New Mars Forums

We finally arrive at an energy economics result favorable to LH2:

Delta IV Heavy Energy Economics Improve Upon Delta IV Medium
447,150,942N-s * 3 (Common Booster Cores) + 124,353,610N-s (DCSS Upper Stage) = 1,465,806,436N-s Total Impulse (both stages)
1,465,806,436N-s / 28,790kg of payload = 50,914N-s/kg of payload to LEO

Falcon 9 Block V Heavy Energy Economics
1,178,870,316N-s * 3 (Booster Cores) + 363,937,025N-s (Upper Stage) = 3,900,547,973N-s Total Impulse (both stages)
3,900,547,973N-s / 63,800kg = 61,137N-s/kg of payload to LEO

50,914N-s/kg  / 61,137N-s/kg = 0.83

1. I take this to mean that the vehicle is very sensitive to thrust performance.   Tahanson43206 set the payload target to orbit at 100,000kg.  Therefore, we want to know the energy economics after we scale-up the LH2 powered SSTO to lift 100t.  There is at least some potential here which indicates favorable energy economics associated with using Hydrogen fuel.  At the 25t payload level or less, RP1 fueled solutions or LH2 paired with solid rocket motors appear to result in much smaller / lighter / equally performant vehicles in the realm of energy economics.  This partially explains why vehicles like Ariane V, Atlas V, and Delta IV Heavy have all failed to compete with Falcon 9 Block V.  For small launchers, no economy of scale is possible.  You either have highly favorable energy economics or the cost of your engines and fuels override whatever minor performance advantages exist.

2. This is an actual desirable result showing a clear advantage in energy consumption required to attain orbit.  Whereas the Delta IV Medium is little better than a Falcon 9 Block V, Delta IV Heavy consumes measurably LESS energy than Falcon 9 Block V Heavy.  Therefore, it must be the case that using LH2 as the fuel of choice doesn't produce the desirable energy economics until the launch vehicle size is scaled-up quite substantially.

3. To make LH2 more competitive, we need better performing engines, such as RDEs, made from somewhat exotic but now much more affordable materials like RCC, in order to minimize engine mass or improve engine TWR, however you prefer to think about it.  RDE alone would result in 150:1 TWR, and using RCC instead of stainless or Nickel-Copper alloys could increase that to 600:1, making LH2 engines suitable for SSTOs.

4. Any reusable SSTO must use CFRP propellant tanks.  No other material is suitable because TPS mass will be added for reentry protection.  However, I was considering the propellant mass fraction of the stainless steel balloon tanks used by the Centaur Upper Stage.  Recent advances in Mangalloy welding make me think that Mangalloy, which has double the yield strength of 304L stainless, could be a lower cost and stronger alternative to fabricate much larger SSTO propellant tanks.  5.1m diameter 304L tanks are clearly doable, but ideally we want 10m diameter tanks.  The Japanese and Koreans have done extensive testing with this material for storage of LNG, LN2, and LH2.  Since they have test data to back up their assertions, I'm going to take them at their word.  Some kind of coating, perhaps Silicon-based, would be required to protect the alloy from LOX, perhaps LH2 as well, but the end result could be relatively inexpensive, strong enough, and light enough.  I was thinking about using steel cable with weldments inside the tank that keep the structure in tension and pull double duty as slosh baffles.  I've never seen this done for this application, but we use it in bridges and other applications.

5. Someone needs to run an analysis to determine what flight trajectories are optimal for LH2 SSTOs, given their lower TWR.  Is there a payload performance benefit to adding some forward / perpendicular velocity during part of the flight within the lower atmosphere, so long as the aerodynamic heating and drag loss is not too severe?  If we're going to use steel balloon tanks, this might be more tolerable.  All LH2 tanks have some external thermal insulation applied so that heat transfer rates remain tolerable.  What kinds of lightweight thermal insulation do we have today?  Perhaps we have something based upon spray-on aerogels that minimize weight and do a better job than polyurethane foam.

