of the atmosphere, execute the roll-pitch program / gravity turn, and then how long it takes to achieve orbital velocity. Let's use the Space Shuttle as an example. A nominal ascent profile for the Space Shuttle lasted between 510 and 520 seconds.
Cd / aerodynamic drag coefficient matters quite a lot if you don't ascend out of the sensible atmosphere quite rapidly.
g0 / Earth's gravitational acceleration constant matters quite a lot if it takes longer to vertically ascend out of the sensible atmosphere, but then it matters a lot less after you complete your gravity turn and start moving down range at a rapidly increasing velocity.
Adding or subtracting 10 seconds of Isp is almost meaningless for a LOX/LH2 engine. A straight Delta-V calculation will show that the difference between 452.3s of Isp and 464.9s of Isp is 85.65m/s of additional acceleration possible for the same propellant load and dry vehicle mass. You will not see any sort of night-and-day payload performance improvement by using that extendible nozzle with the RS-25, but your vehicle will have significant additional dry mass. Unless some other part of your vehicle becomes lighter to compensate for lugging heavier engines all the way to space, then you may actually lose a little payload performance. If we were starting in orbit with 464.9s vs 452.3s of Isp, with full propellant tanks, then yes, you would see a meaningful payload performance increase, but again, it's not going to be a night-and-day type of improvement, especially if the nozzle extension mass is significant.
For a SSTO, adding or subtracting 50 to 80 seconds of Isp matters a lot less if that loss in Isp performance means your vehicle can accelerate twice as fast as it could using a much higher Isp propellant combo that cuts your engine TWR in half, and therefore acceleration in half, as is the case with LOX/LH2 vs LOX/RP1 or LOX/LCH4.
The reason LH2 engines have such pedestrian TWRs compared to RP1 and LCH4 engines has everything to do with the mass and volume of propellant that must be pumped per second to generate equivalent kinetic energy as engines producing heavier exhaust products. LH2's density is so lousy that you'd have to throw a lot more of it out the back of the engine to achieve equivalent thrust because the exhaust product is so much lighter, relatively speaking. You have to just about double the mass flow rate of a LH2 engine to match the kinetic energy of RP1 or LCH4. While you don't need to precisely double the mass of the pumps to make that happen, in actual practice the mass of a LH2 fueled engine as a whole has to just about double to achieve thrust values similar to RP1 or LCH4 engines. Raptor 3 is a 200:1 TWR engine. RS-25 will never be a 200:1 engine unless it's made from radically lighter CMC materials. Any CMC materials that could make the RS-25 a 200:1 TWR engine would make the Raptor 3 or Merlin 400:1 TWR engines.
Use the link shown above to download the high resolution "TIFF" formatted image file version from the Library of Congress, for the image shown below:
If you look at the propellant mass of the SRBs and compare it to the LOX/LH2 propellant mass in the Space Shuttle External Tank, what you'll note is that the majority of all propellant is consumed during the first few minutes of flight, in order to clear the atmosphere. The same applies to the Saturn V, Falcon, Starship, and most other rockets.
Adding an extra 10 seconds of Isp to the engines on the Space Shuttle using a complex nozzle extension simply does not buy very much additional payload performance for several reasons:
1. The additional mass of the nozzle extension has to be carried all the way to orbit. When they made the External Tank (ET) lighter, since it was carried all the way to orbit, ET mass reduction resulted in a nearly equal increase in payload performance to orbit.
2. The additional nozzle extension mass retards vehicle acceleration when the most propellant is being burned, though obviously not by very much.
3. A very healthy chunk of the total propellant load has already been consumed by the time you attain an altitude where the nozzle extension can be used to add thrust through more optimal exhaust gas expansion (Isp improvement).
#149 Re: Human missions » Mars Ascent Vehicle - LOX/LCO vs LOX/LCH4 » 2025-05-15 18:56:30
RobertDyck,
The Mars Direct ERV was a 40,000kg class payload sent to Mars using a rocket with Saturn V class payload performance. I've never seen anything approaching a detailed mass breakdown for the Mars Direct ERV, nor merely it's precise external dimensions and dry mass to know if any of the mass and performance estimates are believable. Regardless, the TLI / TMI capability for a Starship with an expendable upper stage is roughly 50,000kg, same as a Saturn V, but with fewer staging events. That means we can send 1X ERV or 2X fully fueled MAVs, powered by NTO/MMH or NTO/HAN, to the surface of Mars.
NASA is not going to cram 4 to 6 people into a vehicle as small as the ERV for a 6 month return trip to Earth. NASA has an ISS habitat module design derived from legacy ISS habitat module hardware for the crew to make the trip to Mars and return to Earth. That is how the real Mars exploration missions will be done, assuming they're done at all.
Mars mission steps using NASA-developed Mars / Moon / Deep Space exploration hardware
1. The Interplanetary Transfer Vehicle (ITV) is launched from Earth aboard a Starship Super Heavy expendable upper stage, into an ISS-like orbit, broadly similar to how Saturn V was used to launch Skylab. The ITV will be powered by a combination of Chemical Propulsion Modules (CPMs) and an onboard Solar Electric Propulsion (SEP) module.
2. A CPM is launched aboard a second Starship Super Heavy expendable upper stage. The CPM could be a Starship expendable upper stage or some kind of bespoke design optimized to provide ITV propulsion. Either way, it's going to use existing engines.
3. The Mars exploration crew rendezvous with the ITV / CPM stack aboard a Dragon or Starliner capsule or Dream Chaser mini shuttle launched by a Falcon 9 or Vulcan.
4. The ITV departs for Mars using CPM(s) attached to the ITV.
5. The ITV arrives in Mars orbit approximately 6 months later, using SEP to insert the ITV into orbit around Mars.
6. The ITV docks with one of several tuna can Surface Habitat Modules (SHMs) already waiting in orbit. The crew transfers to the SHM. The SHM takes the crew to the surface of Mars.
7.. The crew performs their exploration mission on the surface of Mars.
8. When it's time to come home, the crew boards one of several fully fueled MAVs waiting for them on the surface of Mars. The MAV returns the crew to the ITV.
9. The ITV departs for Earth and 6 months later arrives near Earth, whereupon the crew departs the ITV aboard their capsule or mini-shuttle for a direct reentry. The direct reentry is nearly identical to the reentry associated with a lunar mission. At the same time the ITV either uses its SEP module to slowly reestablish a circularized LEO orbit so that it can be refueled and refurbished for another mission. Alternatively, the ITV burns up during reentry.
10. The crew capsule either splashes down in the ocean or the mini-shuttle lands on a runway.
Some variation of that is how a real Mars mission will be done, because that is the mission hardware set that NASA is actively developing. Nobody else has made any serious attempt to assemble all the bits and pieces of required hardware. SpaceX only has a giant rocket that they're trying to use like a Swiss Army knife for every possible task, and we can already see how well that's going. At this point, I would settle for an expendable Starship upper stage that delivers 200t to a 400km orbit.
#150 Re: Interplanetary transportation » Universal Ion Propulsion » 2025-05-15 17:29:16
Calliban,
Dr Winglee's group developed a way to "tack" into the solar wind, much like a sailing ship.