kbd512 Postings

kbd512 · 2025-07-05 05:55:27

tahanson43206,

LOX/LH2 may be able to perform an expendable SSTO mission, but the real issue is the LH2 propellant volume and the low thrust-to-weight ratio (TWR) of LH2 fueled engines.

LH2 Engine TWR: 75:1
LCH4 and RP1 Engine TWR: 150:1 to 200:1

At the thrust levels demanded for a 100t class payload, about 1.5X whatever GLOW happens to be, using LH2 adds considerable engine mass.

LOX is 1,141kg/m^3. LH2 is 71kg/m^3. LOX has 16X the density of LH2 per unit volume, yet tank mass per unit volume of propellant stored will be the same between LOX and LH2, up to 3g of vehicle acceleration. LH2 internal pressurization loads drives propellant tank mass. This useful tidbit of information comes from NASA.

LOX/LH2 Isp (Vac): 450s
LOX/RP1 Isp (Vac): 360s
LH2 provides a 25% Isp increase over RP1.

1m^3 of LH2, at 71kg/m^3 and 142MJ/kg, stores 10,082MJ/m^3
1m^3 of RP1, at 810kg/m^3 and 43MJ/kg, stores 34,830MJ/m^3
RP1 provides a 345% volumetric energy density increase over LH2.

Can LH2's 25% fuel economy advantage over RP1, somehow overcome RP1's 345% energy density increase over LH2?

Delta IV (all LOX/LH2 stages): 249.5t GLOW; 8.51t payload to LEO (3.411% of GLOW)
Falcon 9 (all LOX/RP1 stages): 549t GLOW; 22.8t payload to LEO (4.153% of GLOW)
Delta IV Heavy (all LOX/LH2 stages): 733t GLOW; 28.79t payload to LEO (3.927% of GLOW)
Falcon 9 Heavy (all LOX/RP1 stages): 1,420t GLOW; 63.8t payload to LEO (4.493% of GLOW)

There's your answer. RP1 fueled vehicles are heavier because they take more payload to orbit than the LH2 fueled vehicles. Even as a percentage of GLOW, payload-to-GLOW is higher for RP1 than LH2. Both rockets are fabricated from Aluminum alloys and both use gas generator cycle engines.

At some point, stage volume, thus stage dry mass fraction, combined with low engine TWR, really starts to become a problem for LH2 because it eats into your payload mass fraction.

Do we imagine that these problems will become more or less severe as we both scale-up the rocket to deliver a 100t payload and push the entire vehicle mass from the ground, all the way to orbit?

Why is the payload mass fraction of Delta IV Heavy lower than a Merlin powered vehicle that never achieves the sea level Isp of the RS-68A and RL-10 engines that power the Delta IV and Delta IV Heavy?

kbd512 · 2025-07-05 21:14:24

Delta IV Common Booster Core
Dry Mass: 28,850kg
Dry Mass Less Engine(s): 22,110kg
Gross Mass: 226,400kg
Total Propellant Mass: 199,550kg (150m^3 of LOX, at 1,141kg/m^3, 171,150kg and 400m^3 of LH2, at 71kg/m^3, 28,400kg)
Fuel Stored Energy: 4,032,800MJ of stored energy, at 142MJ/kg
Dry Mass Fraction (Dry Mass / Gross Mass): 12.74293%
MJ of Stored Energy per kg Dry Mass (4,032,800MJ / 28,850kg): 139.78510MJ/kg

RS-68A Engine Mass: 6,740kg
RS-68A Mass Flow Rate: 1,400kg/s
RS-68A Sea Level Thrust: 3,137,115.1N (sl); 3,560,000N (vac)

Delta IV Common Booster Core Sea Level Total Impulse Calculation
199,550kg / 1,400kg = 142.53571 seconds of total firing time
142.53571 seconds * 3,137,115.1N = 447,150,942N-s of total impulse
447,150,942N-s / 28,850kg = 15,499N-s total impulse per kg of dry stage mass
447,150,942N-s / 226,400kg = 1,975N-s total impulse per kg of gross stage mass
447,150,942N-s / 199,550kg = 2,241N-s total impulse per kg of propellant

Delta IV Common Booster Core Vacuum Total Impulse Calculation
142.53571 seconds * 3,560,000N = 507,427,128N-s of total impulse
507,427,128N-s / 28,850kg = 17,588N-s total impulse per kg of dry stage mass
507,427,128N-s / 226,400kg = 2,241N-s total impulse per kg of gross stage mass
507,427,128N-s / 199,550kg = 2,543N-s total impulse per kg of propellant

Notional Delta IV Common Booster Core equipped with 2X RS-25D Engines
Dry Mass: 28,850kg - 6,740kg + (3,177kg * 2) = 28,464kg
Dry Mass Less Engine(s): 22,110kg
Gross Mass: 228,014kg
Total Propellant Mass: 199,550kg

RS-25D Engine Mass: 3,177kg
RS-25D Mass Flow Rate: 514.49kg/s
RS-25D Thrust: 1,860,000N (sl); 2,279,000N (vac)

RS-25D Powered Delta IV Common Booster Core Sea Level Total Impulse Calculation
199,550kg / (514.49kg/s * 2) = 193.92991 seconds total firing time
193.92991 seconds * (1,860,000N * 2) = 721,419,265N-s total impulse
721,419,265N-s / 28,464kg = 25,345N-s total impulse per kg of stage dry mass
721,419,265N-s / 228,014kg = 3,164N-s total impulse per kg of stage gross mass
721,419,265N-s / 199,550kg = 3,615N-s total impulse per kg of propellant mass

RS-25D Powered Delta IV Common Booster Core Vacuum Total Impulse Calculation
193.92991 seconds * (2,279,000N * 2) = 883,932,530N-s total impulse
883,932,530N-s / 28,464kg = 31,054N-s total impulse per kg of stage dry mass
883,932,530N-s / 228,014kg = 3,877N-s total impulse per kg of stage gross mass
883,932,530N-s / 199,550kg = 4,430N-s total impulse per kg of propellant mass

Falcon 9 Booster Core
Dry Mass: 25,600kg (includes ~2,000kg of reusability hardware)
Dry Mass Less Engine(s): 21,370kg
Gross Mass: 436,500kg
Total Propellant Mass: 410,900kg (287,400kg densified LOX and 123,500kg densified RP1)
Fuel Stored Energy: 5,310,500MJ of stored energy, at 43MJ/kg
Dry Mass Fraction (Dry Mass / Gross Mass): 6.07102%
MJ of Stored Energy per kg of Dry Mass Fraction (5,310,500MJ / 25,600kg): 207.44141MJ/kg

Merlin 1D Engine Mass: 470kg (4,230kg for all 9 engines)
Merlin 1D Mass Flow Rate: 229kg/s
Merlin 1D Thrust: 657,000N (sl)

Falcon 9 Booster Core Sea Level Total Impulse Calculation
410,900kg / (229kg/s * 9) = 199.36924 seconds of total firing time
199.36924 seconds * (657,000N * 9) = 1,178,870,316N-s
1,178,870,316N-s / 25,600kg = 46,050N-s per kg of dry stage mass
1,178,870,316N-s / 436,500kg = 2,701N-s per kg of gross stage mass
1,178,870,316N-s / 410,900kg = 2,869N-s per kg of propellant

When we compare a staged combustion LH2 fueled engine with a gas generator RP1 engine, we finally arrive at a total impulse per unit of stage gross mass value favorable to LH2, but notice how that total impulse per unit of stage dry mass value is still well below RP1. For a SSTO, the stage dry mass is going all the way to orbit with the payload, so the less stage dry mass that you need for a given total impulse, the more payload you get to deliver to orbit. If we remove the reusability hardware mass from the Falcon 9 booster core's dry mass, LH2 will fall even shorter.