LOX/LH2 Propellant for Equal Thermal Output as 1m^3 of Densified RP1 and 1.608m^3 of Densified LOX

NBP LOX = 1,141kg/m^3 (Normal)
MP + 10K LOX = 1,262kg/m^3 (Densified)

37,281MJ/m^3 of RP1 / 10,082MJ/m^3 of LH2 = 3.697778m^3 of LH2 per 1m^3 of RP1
3.697778 * 71kg/m^3 = 262.542238kg of LH2
262.542238kg of LH2 * 6.03 = 1,583.12969514kg of LOX
1,583.12969514kg (LOX) + 262.542238kg (LH2) = 1,845.67193314kg of LOX/LH2 propellant
2,895.78kg (LOX/RP1) / 1,845.67193314kg = 1.568957
LOX/RP1 is 56.8957% heavier than LOX/LH2 for equal thermal output

LOX/LH2 volume increases to:
3.697778m^3 (LH2) + 1.387493m^3 (LOX) = 5.085271m^3 total LOX/LH2 propellant volume
3.697778m^3 (LH2) + 1.254461m^3 (Densified LOX) = 4.952239m^3 total Densified LOX/LH2 propellant volume
4.952239m^3 (Densified LOX + NBP LH2) / 1.608m^3 (Densified LOX + Densified RP1) = 3.079751

Best case scenario, Densified LOX + NBP LH2 occupies 3X greater volume than Densified LOX + Densified RP1 for equal thermal energy output.

Will the propellant tank volume trade for LH2 be quite that bad, due to the much higher Isp of LOX/LH2?

No.  A lighter / faster combustion product means less mass flow at a higher velocity partially equalizes the dramatically lower bulk propellant density of LOX/LH2.  This counts for something, but not enough to favor LH2.

Will a LOX/LH2 powered vehicle dry mass ever be as light as a LOX/RP1 powered vehicle for the same payload to orbit?

No.  The propellant itself will be lighter for equal kinetic energy output and the cube-square law will also partially offset the increased propellant tank mass to store LH2, but the stage dry mass must be higher when using LH2.  There are no two ways about this.

This stage dry mass increase is accurately reflected in the stage dry mass and payload to orbit of Delta IV Medium (all LH2 powered) and Falcon 9 (all RP1 powered).  Both vehicles have very similar stage dry mass, although the Delta IV Medium stage dry mass is higher for both stages, after subtracting the engine mass.  LH2 engines are also heavier than RP1 engines for the thrust output provided, but that's merely another aspect of the compounding problem of low thrust and high stage dry mass.  Delta IV Medium delivers 8,510kg to orbit.  Falcon 9 delivers 22,800kg to orbit.  The fact that the Falcon 9 (549.5t) is about 2X as heavy (with its full propellant load) as the Delta IV Medium (224.6t) has no significant effect on stage dry mass.

549.5t (Falcon 9 wet mass) / 224.6t (Delta IV Medium wet mass) = 2.44657
22,800kg (Falcon 9 payload) / 8,510kg (Delta IV Medium payload) = 2.67920

224.6t (Delta IV Medium wet mass) / 8.51t (Delta IV Medium payload mass) = 26.39248
549.5t (Falcon 9 wet mass) / 22.8t (Falcon 9 payload mass) = 24.10088

The ratio of Falcon 9 wet mass to payload mass is LOWER than it is for Delta IV Medium.

How can that be the case?

The higher stage dry mass of the LH2 powered vehicle (the combination of higher propellant tank mass and much lower engine TWR) offsets more than 100% of its considerable Isp advantage over the RP1 powered vehicle.  This problem would be worse for a SSTO.

The major "clue" that something doesn't add up is the simple fact that the Total Impulse delivered by both stages of the Delta IV Medium, per kilogram of payload delivered to orbit, is only very modestly lower than it is for both stages of Falcon 9.  The only way for that to be true is if the relative propulsive efficiency (total number of Newton-seconds of thrust delivered per kilogram of payload) associated with using LH2 in a real rocket powered vehicle is very similar to RP1.  If that number for LH2 was dramatically lower than it was for RP1, then it's indicative of much greater relative propulsive efficiency.

