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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?
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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.
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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)
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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.
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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.
"Plan and prepare for every possibility, and you will never act. It is nobler to have courage as we stumble into half the things we fear than to analyse every possible obstacle and begin nothing. Great things are achieved by embracing great dangers."
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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.
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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)
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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.
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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.
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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)
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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.
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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.
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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)
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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.
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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)
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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.
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