If the reusability hardware (grid fins, landing legs, cold gas thrusters, etc) weighs ~2,000kg, then the Falcon 9 Stage Dry Mass Fraction is only 5.40664%, so all of the other performance numbers measurably improve as well.

I'm not going to bother with Merlin-1D's vacuum thrust or a staged combustion RP1 fueled engine because the point is already made. Total Impulse delivered per unit of stage dry and gross / wet mass, is higher for a Falcon 9 booster core with its 9 Merlin engines, at sea level, than it is for the Delta IV booster core with its RS-68A engine in a vacuum.

This simple math comparison uses real engine and stage performance numbers applicable to real rocket booster core stages, with the exception of the notional RS-25D powered Delta IV Common Booster Core, thrown in just to show that not even a staged combustion LH2 fueled engine beats a gas generator RP1 fueled engine in terms of total impulse delivered per unit of dry stage mass.

In this example, RP1 provides substantially greater stored energy per unit of dry stage mass AND gross stage mass, despite LH2's clear Isp advantage over RP1. LH2 gross / wet mass is clearly much lower, yet stage dry mass is NOT. We can substitute-in whatever stage fabrication materials we wish to use for both vehicles, but if we use the same materials for both RP1 and LH2 powered vehicles, that result is NOT GOING TO CHANGE! In the same way that gravimetric energy density has major impact on the performance of battery operated aircraft vs those powered by hydrocarbon fueled engines, volumetric energy density greatly affects the total impulse delivered by rockets fueled by RP1 vs LH2.

The "proof" is in the delivered payload tonnage to LEO per unit of vehicle dry mass fraction. It's higher for RP1 than it is for LH2, despite the fact that the gross / wet mass of the RP1 fueled stage is much higher than it is for the LH2 fueled stage. For a SSTO, this matters even more than for a TSTO. Orbital rocketry is about converting mega-joules of stored energy into thrust, for a given dry vehicle mass fraction, which in turn determines how much payload mass fraction you get.

For a SSTO, you do not get to "delete" 2/3rds of the vehicle's total dry mass on the way to orbit. For these two real rockets powered by real engines, RP1 clearly provides more stored energy to convert into thrust than LH2 does, per unit of vehicle dry and gross / wet mass fractions.

In this example, our stage dry mass per Newton-seconds of total impulse delivered, for RP1, is 297% that of LH2. Our gross / wet mass per Newton-seconds of total impulse delivered for RP1 is still 137% greater than it is for LH2. Despite that fact, our dry stage mass is still lower for RP1, even though we're carrying 240% more propellant mass in the RP1 fueled stage, as compared to the LH2 fueled stage.

LH2's Isp advantage over RP1, when burned by the RS-68A, cannot overcome RP1's density impulse and Merlin-1D's thrust-to-weight advantage. I didn't have to guess at anything, extrapolate anything, or use advanced rocketry knowledge.

tahanson43206 · 2025-07-06 20:50:22

For kbd512 re Gemini evaluation of carbon vs LH2

I was surprised to see that your proposal appears to hold up fairly well. I am not surprised the LH2 solution held up. What ** is ** interesting is that Gemini ended up leaning slightly in your direction.

(th)

kbd512 · 2025-07-09 02:37:41

Let's consider the energy efficiency of RP1 vs LH2, purely from an energy consumption per kilogram of useful payload perspective.

D4M = Delta IV Medium (all stages fueled with LH2)
F9B5 = Falcon 9 Block V (all stages fueled with RP1)

Note: D4M is a single Common Booster Core powered by a single RS-68A, with no strap-on solid rockets, and an upper stage powered by a single RL-10. D4M's vehicle dry mass is significantly greater than a F9B5, despite delivering less than half the payload to LEO of a F9B5. Fuel mass for both stages of D4M is 4.9X less than both stages of F9B5, and total propellant mass for the D4M's Common Booster Core is almost exactly 50% that of a F9B5 booster core. One would think that the superior Isp of Hydrogen burning engines and half the propellant mass would translate to a lighter dry stage mass for the D4M Common Booster Core, yet the opposite is true. Both stages of D4M are significantly heavier than both stages of F9B5. Both the propellant tanks and engines of D4M are heavier than the propellant tanks and engines of F9B5.

F9B5 Energy Economics
Fuel Masses: 123,500kg (booster); 34,275kg (upper stage); 157,776kg (total)
157,776kg of RP1 * 43MJ/kg = 6,784,368MJ of total fuel energy
6,784,368MJ / 22,800kg to LEO (fully expended) = 297.56MJ/kg of payload to LEO
297.56MJ / 43MJ/kg = 6.92kg (2.24 gallons of RP1 per 1kg of payload to LEO, at 3.084kg/gal)

D4M Energy Economics
Fuel Masses: 28,400kg (booster); 3,763kg (upper stage); 32,163kg (total)
32,163kg of LH2 * 142MJ/kg = 4,567,146MJ
4,567,146MJ / 8,510kg to LEO (expendable-only) = 536.68MJ/kg of payload to LEO
536.68MJ / 142MJ/kg = 3.78kg (14.12 gallons of LH2 per 1kg of payload to LEO, at 3.733gal/kg)

LH2 vs RP1 Payload-to-Orbit Fuel Efficiency Comparison
536.68MJ/kg (D4M/LH2) / 297.56MJ/kg (F9B5/RP1) = 1.80360

D4M vs F9B5 Payload to Orbit
8,510kg (D4M payload to orbit) / 22,800kg (F9B5 payload to orbit) = 0.37325

D4M vs F9B5 Booster Total Impulse
447,150,942N-s (D4M) / 1,178,870,316N-s (F9B5) = 0.37930

D4M vs F9B5 Booster Dry Mass
28,850kg (D4M) / 25,600kg (F9B5) = 1.126953125

Conclusions
D4M consumes 80% more fuel energy than F9B5 while delivering only 38% as much payload to orbit, with a mass increase of 13% over F9B5. On top of that, D4M also costs more than F9B5, which is why it was retired. If you increase the booster dry mass of the LH2 solution by 338%, using a triple booster core Delta IV Heavy configuration, then you can get 5,990kg more payload to orbit than F9B5.

Key Takeaways
Whatever propellant mass efficiency advantage LH2 provides over RP1, it clearly doesn't translate into lower vehicle dry mass, lower total energy consumption, nor even improved payload performance. If anything, the opposite appears to be true. As the size of the vehicle scales-up to deliver more payload to orbit, that vehicle dry mass problem only gets worse, never better, because materials don't become any stronger as you increase the size of a structure. You will see a relative improvement in vehicle dry mass as you increase the propellant tank diameters, thanks to the cube-square law, yet any LH2 powered vehicle is still going to end up being significantly heavier than a RP1 powered vehicle for the same delivered payload mass.