Delta IV Medium: 67,157N-s/kg of payload
Falcon 9 Block V: 67,667N-s/kg of payload

510N-s of Total Impulse reduction per kg of payload.  That's the extent of what LH2's fantastic "efficiency" provides to a real world Delta IV Medium rocket, which is an almost meaningless figure, because it's equivalent to extending the firing duration of the booster engine(s) in either vehicle for 1.5 seconds or less for either vehicle.  I was expecting a N-s/kg reduction commensurate with the Isp increase of LH2 over RP1 for the engines in question, meaning N-s/kg should be at least 29% lower for LH2.  RS-68A produces a higher Isp at sea level than the Merlin-1D Vacuum nozzle engine variant produces in a hard vacuum.  In real life, RP1's thrust efficiency in a Falcon 9 Block V is less than 1% lower than LH2 in a Delta IV Medium, by computing total impulse delivered by both stages and then dividing by the total kilograms of payload to orbit.  That is spectacularly bad, but only reinforces how important engine thrust and stage dry mass truly are to overall vehicle performance.

If we wanted to make the payload performance of Delta IV Medium and Falcon 9 "equal", what does that imply?

22,800kg / 8,510kg = 2.6792

Stage dry mass is going to increase by about 2.6792X to make a LH2 gas generator cycle engine vehicle "equal" to a RP1 gas generator cycle engine vehicle.  The payload to LEO of Delta IV Heavy, which is about 3X heavier than Delta IV Medium, is only 5,990kg higher than Falcon 9.

Delta IV Medium
26,850kg (1X Common Booster Core) + 3,480kg (5.1m DCSS Upper Stage) = 30,330kg (total dry mass)
30,300kg dry mass / 8,510kg payload mass = 3.56kg dry mass per 1kg payload mass

Delta IV Heavy
80,550kg (3X 5.1m Common Booster Cores) + 3,480kg (5.1m DCSS Upper Stage) = 84,030kg (total dry mass)
84,030kg dry mass / 28,790kg payload mass = 2.92kg dry mass per 1kg payload mass

Falcon 9 Block V
23,600kg (expendable booster core) + 4,000kg (upper stage) = 27,600kg (total dry mass)
27,600kg dry mass / 22,800kg payload mass = 1.21kg dry mass per 1kg payload mass

Falcon 9 Block V Heavy
70,800kg (3X expendable booster cores) + 4,000kg (upper stage) = 74,800kg (total dry mass)
74,800kg dry mass / 63,800kg payload mass = 1.17kg dry mass per 1kg payload mass

How do we make LH2 stage dry mass "more equal" to RP1 stage dry mass?

That's the real question we need to answer to make LH2 a viable candidate fuel for SSTOs.

For a practical purely rocket powered SSTO, what does that mean?

If you insist on using LH2, then you get a lot less payload to orbit (that is not the vehicle itself) for the same stage dry mass.  I can't speak for anyone else, but that looks like a bad trade to me.  All that "potential" thrust efficiency (N-s/kg of payload) amounts to almost nothing due to the poor stage dry mass efficiency, exacerbated by poor LH2 engine TWR.  RDEs and RCC vs metal alloys could potentially solve the TWR issue for LH2 engines, but they cannot reduce the propellant tank mass, which is the largest portion of total stage dry mass.

RobertDyck,

The US does engage in trade, and always has, but it's not a major part of our economy, except for the part that drains resources and jobs to actually buy the products.  If trade was such a great deal, then there would be far fewer people living in relative poverty in America, not more.

I haven't been in a Wal-Mart for many years.  I honestly have no clue where their products are made these days.

Edit:
Back when we still shopped at Wal-Mart, my wife bought dishes, glasses, and other household items (towels, blankets, linen), specifically because they were "Made in USA".

Your parents paid $25 in 1984 for a device which I can buy for $10 in 2025:
EZ DUZ IT - USA Made Can Opener - $9.95

It's available on Amazon with a $4 markup.  There are other American-made can openers which are also available for less than $25.

I feel like you were a little over-zealous in typing out all those zeroes.  You're "off" by a factor of 10.