Even when we compare the total impulse of a notional RS-25 powered D4M operated purely in a vacuum vs a Merlin powered F9B5 operated purely at sea level, the total impulse generated per unit of vehicle dry mass advantage is lopsidedly in favor of F9B5.

If you invoke a D4M with Carbon Fiber propellant tanks with only 60% of the mass of the actual historical D4M's Aluminum propellant tanks, powered by a pair of lighter and higher-Isp RS-25D engines, you will still fail to generate a total impulse per unit of dry vehicle mass figure that's higher than a Merlin powered F9B5, constructed primarily of Al-2195 alloy. That's how horrendously awful LH2 is as a fuel choice for a SSTO or TSTO.

Total impulse generated per unit of dry vehicle mass, not propellant mass, is the figure of merit that either makes or breaks SSTOs, and the primary reason why there are none. That figure of merit is either sky-high, as it would be for a RP1 fueled rocket with Carbon Fiber propellant tanks and staged combustion engines, or SSTO doesn't work.

Calliban · 2025-07-09 03:18:03

This suggests that overall propellant volumetric energy density is a more important consideration than mass energy density. Musk chose LNG propellant as a compromise between the mass energy density of H2 and the volumetric energy density of heavier hydrocarbons.
https://commons.m.wikimedia.org/wiki/Fi … to-license

Hydrogen is even weaker than these figures show it to be. A great deal of infrastructure and additional energy are needed to manufacture it. Storing and handling hydrogen on the ground is also very costly and dangerous.

kbd512 · 2025-07-09 13:44:58

Calliban,

As an engineer, do you understand all the calculations I presented to evaluate Total Impulse per unit of vehicle dry mass?

tahanson43206,

Let's compute the volumetric energy density efficiencies for the following fuels:
LCH4 (not the same thing as LNG), at 422.8kg/m^3 and 55.5MJ/kg
HTPB (solid propellant), at 920kg/m^3 and 20MJ/kg
N2H4 (Hydrazine), at 1,021kg/m^3 and 19.5MJ/kg
HFO (Heavy Fuel Oil / "Bunker Fuel"), at 1,010kg/m^3 at 41MJ/kg

LH2: 71kg/m^3 * 142MJ/kg = 10,082MJ/m^3
HTPB: 920kg/m^3 * 20MJ/kg = 18,400MJ/m^3
N2H4: 1,021kg/m^3 * 19.5MJ/kg = 19,909.5MJ/m^3
LCH4: 422.8kg/m^3 * 55.5MJ/kg = 23,465.4MJ/m^3
RP1: 810kg/m^3 * 43MJ/kg = 34,830MJ/m^3
HFO: 1,010kg/m^3 * 41MJ/kg = 41,410MJ/m^3
CBP (Carbon Black Powder): 1,700kg/m^3 * 32.8MJ/kg = 55,760MJ/m^3

Volumetric Energy Density of Alternative Fuels Compared with RP1
CBP: 160.092%
HFO: 118.892%
RP1: 100%
LCH4: 67.371%
N2H4: 57.162%
HTPB: 52.828%
LH2: 28.946%

When over 90% of the vehicle's volume is filled with propellant, as is the case for all orbital class rocket vehicles, vehicle dry mass becomes sensitive to propellant density. Since CBP and HFO would pose a severe coking problem for a rocket engine, and HFO is also loaded with Sulfur in most cases, RP1 is the most energy dense and readily available propellant for minimizing vehicle dry mass.

There are no clever "tricks" to skirt around major volumetric energy density differences when fractional improvements in propellant mass efficiency can't overcome volumetric inefficiencies which are multiples of propellants with much higher volumetric energy densities. You need as high an Isp as you can manage, but more importantly, you need sufficient density to minimize vehicle dry mass if you intend to accelerate 100% of the vehicle dry mass, plus the useful payload, to orbital velocities. The only known way to do that is to minimize the vehicle's volume, which is enabled by RP1 and hampered by far less dense propellants like LH2 and LCH4, or propellants with much lower thermal output per unit of mass, such as Hydrazine and solids like APCP and HTPB.

tahanson43206 · 2025-07-09 17:27:54

For kbd512 re #306

RP1 is the most energy dense and readily available propellant for minimizing vehicle dry mass.

This is a statement that deserves a moment of focus.

I am hoping you (or someone) will focus intently on this statement, and provide a simple 100 word or so confirmation it is correct. I don't know if it is correct or not. That's a benefit of being an observer when situations like this come up.

What I'm imagining is a very simple comparison.

We could define some arbitrary amount of energy.

We should be able to accurately compute the volume and mass of that amount of energy for LH2 and for RP1.

What I am expecting is that the mass of RP1 would be far greater for two reasons.

However, the volume of LH2 would be far greater.

I wonder if you might be thinking about tankage to hold these two liquids?

(th)

kbd512 · 2025-07-10 00:22:25

tahanson43206,

Using RP1 (kerosene) as the fuel of choice permits rocket vehicle designers to minimize the propellant volume and maximize the total units of thrust generated per unit of engine mass and total vehicle dry mass for an orbital launch vehicle. This feature of RP1 is particularly important for SSTOs, but also important for TSTOs.

Vehicle dry mass scales by engine thrust-to-weight ratio and propellant tank volume, because for any significant payload to orbit, the propellant tank mass represents the majority of the vehicle's total dry mass (engines + propellant tanks). The cube-square law does help limit propellant tank mass increase as the propellant tank volume increases, but the end result will still be heavier propellant tanks and therefore higher vehicle dry mass when substantially less dense propellants, such as LH2, are substituted for RP1, when the goal should be equal or better payload performance with RP1. The vehicle dry mass comparison between Delta IV Medium and Falcon 9 Block 5 outright proves this. This is merely another way of saying that you cannot increase propellant tank volume by 10% or more, on account of LH2's exceptionally low volumetric density, yet still somehow arrive at a lower vehicle dry mass, as compared to RP1.

D4M Booster vs F9B5 Booster Propellant Volume and Tank Mass Comparison
D4M: 450m^3; mass, less engine: 22,110kg
F9B5: 370m^3; mass, less engines: 21,370kg or ~19,370kg without reusability hardware installed
F9B5 tank mass is 87.61% that of D4M
F9B5 tank volume is 82.22% that of D4M
F9B5 payload to orbit is 2.68X greater than that of D4M, but only 2X the gross liftoff mass of D4M

D4M Upper Stage vs F9B5 Upper Stage Propellant Volume and Tank Mass Comparison
D4M: 76m^3; dry mass, less engine 3,189kg; 23,264kg LOX + 3,956kg (561,752MJ) LH2
1X 301kg RL10B-2 engine, 110.1kN, 465.5s Isp; 24.1kg/s mass flow rate
1,129.460580912863071s * 110,100N = 124,353,610N-s
226,770kg total LOX/LH2 (both stages) / 8,510kg = 26.64747kg propellant per 1kg of payload
28,850kg + 3490kg = 32,340kg total dry mass

(447,150,942N-s + 124,353,610N-s) = 571,504,552N-s Total Impulse (both stages)
571,504,552N-s / 8,510kg = 67,157N-s/kg of payload

F9B5: 98m^3; dry mass, less engine 3,450kg; 75,200kg LOX + 33,200kg (1,427,600MJ) RP1
1X 550kg Merlin-1D Vac engine, 801kN, 348s Isp; 236.6kg/s mass flow rate
454.353338968723584s * 801,000N = 363,937,025N-s
518,400kg total LOX/RP1 (both stages) / 22,800kg = 22.73684kg propellant per 1kg of payload
25,600kg + 4,000kg = 29,600kg total dry mass (with booster reusability hardware installed)
23,600kg + 4,000kg = 27,600kg total dry mass (expendable)

(1,178,870,316N-s + 363,937,025N-s) = 1,542,807,341N-s Total Impulse (both stages)
1,542,807,341N-s / 22,800kg = 67,667N-s/kg of payload

Total Propellant Volume Required per kg of Useful Payload
D4M: 526m^3 of total propellant volume; 16.17871kg of payload per 1m^3 of propellant volume
F9B5: 468m^3 of total propellant volume; 48.71795kg of payload per 1m^3 of propellant volume

Total Thrust Required per kg of Useful Payload
D4M: 67,157N-s/kg of payload
F9B5: 67,667N-s/kg of payload

D4M "beats" F9B5, a vehicle pushing 2X as much total mass into the sky, in this one category that doesn't help make the case for LH2 fueled SSTOs, nor TSTOs for that matter. It's fractionally better than RP1 here, meaning you could run the main engine(s) for about 1.5 seconds longer, or less, on a LH2 fueled vehicle before total thrust (an indirect measure of total energy expenditure) per kg of delivered payload equalizes between both vehicles.

Hopefully we now understand how greatly vehicle dry mass affects our vehicle's total energy expended per kg of useful payload. We're getting 3X more payload mass per cubic meter of propellant volume for the F9B5 when we include both stages of the D4M and F9B5 rockets. That's a reasonably good efficiency metric if you ask me. Rather than showing how great LH2 is, we're continuously illustrating how freakishly efficient a modern RP1 fueled vehicle truly is. Even for a TSTO, every little bit of vehicle dry mass increase has a serious effect on vehicle performance.

RP1 has an acceptable Isp (total units of thrust generated per unit of propellant mass expended) for SSTO and TSTO. It's not even close to being "the best", yet LH2's much higher Isp doesn't seem to help improve total vehicle dry mass. Delivery of payload to orbit is primarily a question of Total Impulse delivered per kilogram of vehicle dry mass, and this is where RP1 is a stand-out performer.

RP1 allows the vehicle's propellant tanks to occupy the least amount of physical volume, therefore reducing their mass, amongst all of the common liquid rocket propellants. Amongst the propellants I listed, RP1 (kerosene) is the most widely used energy dense fuel combusted in gas turbine engines. The turbopumps that feed liquid rocket engines are specialized gas turbines capable of force-feeding a tremendous mass of gaseous propellants into the main combustion chamber.

kbd512 · 2025-07-10 12:16:55

As the above comparison shows, our RP1 fueled TSTO is delivering 3X more payload mass per cubic meter of propellant volume, as compared to LH2.

Using the Silverbird Astronautics Launch Vehicle Performance Calculator:

Upgraded Staged Combustion Merlin-powered SSTO
Dry Mass: 17,088kg (40% lighter than Al-2195 tanks using CFRP; Merlin engine mass same as existing engines)
Propellant Mass: 518,400kg (same propellant mass as both stages of F9B5)
Thrust:5,913kN (same as Merlin-1D sea level thrust X9)
Isp: 324 seconds (Merlin-1D converted to staged combustion 360s * 0.9 for time-avg thrust output to orbit)
Launch Site: Cape Canaveral / KSC (USA)
Default Propellant Residuals?: Yes
Restartable Upper Stage?: No

Orbit: 185km by 185km, 45deg inclination

Estimated Payload: 10,654kg; 95% Confidence Interval: 5,704kg to 16,639kg

Why might we still want an expendable SSTO version of Falcon 9?

If we can figure out how to get the engines back, once on-orbit, CFRP tanks could be made 10% lighter than the mass estimate provided here for a single use. Total stage mass could be as little as 14,940kg, with 10,710kg being mostly CFRP. At $100/kg, if we figure on throwing away $1M per launch to discard the propellant tanks while recovering the engines, then I think we can live with that.

For this lighter single-use vehicle...
Estimated Payload: 12,802kg; 95% Confidence Interval: 7,852kg to 18,784kg

For RCC vs stainless and Nickel-Copper engines:
1,057.5kg (engines) + 10,710kg (propellant tanks) = 11,767.5kg dry stage mass
Estimated Payload: 15,970kg; 95% Confidence Interval: 11,023kg to 21,957kg

50% stronger T1200 fiber with CNT-infused resin vs IM7 fiber:
1,057.5kg (engines) + 6,426kg (propellant tanks) = 7,483.5kg dry stage mass
Estimated Payload: 20,254kg; 95% Confidence Interval: 15,310kg to 26,241kg

If we can make the tanks light enough, then at some point we give up on full reusability because it's cheaper to just discard the propellant tanks, collect the engines with a space tug, and return them to Earth using a HIAD. Alternatively, we accumulate some engines in space for use to propel ships to other destinations. Storing RP1 in space is relatively easy.

Someone also recently devised a method to make RCC 100X cheaper than it previously was, so fabricating high-temp engine components from RCC is now a plausible means to greatly reduce engine mass and cost.

SSTO will never be as-performant as TSTO, but the payload performance difference between the partially reusable TSTOs we have today and the expendable SSTOs we could have tomorrow may not be that great. If the cost is lower than the cost of stage recovery, that's still a "win" in my book.

tahanson43206 · 2025-07-10 13:02:47

for Kbd512 re #309 and others...

Re carbon tanks to orbit / return the engines ...

Can those carbon tanks be repurposed for use as habitats?

I would imagine they might, but they would require some enhancement, I would think.

Why not just melt the engines down and reuse the atoms on orbit?

The concept of SSTO delivery of fuel and supplies to a refueling depot might blend nicely with the SSTO concept if total reuse of material is possible.

PS ... GW sent me a detailed analysis of tank weight for various propellants, taking oxygen into account.

I have it saved in a folder called "tanks" is we want to look at it on Sunday.

Tankage for LH2 weighs more than tankage for CH4, but oxygen is involved and the amount may not be the same for the two fuels.

There didn't seem to be an easy winner to pick from the data.

(th)

kbd512 · 2025-07-10 14:40:16

tahanson43206,

The propellant tanks could be retained in-orbit to serve as the basis of a propellant depot, but only if all the rockets are taking payloads to the same orbit. Falcon 9 is a general purpose workhorse rocket. If all the payloads Falcon 9 delivers to orbit are going to different orbits, then you're not going to have them all in one place.

You want to "melt down" an engine made from RCC? Why not use it again as an engine?

Using RCC (2g/cm^3) serves two purposes, the first being to greatly reduce engine mass to a value about 1/4 that of stainless (8g/cm^3) or Nickel-Copper (8-9g/cm^3) alloys. The second is that it does not require regenerative cooling to survive the temperatures the engine generates. NASA has tested uncooled RCC engine nozzles burning LH2 / LCH4 / RP1 fuels. They coat the RCC nozzle with UHTCs to prevent oxidation at extreme temperatures. Melting them down is not easy. You need electrical power to do it. Think about it. RCC survives combustion temperatures of LH2 / LCH4 / RP1 fuels without cooling, doesn't melt or disintegrate, and survived multiple firings without refurbishment.