Purism's US-assembled Smart Phone with US-made Electronics and US-sourced Materials:
Purism Liberty Phone

Purism's 128GB Liberty Phone: Starts at $1,999
Apple's 128GB iPhone 16 Pro: Starts at $999

Edit #2:
Most of the Apple mobile chips are now made in the US before being shipped to China for final assembly into iPhones and other Apple products, then shipped back to the US.  TSMC moved their chip fab to Arizona citing security concerns about having all of their factories in Taiwan.

Edit #3:
Amprius now has 500Wh/kg Lithium-ion batteries, which are made in Brighton, Colorado.  This is far in excess of what the present iPhones have in them, but will obviously help power the next generation of mobile devices.  Purism could source their batteries from Amprius, for example, because their batteries are user-replaceable by design.

Yes they are and I just proved it.  Actually, the real world market economy just proved it.  I didn't have to prove anything.  You should've done some basic research before making your outrageous claims, but you didn't because basic research wouldn't support your claims.

RobertDyck,

Isn't President Trump doing Canada a favor by encouraging Canada to seek more profitable trade deals with other nations?

Yes, free trade has been so "good" for Americans that since offshoring of American manufacturing jobs began during the 1970s, real wages haven't increased, relative to inflation, except ever so briefly during President Trump's first term in office.  We're all too stupid to notice that we're somehow getting "poorer" despite all those "cheap" goods pouring in from overseas.  In other news, water is no longer "wet".  Sorry, but all the pathetic attempts at "jedi mind tricks" no longer work on people who have to choose between rent, food, gas money to drive to work, and health care.  Free trade has been great for rich people- just one more way to exploit desperate people elsewhere in ways they're not allowed to here in America due to labor laws.  For everyone else, it's become yet another "utopia".  I can buy lots of things I don't need with money I don't have, but I can't buy an appliance that lasts longer than 5 years.

When you pay your neighbor to make something you truly need (need vs want, endless choice vs meaningful choice), the benefit is that he gets to use the money to support his family, you get your coffee machine, and because his company actually makes coffee machines, any coffee machine manufacturing innovations are likely something that our fellow countrymen get to benefit from, rather than someone living in a foreign land.  I don't need or want an internet connected coffee machine with 50 different settings and more lights than a Christmas tree.  I could care less if it sings to me in the morning.  I'm buying it because I want a hot cup of coffee in the morning.  If it can do that, reliably, for the next 20 years or so, then it was something worth spending my money on.  It wouldn't matter if it costs $100 vs $50 when it lasts 4X longer because it's not loaded with useless features nor made from the cheapest materials imaginable.

The US didn't engage in enough trade immediately before and during the Great Depression for trade to make much difference to the overall US economy.  For starters, most Americans couldn't afford imported goods.  We didn't export much, either.  Productive output nonetheless decreased by half and investments fell to near-zero.

This popular line of argumentation about Smoot-Hawley is frequently used to explain the worsening of the Great Depression, yet there's clear evidence that trade was never a sufficiently large part of the American economy to explain away a 10 year long event.  To this day, trade continues to be a minor part of the overall US economy.  The real issue is that when it comes to making things that truly matter, all the raw materials inputs required to create the machines, we largely stopped doing that.

We're not allowed to open new mines, smelters, and other heavy industry inputs required to make things in quantity.  The Democrats are largely to blame for that.  We tried to open a Lithium mine and they did everything in their power to prevent that.  The same applies to coal, oil, natural gas, and lumber.  I'm willing to meet Democrats half-way on the lumber issue by growing our own bamboo in the Southern US because I don't want all of our tress cut down, either.  You won't find anyone who thinks cutting down all the trees is a good idea, except maybe the green energy advocates who want to clear-cut forests for their wind and solar projects.

Green energy tech is finally dead because they fixated on so much stuff that simply did not matter and imported all of it from overseas.  They could've built solar thermal power plants that operated 24/7/365, and the fact that they used more steel and concrete would've been shrugged off as acceptable because they last for 75 years like any other thermal plant and require the type of maintenance that ordinary people can do.  More electricity without having to burn something is more better, but only when you get it through sustainable means and it's boringly reliable.  That's not what they did, and now there's no more money to finish the over-arching idea, because they're focused on meaningless details while ignoring the important ones.