CFRP propellant tanks built strong enough to survive LH2 internal pressurization are also built strong enough to carry LOX and survive 3g accelerations. 52psi vs half that much for LOX and RP1 is an extreme amount of force.

52psi = 36,560kgf/m^3 (drastically more force per unit area than 3g acceleration or aero loads during ascent)

Lockheed-Martin's box-stiffened CFRP tank concept has a 634m^3 capacity, max design pressure of 46.2psi, and it weighed 6,573lbs / 2,981kg. Therefore, the proposed tank mass of 10,710kg is quite realistic as a stage mass (tanks, engine mounts, sensors), and probably gross overkill for 468m^3 of LOX/RP1 tank capacity. I'm talking about loading 518,400kg of propellant in total, and the load is definitely distributed over more than 14m^2, so the idea that this proposed tank somehow isn't strong enough is bunk. It's probably far stronger than it needs to be, yet still a minor fraction of the mass of the original TSTO F9B5 tank mass, less engines.

Sub-scale testing results with the 5.5m diameter tanks:

In other tests, the tank successfully maintained fuels at extremely low temperatures and operated at various pressures. Engineers filled the tank with almost 30,000 gallons of liquid hydrogen chilled to -423F, and repeatedly cycled the pressure between 20 to 53 pounds per square inch -- the pressure limit set for the tests.

30,000 gallons = 113.562m^3 (not too large, but not small, either)

3 different designs from the 3 major aerospace primes. All of them passed all tests.

kbd512 · 2025-07-10 15:08:55

The Rocket Labs Electron rocket is LOX/RP1 TSTO with propellant tanks made entirely from CFRP. It's successfully delivered payloads into orbit 64 times now. The turbopumps are electric and use Lithium-ion batteries to power them. Engine Isp is right up there with the best staged combustion engines, meaning 311s to 343s. I think we can say that it works now.

tahanson43206 · 2025-07-11 12:21:37

For kbd512 re post in SSTO topic:

https://newmars.com/forums/viewtopic.ph … 67#p232767

It is time to begin developing the proposal for a carbon based solution for SSTO. If the data you've collected and shown stands up to Real Universe demands, your design should win funding easily.

Your design will include specifications for all the inert mass of the vehicle, along with the quantities of carbon fuel and LOX you'll need.

In addition, your design will include prescriptions for engine performance and operation throughout the flight.

Your design will include details of diameter and length of the vehicle, and shape of the nose.

Your design will include specifications of engines, pumps, plumbing and tanks, and the manner in which tanks are integrated into the hull to achieve the required performance.

Because you have shown that carbon is capable of achieving LEO in an SSTO design, it should be possible to find a funder interested in that capability willing to build and fly the system.

NewMars is unproven as a venue for serious project work, but your initiative provides an opportunity for a test of capability.

It may turn out that a blend of NewMars posts and other cloud resources will be needed to capture all the elements.

(th)

kbd512 · 2025-07-11 13:14:58

Let's evaluate how using the densest form of Carbon, Carbon Black Powder (CBP), does or does not "work" for rocketry:

The F9B5 booster contains 123,500kg of RP1. We want to evaluate how much CBP is required to deliver equivalent thermal energy, to "know" if CBP's spectacular density of 1,700kg/m^3 can reduce total propellant volume and thus vehicle dry mass.

123,500kg of RP1 * 43MJ/kg = 5,310,500MJ
123,500kg of RP1 / 810kg/m^3 = 152.469m^3 RP1 volume
123,500kg of RP1 / 867kg/m^3 = 142.445m^3 Densified RP1 volume

5,310,500MJ / 32.8MJ/kg (CBP) = 161,905kg of CBP fuel (31% more mass than RP1 for equal thermal output)
161,905kg CBP / 1,700kg/m^3 of CBP = 95.238m^3

161,905kg * 2.67kg of LOX = 432,286.35kg of LOX
432,286.35kg (LOX) + 161,905kg (CBP) = 594,191.35kg of total propellant mass
432,286.35kg LOX / 1,141kg/m^3 = 378.866m^3 LOX volume
432,286.35kg LOX / 1,262kg/m^3 = 342.541m^3 Densified LOX volume

95.238m^3 + 342.541m^3 = 437.779m^3 total propellant volume (CBP Isp is clearly too low, so both mass and volume increases)

Here we can see the effect of low gravimetric energy density on low-Isp high volumetric density fuels. The density of CBP is fantastic, but 32.8MJ/kg is insufficient gravimetric energy density to result in a propellant tank volume reduction for a rocket powered vehicle. The Isp is so low that both propellant mass and volume have increased, as compared with RP1, which means CBP is of no practical use for a SSTO, nor TSTO for that matter. CBP will produce a spectacular amount of thrust as a strap-on hybrid solid rocket booster, much like APCP and HTPB, but the quantity of LOX required is so great that the net effect is to increase the propellant tank volume and thus vehicle dry mass to achieve a given level of payload performance.

Solids are frequently used in conjunction with LH2 to provide that critical but missing thrust at liftoff to "get the rocket moving smartly downrange", as GW puts it. However, solids of the size useful for orbital launches are neither cheap nor rapidly reusable. If the motor casing refurbishment and propellant refill was cheap enough, then we'd figure out an approach for reusability, such as having a stock of motor casings on-hand. Unfortunately, large solids are expensive and time consuming to manufacture, with the propellant poured in multiple batches, so they're more useful for infrequent exploration missions and military weapons than routine orbital launch with a high launch cadence. If CBP was very cheap / easy to produce, and was far less dangerous to transport without an oxidizer mixed into the fuel, then it might have wider utility.

Oddly enough, a German rocket company is working on a hybrid solid rocket uses CBP mixed into Paraffin wax to control burn rate and prevent melting, with LOX as the oxidizer, giving them a solid rocket with a throttle and Isp almost identical to RP1. It's technically interesting because a LOX/LH2 core stage could hold additional LOX to feed the solid boosters, which are essentially inert for transportation purposes and very compact since they lack the oxidizer. LOX is more effective than traditional toxic oxidizers used in solids, such as Ammonium Perchlorate, meaning it increases the Isp of the solid well above what is possible with APCP and HTPB. It would be possible to "ring" a core stage with what are essentially "fuel only" solid rocket boosters, and then to increase the length / height of the core stage to hold more LOX. You could put the LOX in a tank above the booster's fuel grain, but then it's more complex and difficult to recover, or more expensive to discard.

tahanson43206 · 2025-07-15 06:14:43

For kbd512 re new Iterative Rocket Design for LH2

https://newmars.com/forums/viewtopic.ph … 39#p232839

To help our readers keep track of who said what, GW Johnson is the advocate for 100,000 kg to LEO for an LH2 vehicle.

I set up the topic with a stated goal of 1 metric ton as the payload above and beyond the vehicle itself.

That said, I like the idea of having a topic that is focused upon 100,000 kg because that should be an attractive number for serious customers.

The simplest solution is to simply change the goal of the LH2 topic to 100,000 kg, and create a new topic for the 1 ton variant.

(th)

kbd512 · 2025-07-15 08:05:52

tahanson43206,

I'm genuinely curious to know if LH2 works better as a fuel than RP1 as you scale-up the size of the rocket to deliver a 100t payload. I don't have any ideological beliefs about this. I only care about what the basic math tells us about the nature of the problem. If LH2 energy economics work more favorably at 100t than 25t, then that's what the ultimate answer is, regardless of what I or anyone else believes about it. Math won't change to suit anyone's beliefs.

I've done the math on Delta IV Medium and Falcon 9 Bock V to evaluate Total Impulse per kilogram delivered. I could get more sophisticated with the analysis, but really don't need to. What it tells me is that for the medium lift payload delivered, RP1 is the fuel of choice. However, when I look at Delta IV Heavy and Falcon Heavy, I see a clear advantage attributable to LH2. Delta IV Medium / Heavy and Falcon 9 / Falcon Heavy are the closest "apples-to-apples" real world rockets we have to compare / contrast LH2 and RP1.

For the boosters, I evaluated Total Impulse of each stage at sea level only. I'm going to use 90% of vacuum Isp, which is pretty standard for booster evluation, and re-run the numbers. The point to the sea level only evaluation was that if LH2 didn't show a reduction in Total Impulse per kilogram of payload delivered there, then it wouldn't show one with 90% of vacuum Isp, either, since it goes up quite a bit for LH2. However, what I see thus far seems to favor LH2 for much larger payloads, because the simplistic sea level Total Impulse evaluation for Delta IV Heavy and Falcon Heavy indicates much lower Total Impulse required to deliver a kilogram of payload to orbit, relative to Falcon Heavy. That means Delta IV Heavy is actually markedly more energy efficient than Falcon Heavy, which is critical for both a SSTO and for any LH2 fueled vehicle to make good economic sense. LH2 is a premium fuel, easily double the cost of RP1 per gallon. If we're going to use it, then we need to make a case for why it's more efficient to use it. For a small payload, we cannot make that case. For a much larger payload, initial results look very promising.

For a 25t payload or less, LH2 doesn't seem to provide a favorable energy trade, which indicates that the combination of thrust "to get the vehicle moving smartly downrange", as GW would say (what he means is that sluggish acceleration is not acceptable for a SSTO), and cube-square isn't working in our favor for smaller rockets, because the RP1 fueled rocket can be so much smaller and lighter, relatively speaking, that the engine and/or propellant tank mass ratios don't work in favor of LH2.

As with most other things in engineering, everything from gas turbine engine efficiency to economical fission reactors, it appears that scaling laws apply to SSTOs as well.

tahanson43206 · 2025-07-30 06:45:44

For kbd512 re analysis of aluminum energy storage system developed in India...

Recharging it requires about 12X to 14X more energy in than you get out, because the Aluminum-oxide has to be converted back into Aluminum powder. Energy density is pretty fantastic. It's essentially a rebuildable fuel cell. Using it takes away any argument about the battery being more efficient than a combustion engine, because it's not. A gasoline powered car will use ~937.5kWh of fuel energy to drive the same distance, meaning 500W per mile. Converting 100kg of Aluminum-oxide powder, back into Aluminum powder, 1,200kWh to 1,400kWh.

The analysis you provided is certainly interesting.

What I don't see is an apples to apples comparison.

An apples to apples comparison would show how much energy is needed to convert carbon dioxide and water back to gasoline.

My guess is that whatever method of energy storage is used will have exactly the same full cycle energy flows.

The difference would be whether a particular user is "cheating" by consuming resources stored by Nature.

The Indian plan (as I understand it) is to use a full cycle process, and NO stored resources at all.

If you are inspired by this objection to perform the math, please publish a comparison of energy flows between the Indian concept and your favorite hydrocarbon fuel.

One of the two systems will be more efficient than the other, when compared in a full cycle mode. I will be watching to see which comes out of the comparison.

(th)

kbd512 · 2025-08-02 03:00:39

tahanson43206,

The difference would be whether a particular user is "cheating" by consuming resources stored by Nature.

Metals, coal, oil, and gas are all natural resources which have been "stored by Nature". If growing hydrocarbon foods in the soil vs full artificial chemical synthesis of sugars and starches is not considered "cheating", then neither is mining coal out of the ground. It's a false double-standard that speaks to a desperate attempt to contort fundamental reality for ideological reasons. If you need food and fresh water, that primarily comes from the ground. If you need metal, then you're still talking about digging in the ground. Similarly, the densest forms of energy also come from the ground. All the machines that extract and transport bulk resource are only possible because they're powered by fuels that come from the ground. To replace these underground resources with resources extracted from the air or oceans requires orders of magnitude more energy and materials, which must initially come from the ground, since they don't presently exist in the required quantities.

Onwards...

My personal "favorite" fuel is pure Carbon mixed into water, essentially Coal-Water Slurry (CWS). It's not prototypical "coal" in the sense that it's not the stuff we dig out of the ground, it's high purity Carbon powder made from CO2 bubbled through a column of Gallium-Copper eutectic, and then mixed into water. Pure Oxygen is released in that room temperature process while pure Carbon powder floats to the top where it can be easily collected. The O2 could be retained for Oxy-Fuel combustion, if so desired, given that only compression is required for storage.

We commonly refer to the powder generated by that process as "Carbon Black", which is normally produced by combusting Methane. Carbon Black is an essential rubber additive found in all motor vehicle tires, Carbon / Graphite is used by most battery anodes, it can serve as the base stock for Carbon Fiber or Reinforced Carbon-Carbon (RCC) for enhanced lightweight motor vehicle components such as brakes. Last but certainly not least, Carbon Black mixed into fresh water creates a very high flashpoint fuel with much greater gravimetric energy density than any existing battery technology, to include Aluminum-air. CWS is capable of powering electric power plants, ships, semi-trucks, mining equipment, trains, and passenger motor vehicles using sCO2 gas turbines. CWS produced in this manner is not a refined petroleum product, nor is it dug out of the ground.

sCO2 gas turbines are high performance / high efficiency external combustion engines, similar in concept to steam engines, except they use sCO2 instead of water as their working fluid, making them about 10X more compact than steam turbines for a given power output. They're also noticeably more compact and much quieter than traditional gas turbines, which are already quite compact for the power they generate. Perhaps most importantly, achievable efficiency using sCO2 vs steam as the working fluid is greater across a broader power band.

The refining process for synthetic Carbon Black would start with Direct Air Capture, Direct Plant Capture from sCO2 electric power plants using RamGen to compress the CO2 exhaust effluent into LCO2, or possibly ocean capture methods in countries with access to oceans. Regardless, acquiring CO2 is not particularly problematic for the parts of humanity that consume the most hydrocarbon fuels.