All that potential, all the money, all that intellectual effort... squandered.  That's the real travesty.

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.

President Trump's tariffs are intended to make it cheaper for American businesses to return manufacturing to America.  All industrialized nations require robust domestic manufacturing bases.  The end goal is to make America self-sufficient so that it does not need to import anything from other countries.  If other nations are upset that America's government is now creating conditions favorable to American businesses manufacturing their products here in America, for consumption by Americans, that's their problem.

Rather than treating this as an opportunity to make themselves self-sufficient as well, other nations like Canada are treating it as a pissing contest with President Trump.  President Biden's administration was busily implementing the same policies for microchips, munitions to feed the war in Ukraine, and other tech items we feel we need.  RobertDyck is upset that Canada will ultimately need to become self-sufficient, even though Canada has the energy, raw materials, education, manufacturing base, and technology to make Canada broadly self-sufficient.  It will cost more money to do, but at the end of the day Canadians will have an independent and self-sufficient nation that is uniquely Canadian and not dependent on remaining in the good graces of any other nation.

Maybe Canadians don't think that will ultimately benefit Canada, but I would like the counter-arguments for why it won't benefit Canada.  I would also like those arguments from the standpoint of long-term generational prosperity- something that goes beyond the "here and now".  A lot of problems have been artificially created by very short-term thinking.

America is now treating our military and economic alliances as, "Let's collaborate if all parties think they're getting fair value for whatever they contribute, or go our separate ways and remain as friends / colleagues if we don't think any particular deal is mutually beneficial."

As an American, I don't want my own fellow Americans to continue to suffer through this systematic "siphoning off of accumulated wealth".  I do not want Canadians to suffer, either.  I have no bone to pick with Canada or Canadians.  That said, my first and most important loyalty and duty is to my fellow Americans.  I cannot control how people in other nations view America's imposition of tariffs on foreign-made products.  Foreign nations have imposed tariffs on American products for many decades now.  If it's good for the goose, then it's good for the gander.

If tariffs are "good / necessary" when Canada applies them to American-made products, then they cannot be "bad / pointless" when America applies them to Canadian-made products, by that very same logic.  Regardless or moral valuation on trade policies, what we're presently doing is slowly but surely draining the wealth out of America.  That process will inevitably end, so whatever benefits it brought to other nations will end as well.

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.

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,

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.

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.

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.

My personal take on Russian armor is that their T-72 is a roughly correct design, but even their frontal armor is ineffective against modern sabot and shaped charge rounds, their fire control system needs modernization, and their thermal imagers used to surveil the battlefield and acquire targets are woefully inadequate.  The 125mm smoothbore cannon is adequate, but their sabot rounds could be improved.  I would prioritize the improved thermal imager and fire control upgrades.

Here's how I would overhaul a NATO tank program:

1. Give up on sabot rounds for direct penetration of modern composite armor.  Kinetic penetrator rounds do work, but the extreme heat involved rapidly "burns out" the barrels through erosion caused by hot high pressure gases, especially the latest and greatest sabot rounds, which renders accuracy unacceptable after far fewer rounds fired.  Barrel life has seen a steep decline with successive "improved penetration" versions of the M829 "silver bullet" sabot round and similar analogs.  M256 L44 cannon barrels are now seeing service lives in the low thousands of rounds at most.  Each M829 now costs as much as a subcompact car.  The Tungsten penetrators fielded by other NATO nations have only made the barrel erosion problem worse, because those require extra velocity to equal the performance of our DU penetrators, which is why you see longer L55 barrels on the latest version of the Leopard 2, prompting Rheinmetall's experimentation with 140mm caliber guns.  Either way, the gun tubes are becoming unwieldy in urban warfare scenarios and if we're increasing caliber to 140mm, then why not use 155mm and call it a day?