This is the Google AI answer to how much energy is required to convert Alumina-oxide into Aluminum metal:

Theoretical minimum energy: The chemical reaction enthalpy, representing the theoretical minimum energy needed to reduce alumina to aluminum, is approximately 8.1 kWh/kg or 29.2 MJ/t (megajoules per tonne) of aluminum. This translates to roughly 6.23 kWh/(kg of Al) when considering the standard Gibbs energy of formation.
Practical energy consumption: In actual industrial settings, the Hall-Héroult process typically requires more energy due to factors like inefficiencies in the electrolysis cells and the need to overcome electrical resistance. Modern aluminum smelters require close to 13 kWh to produce 1 kg of aluminum, while the world average may be closer to 14 kWh/kg Al. Some sources report the practical annual operating minimum between 12.9 DC kWh/kg Al and 13.2 DC kWh/kg Al. Best-practice settings require around 13,000 kWh/ton (47 GJ/ton).
Breakdown of energy consumption:
Smelting: The most energy-intensive step, consuming about 68.6% of the total energy, or roughly 193.6 MJ/kg of aluminum.
Alumina production (Bayer process): Refining bauxite into alumina requires approximately 25 GJ/tonne of aluminum, or around 7.26 kWh/kg Al.
Other stages: Bauxite mining and carbon anode production also require energy, albeit in smaller proportions.
Comparison with recycling: Recycling aluminum is significantly more energy-efficient than producing it from virgin materials. Recycling aluminum saves about 95% of the energy required for primary production. This translates to approximately 14,000 kilowatt-hours of energy saved per ton of recycled aluminum.
Impact of energy source: The carbon footprint associated with aluminum production varies depending on the energy source used to generate the electricity. Smelters using renewable energy sources like hydroelectricity can have lower carbon emissions compared to those relying on coal-fired power plants.

Anyone who is truly interested can read more about Aluminum production here:
U.S. Energy Requirements for Aluminum Production - Historical Perspective, Theoretical Limits and Current Practices

The section to read starts on Page 39 of 150 of the above PDF file, and is entitled "5. Primary Aluminum Production". Since the document contains clickable links that will take you directly to the section of interest, I recommend using them.

Globally, the total input energy ranges between 12kWh and 17kWh per kg of pure Aluminum metal. Most plants burn coal or natural gas to generate the electricity, but some use hydro dams. Aluminum smelters run 24/7/365, and absolutely cannot shut down due to intermittent energy sources, else the cost of bringing the plant back online is often uneconomical.

CO2-to-Carbon
Room-temperature CO2-to-carbon conversion facilitated by copper-gallium liquid metal

Hightlights:
1. Room temperature CO2 reduction to solid carbon with Cu-based system.
2. Liquid metal Ga enhances electron transfer for CO2 reduction.
3. GaIn-Cu catalysts yield 2598.7μmol/h of solid carbon with 100% selectivity.
4. Catalyst system stable for 360 hours, promising for industrial use.

I will presume that any process which proceeds at a temperature of only 40C doesn't use too much energy.

This is the Google AI answer to how much energy is required to capture CO2 directly from the air:

Overall DAC Energy Consumption: Contemporary assessments indicate that DAC systems necessitate an energy input ranging from 2000 to 3000 kWh (or 7.2 to 10.8 GJ) to sequester one metric ton of CO2, according to Springer. This includes the energy for fans to move air, the heat needed for sorbent regeneration, and the electricity required for CO2 compression.

While ocean capture systems may consume significantly less energy, air capture can be used anywhere there is dense air, which would be everywhere at or near Earth sea level. Air and power plant exhaust effluent CO2 capture are the most practical options available, with ocean capture being an attractive option which may be more practical for nations with extensive coastlines and infrastructure capabilities, such as America and China.

The energy cost to produce deionized (DI) water is relatively low, typically around $0.00278 per gallon, or $2.78 per 1,000 gallons, according to one industry estimate. This cost is associated with the electricity used in the deionization process, assuming a rate of $0.10 per kWh. However, the total cost of DI water can vary depending on several factors, including the specific deionization system used, the source water quality, and the scale of production.

~3,783kg of water per 1,000 gallons.
3,783kg / 27.8kWh = 136kg of deionized water per kWh
27,800Wh / 3,783kg = 7.348665W/kg

By my math, that means Coal-Water Slurry (CWS) requires about 2kWh to 3kWh per kg of CO2, the pure Carbon production requires very little energy input, plus the energy for the water. CWS is 60% to 70% coal by volume.

3.79kg per gallon of water * 0.65 = 2.4635kg of pure Carbon

2.4635kg * 3kWh/kg = 7.3905kWh of input energy
1.3265kg of deionized water * 7.348665W/kg = 9.748004Wh of input energy

2.4635kg * 32,800,000J/kg = 80,802,800J = 22,445Wh per gallon of CWS

We should be able to get 50% of that energy back out using a sCO2 gas turbine engine, which will be positively tiny for a car, given that a 250kW turbine wheel is roughly the size of a Silver Dollar. I presume we'd connect the power turbine to a flywheel, and then connect the flywheel to a geared CVT transmission. We'll tack on another 15% worth of losses within the drivetrain. That seems pretty close to reality, since overall efficiency is 35%, about the same as a modern gasoline fueled piston engine. I'm sure someone will argue with that figure, but they're arguing over well-established engineering principles.

Geared CVT:
Screenshot-2024-12-18-091901-copy.jpg?quality=100&w=2048

Traditional CVTs use belts or chains that stretch and wear out over time, but a geared CVT can continuously adjust its gearing ratio like a traditional CVT while using more compact and durable spur-type gears.

We end up with 7,855.75Wh of mechanical energy per gallon of CWS fuel, and we need to input about 7,400.248Wh of total energy. We need to add pipeline transport energy costs as well, which will be substantial in aggregate but are also highly variable since they're dependent upon how far the LCO2 and CWS must be pumped. Additional energy will be expended to transport CWS by truck to fueling stations. CO2 can presumably be captured from gas or coal-fired power plants to reduce total energy input. However, the actual refinement process for this "synthetic coal", produced using Gallium-Copper eutectic, comes pretty close to energy "break-even". Turning Alumina into Aluminum requires 2,000C temperatures. After more than a century, that's still the basic process for bulk Aluminum production, and it still requires Carbon anodes that produce CO2, regardless of how the input electricity has been generated. In contrast, turning CO2 into pure Carbon requires 40C temperatures. That is primarily what dictates how much input energy is required for recycling Alumina vs CO2.

I wanted to use synthetic coal CWS because the fuel is readily storable, not particularly hazardous since it's the same material found in fresh water filtration systems, with the water that this "coal dust" has been mixed into being the only "volatile" part of the fuel. The Carbon particles don't readily fall out of suspension after they've been mixed into the water. Since the Carbon is mixed into water, it's exceptionally difficult to accidentally ignite CWS, unlike any other hydrocarbon fuel. A CO2-derived pure Carbon CWS contains no significant impurities like Sulfur or heavy metals or radioactive gases like Radon. Recycling atmospheric CO2 into fuel is a truly sustainable process, because CO2 and water is so abundant on Earth, relative to metals. Synthetic CWS requires only storage tanks and pipeline networks to efficiently deliver the precursor materials and fuel, same as crude oil, gasoline, diesel, and kerosene. Plain Carbon steels are entirely suitable. We already pump Methane and Ammonia around the country, so we can also pump LCO2 and CWS just as easily.

If the energy required to convert Alumina into Aluminum can be reduced to something approaching the theoretical minimum energy, then maybe pursuing that option makes more sense. If Tesla's rechargeable Aluminum-air battery has the energy density of the latest Lithium-ion batteries, then it still results in vehicles that are heavier than existing gas powered vehicles. sCO2 powered vehicles would use power train components substantially lighter and more compact than all existing gasoline and diesel engines, which means the vehicles could get lighter instead of heavier. A CWS fuel spill is not particularly hazardous, and could be washed off a roadway with water. Technically, this high purity Carbon fuel can be ingested in small quantities without much in the way of health effects. All other common fuels pose health and flammability hazards, to include alcohol.

Teslas use ~250W/mile (not the absolute most efficient but not the worst, either), which implies a 375kg battery at 300Wh/kg at the pack level, for a vehicle with a 112.5kWh pack capable of achieving a driving range of 450 miles. That's about identical range to existing gas powered compact / subcompact commuter car, none of which are powered by engines that weigh as much as an all cast Iron Big Block Ford engine. People who are ideologically motivated can try to fight physics all they want, but in the end they're going to pay significantly more money for that excess weight, in terms of purchase and maintenance costs. Every base model gas powered truck is about $10K to $15K more than a base model gas powered sedan, because moving and transforming natural resources into materials and then components cost money and energy.

That Cadillac Escalade my wife drove more or less proves that excess weight costs a lot more money. She doesn't drive it anymore. She drives a Toyota RAV4 now. For very heavy vehicles, such as BEVs or oversized SUVs, the tires and brakes never last as long as the warranty period says they should and they always cost more to replace, specifically because they're so heavy and made from high embodied energy materials. Playing a convoluted game of pretend with efficiency doesn't change physical reality. There have been numerous disingenuous attempts to imply that electricity is "more efficient than Carbon-based fuels", but any time you increase ordering of energy forms you radically increase the energy and materials inputs. Any purely electrical system based upon photovoltaics, wind turbines, and batteries will require more materials and thus energy inputs than a system based upon chemical energy.

You buy a Tesla, then you accept that you're looking at new tires and brake pads every 12 to 18 months, you're paying 50% than what a comparably-sized but lighter gas powered vehicle costs. You're never recovering that additional money, in terms of gas money saved, over 10 years of normal driving. You either accept that at least 50% of your power is coming from burning something, likely closer to 75% if you mostly recharge your BEV at night when you're not using. Alternatively, you accept that Carbon-based chemical energy requires fewer materials and energy inputs than electrical energy, especially when recycling is an integral part of long-term sustainability.

If theory matched reality, then a country like Norway, where nearly all of the vehicles in active use are EVs, ought to see a 20% reduction in total oil consumption, rather than 10%. 10% would be a very big number for America, but hardly a game changer. Only big numbers matter, though. Room temperature CO2 recycling allows us to post very big numbers because a near-100% recycling rate implies that we're adding little to nothing to the atmosphere or oceans over time. Gradual changes using photovoltaics and wind turbines are already completely overwhelmed by the new coal and natural gas power plants in China, India, and Africa. The new coal power plants in China alone have already "cancelled out" more than 100% of the supposed CO2 emissions reductions of Western nations from photovoltaics, wind turbines, and batteries. That's why such solutions are utter nonsense. They squander lots of time and money while making negative progress towards their stated global CO2 emission goals. Unless our "green evangelists" are planning to genocide most of the people in developing nations, which would only indicate that they're more profoundly evil than I know them to be, this will continue to be the case. Recycling CO2 is the best option available, because it's one that doesn't require nonexistent electrical energy storage technology capable of matching the energy density of chemical fuels.

You can't successfully fight physics by continuously increasing the total mass of all vehicles, the energy intensity of the materials required to make them, and the energy generating systems that power them. Nothing about this Aluminum battery solution requires less total mass of materials than chemical energy. The grid infrastructure to power a fleet of primarily electric vehicles, without burning something, requires doubling or even tripling of its existing carrying capacity. Insisting upon the use of intermittent energy as primary sources only increases requirements, which is why all "green energy" systems must be backstopped gas turbines or nuclear reactors. Electrical systems do not last as long in operation as pipelines and pumps, either. There's not a week that goes by where I don't see the power company trucks and crews replacing a piece of electrical equipment heavier than a car. Ignoring all the energy inputs into continuous complete replacement, especially if little of it gets recycled, as is the case for all electronics, is intentionally deceptive. The only part that's shockingly real is the absurd cost. Electrical and electronic systems are truly fantastic when they work, but when they don't the only "solution" thus far is complete replacement of whatever failed. Nobody "repairs" batteries, they replace them. The fuel-cell-like Aluminum-air batteries might be the only notable exception.

Collectively, all of this new electrical and electronic technology doesn't do what it claims to do, namely reduce CO2 emissions, or merely reduce the rate of increase of CO2 emissions. Incentivizing CO2 capture and recycling by using simplistic fuels and processes has a much better chance of arresting the rate of increase and drawing down CO2 over time, because CWS can be stockpiled in simple steel storage containers. Rather than play along with an absurd idea that clearly isn't working, namely consuming more energy and materials to somehow eventually consume less, I'm proposing a more realistic alternative that could work because it uses existing technology.

Here in Texas we're building sCO2 gas turbines to pump our natural gas and to generate electricity. We needed something better than existing gas or steam turbines or piston engines. sCO2 provides that "something better". We spent the past 25 years developing sCO2 tech, and now we're finally seeing the first commercial applications. As near as we can tell, it actually works, it is meaningfully more efficient than traditional gas or steam turbines, and doesn't require quantities of materials we don't have.

tahanson43206 · 2025-08-04 06:24:19

For kbd512 re post in Google Meeting...

https://newmars.com/forums/viewtopic.ph … 61#p233261

That is a terrific suggestion! Thanks! It seems to me if GW's spreadsheet works with a well documented model, and makes a prediction that is close to what was actually observed, then we can trust it to try some of your ambitious ideas.

It is possible currently active companies may not want to make details of their designs public for competitive reasons. However, they ** might ** be willing to test GW's software in their own offices, using their own data, and provide feedback to GW about how well the software performed. That would be a major win, if we can achieve it.

(th)

tahanson43206 · 2025-08-05 08:48:07

For kbd512 re #319 and new request...

GW accepted your suggestion to prepare spreadsheets for calculation of stages for the Apollo as a proof method.

He has spreadsheets for stage design in the queue, but right now he is concentrating on the SSTO concept.

The requirements for an SSTO are sufficiently different, that (here I am guessing) perhaps the two might require a different approach?

***
Our new member is on the board, with an interesting post in one of our more active topics. I am inspired by his arrival to ask the question:

Could we convert all our "banned" Users to Inactive? SpaceNut was trying to upgrade users from banned to inactive, but he was doing it a few at a time, and he ran out of time. Could you write a database command that would convert all users currently banned to inactive? If you can, it would remove what I see as an embarrassment, because we have over 1000 members who were not actually banned. The banned state is a consequence of the Great Hacker Attack, and our (SpaceNut and mine) attempt to deal with it.

I think that if someone who was truly banned were to apply for return to active status, they would have to apply through the newmarsmember portal, so we could make a decision at that time. I'm hoping this might be something you could do that wouldn't take too much time?

(th)

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#310 2025-07-10 13:02:47

Re: kbd512 Postings

#311 2025-07-10 14:40:16

Re: kbd512 Postings

#312 2025-07-10 15:08:55

Re: kbd512 Postings

#313 2025-07-11 12:21:37

Re: kbd512 Postings

#314 2025-07-11 13:14:58

Re: kbd512 Postings

#315 2025-07-15 06:14:43

Re: kbd512 Postings

#316 2025-07-15 08:05:52

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#317 2025-07-30 06:45:44

Re: kbd512 Postings

#318 2025-08-02 03:00:39

Re: kbd512 Postings

#319 2025-08-04 06:24:19

Re: kbd512 Postings

#320 2025-08-05 08:48:07

Re: kbd512 Postings

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