2. Standardize on 155mm caliber for all towed artillery pieces, self-propelled guns, tank main guns, naval guns, and airborne gunships.  The US Army discovered long ago, much to its chagrin, that their tanks were nowhere near as well protected against 155mm artillery shells as they thought they were, whether fired at high angle or low angle, whether air burst above the tank or allowed to make contact and detonate.  Air burst 10m to 20m above the target proved to be the most destructive, but direct hits also rendered the target tanks inoperative as tanks.  155mm did not penetrate the frontal armor, but splinters from the shells punched right through the main gun tube of the target tanks, totally wrecked their optics and radio antennas, completely sheared off parts of the running gear (road wheel bogies and tracks), and punched through the engine deck in airbursts, setting fire to the vehicle after destroying the engine beneath the thinner engine deck armor.  Given that the M48 and M60 tanks used in the test were no longer usable as tanks, and would require depot level repairs to return to service, I'd say that still counts as a "kill".  No modern Russian or NATO tank would fare any better when hit with 155mm.  The M776 cannon that equips the M777 towed howitzer is a L39 weapon that weighs 3,700lbs for the barrel and breech, about 6,370lbs in total when configured as an artillery cannon.  The M256 L44 cannon that equips the Abrams weighs 8,330lbs.  The L55 cannon of the Leopard 2 weighs 9,170lbs  I would reduce the M776 cannon's length to L29, same as ye olde M109s, and Charge 4 to reduce the length of the breech and recoil.

3. Cease and desist with loading up every imaginable type of weapon and round of ammo aboard every vehicle.  The sheer quantity of equipment we try to cram into each vehicle is, in point of fact, driving the horrendous cost and weight of our tanks and other armored vehicles.  A tank should have its driver and vehicle commander located inside the hull and nobody inside the turret, which implies an autoloader for the main gun, a high quality thermal imager, panoramic fiber optic cameras that provide a 360 degree view of the battlefield without having the crew constantly expose themselves, a top notch fire control computer, and a battery of grenade launchers for smoke and HE or buckshot for close-in vehicle defense.  It doesn't need multiple machine guns in different calibers, a small caliber automatic cannon, Javelin missile launchers, and partridge in a pear tree.  The more people and weapons you stuff into a tank, the higher its silhouette becomes, which makes it easier to spot from the ground by casual observation, and the heavier yet more vulnerable you make the vehicle and people in it, simply because you cannot adequately armor the top and sides of a tank with a gigantic 3-man turret.

For a tank to support infantry assaults, you need a big gun that fires powerful shells, mostly to deal with light vehicles and dug-in infantry using fragmentation effects.  You can and will fire that $15K M829 "silver bullet" at a Toyota pickup truck if that's all you have, and it will do what it does, but to what end?  Whether or not it can outright "punch through" opposing tanks is largely irrelevant unless you're fighting hordes of tanks over open ground.  Even if you are, that's what Javelins and airborne tank destroyers are for.  What's the point of deploying infantry and combat drones in support of armored ground vehicles if you're not going to use all of their strengths to deliver asymmetric counter-attacks to defeat opposing forces?  The western world has an unhealthy ongoing fascination with the idea of "dueling" and "aces".  War is, "We're using everything we have against everything our enemy has."  It's not, "We're going to defeat your tank with our superior tank in one-on-one combat."  You get "quality" not only from technological superiority, but also from exhaustive realistic training, well-maintained equipment with plentiful spares, local numerical superiority, and the element of surprise.

My concept of "useful combat drones" involves small-ish (~750lbs) remotely-piloted fixed wing piston engine aircraft that can fly at speeds up to 200mph or so for up to 1 hour, and drop a single 155mm glide bomb shell on enemy armor or field fortifications.  If the air defense threat is minimal, perhaps Apaches and Warthogs can fulfill our "airborne tank destroyer" doctrinal role to protect our combined armor / infantry advances.  Apache and Cobra gunships ultimately replaced the M18 and M36 tank destroyers of WWII, as well as armored cavalry scouts for most missions, but now those roles appear ready to assign to small combat drones.  I would retire the Apaches and Warthogs at this point because the threat posed by radar-aimed autocannon and shoulder-fired missiles to piloted close air support assets is far too great against a peer level adversary.  Afghanis may not have any Stingers.  Russia and China, however, have thousands of them.  Flying at treetop level against them is a better than average way to lose good men and machines on every mission.  When, not "if", some of our combat drones get shot down, we build more and our drone operators live to learn from their failures since they weren't aboard the machines that were shot down.

We'll use trucks and cargo ships to transport / fuel / arm / repair our combat drones.  We'll launch them from gun-based catapults that use fuel and compressed air in conjunction with the high-low pressure principle our 40mm grenades use for "soft launch".  We'll recover them with nets so we don't need runways.  This worked well for drones of that approximate size carried aboard American battleships used during Gulf War I to spot the fall of rounds from the 16 inch guns.

The XM-803 (forerunner to the M1 Abrams, equipped with a 152mm rifled gun):

This is what we need to get back to, but without the oversized / overweight 3-man turret, meaning only 2 men in a highly protected hull compartment.  We must use electro-optical sensors to surveil the battlefield instead of exposing tank crews to enemy small arms and artillery fire while they're actively operating their tank.  The entire point of putting them in a tank was to protect them so they could focus on neutralizing targets that are holding up infantry advances.  Keeping all their hatches open, because they're still using their eyeballs to observe, has become a highly lethal anachronism.  This new tank concept deals with enemy field fortifications and infantry by direct-firing 100lb HE shells at them.  When you start direct-firing 155s at people, then everyone gets down, or everyone dies.

The infantry themselves will start manning self-mobile 30mm cannons that the Apaches previously carried, using small arms for close range self defense only.  As they advance, their first objective is to tear apart light vehicles, machine gun nests, mortar pits, and infantry formations using HEI cannon shells.  The Chinese are sure to fight island defense campaigns in the South Pacific roughly the same way the Japanese did during WWII, because there's not many viable alternatives to what the Japanese did, even with modern weapons.

This new way of fighting is firepower-based, rather than a specialized / limited supply weapon for this / that / the other.  Most ground targets can be dealt with using 155s.  Anything that requires a JDAM or bunker buster or cruise missile will be delivered by F-35s, B-21s, B-52s, ships, and submarines.  Missiles will be used sparingly, mostly limited to attacking enemy air defense systems, ships, and aircraft.  There are not enough of these advanced weapons, we cannot easily and rapidly produce more, and huge amounts of resources are captured in their production, stockpiling, and maintenance.  Iron bombs and artillery shells can be manufactured by the millions for comparatively little cost, especially if the shell bodies or bomb casings are stockpiled without filling them with explosives.

We need all services to adopt 155mm.  L29 155s for our tanks to minimize barrel length / weight / recoil force, L39 155s for our towed M777s and M109s, perhaps a few L52 M777ERs for niche applications, and 58 caliber 155s for the US Navy to replace their less effective 5 inch Mk54 guns.  When everyone is using 155, figuring out logistics becomes rather easy.  You need 155 shells and their modular propellant charges.  For longer shots, we'll splurge on Excalibur munitions so we don't burn out the barrels while "firing for effect".  All the services have spent mad money on special application guns.  The US Navy spent "stupid money" trying to make a 155 fly 100 miles.  Each 110lb guided shell cost $1M, as much as a Tomahawk cruise missile that flies 1,500 miles with a 1,000lb warhead.  A 58 caliber cannon can still reach out 40 miles or so with a $50K guided shell.  That would have been a very useful capability for the US Navy to have aboard ships, because it roughly doubles engagement range over the standard 5inch L62 Mk54 gun.  That would put the ship that fired it over the horizon, against most islands.

The objective of this program is to build lots of guns and shells, not "special guns" that require their own "special shells".  Since 155 works for the Army and Marine Corps, it works for the Navy, too.  Heck, we're going to figure out how to mount 155s in Spectre gunships so we can shell people from the air as well.  Against adversaries like Russia and China, more shelling equals "more better".  Russia is not wrong, as it pertains to their liberal use of artillery.  They don't call artillery the "King of Battle" for nothing.  The fact of the matter is that a 100lb HE shell is adequate to kill most military targets, if accurately delivered.  Bunkers, ships, and certain kinds of area targets like runways are the only kinds of targets that require more powerful guided bombs and missiles or torpedoes.