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GW,
That is why I set the dV requirement higher than the bare minimum required to attain orbit. My presumption is that the RCS (GOX/RP1) and OMS (LOX/RP1) are integrated systems that burn residual propellants. Mission duration is 24 hours or less. If passenger transfer to an interplanetary colonization ship can't be accomplished inside of 24 hours, then you de-orbit and land. 1.5% of 2,000t of propellant is 30t. That's plenty of propellant for orbit circularization, on-orbit maneuvering, and de-orbit burn. The Space Shuttle had 14,297kg of NTO/MMH propellant in total for its OMS and RCS systems. My proposed SSTO vehicle has more than 2X as much residual propellant in the feed lines to the main engines, which is where it will be stored until it's fed into the OMS and RCS. The primary power subsystem will consume around 750kg to 1,000kg of propellants over 24 hours to generate electrical power to provide life support and power avionics as well as flight control surfaces and vehicle subsystems. That's a reasonable propellant reserve- more than the Space Shuttle had, and since flights are regularly scheduled, if it proves necessary a rescue vehicle will be launched the next day to offload the passengers and either take them to their ship or take them back to Earth.
Edit:
The dV increment I factored in included enough propellant for to achieve orbit while covering anticipated / prototypical gravity and drag losses, on-orbit maneuvering, de-orbit burn, and running an APU to provide electrical power and propellant pressurization during the entire flight. Airliners don't use Hydrazines. They burn O2 and kerosene. Since this vehicle is an orbital class airliner, if it needs power, then it gets that power from the same propellant sources that feed the main engines. Containing dry vehicle mass growth is critical, so eliminating separate subsystems and multiple propellant types is a form of vehicle mass control. LOX or GOX and RP1 requires spark plugs, but modern Iridium-tipped spark plugs are pretty reliable. Whether that is seen as a major risk or complication, I don't know, but the RS-25s had Iridium-tipped spark plugs to assure ignition within the pre-burners, and they were fired continuously all the way to orbit.
Since this vehicle doesn't have any Hydrazines onboard, all the special handling equipment, training, and tasks related to loading Hydrazine type propellants can be eliminated. Separate Helium tanks for propellant pressurization can also be eliminated. We must simplify the entire vehicle design to the degree possible, use the latest and greatest materials and technologies for the performance benefits provided, and refrain from stubbornly fixating on how something was done in the past, when it was self-evident that made SSTO infeasible.
Last edited by kbd512 (2024-06-15 22:01:57)
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I'm curious if we at least agree that total impulse per unit volume of propellant is a figure of merit for SSTO, rather than propellant weight alone. Propellant weight absolutely does drive vehicle dry mass fraction, but not to the same degree as LH2 tank internal pressurization, which is double that of LOX or RP1.
Total impulse per unit volume of propellant results in at least 2X greater structural mass allocation for a LH2-fueled vehicle, and more TPS mass allocation, albeit not 2X more due to cube-square, maybe 1.2X to 1.4X more for a lifting body design. As you scale-up, that problem only becomes worse, not better, because the material doesn't magically become any stronger or stiffer as volume and surface area increases. Cube-square, applicable to all solutions, ameliorates the surface area and TPS mass problem, but at least 50% greater dry vehicle mass for LH2 is unavoidable due to LH2's extremely low density per unit volume and high internal pressurization to force-feed fuel into the feed lines / pump inlets for the engines.
Even LH2's 100s greater Isp, relative to LOX/RP1, is not enough to overcome that problem. The low thrust per unit engine weight of LH2-fueled engines only aggravates the total impulse problem, because it results in all major vehicle components becoming heavier - propellant tanks, engines, TPS, and landing gear. LH2 is a disabling, rather than enabling, propellant selection for SSTO.
Run a straightforward total impulse calc to demonstrate the total impulse provided by 1m^3 of LOX/RP1 and LOX/LH2 at avg Isp for staged combustion engines (338s for LOX/RP1, 409s for LOX/LH2). I even tried the calc using 452s as the Isp for LOX/LH2, but the end result is that the LH2 fueled vehicle STILL has a greater dry mass fraction than LOX/RP1 with avg Isp.
After that's done, then factor in 200:1 TWR for LOX/RP1 engines vs 77:1 or even 100:1 (imaginary engine tech) for LOX/LH2 engines.
Is it at least obvious that pressurizing LH2 to 45psi represents a far more significant mechanical load on the propellant tank structure than a 5g acceleration with a full LOX or RP1 tank of equivalent volume?
By the time you can achieve 5g in the LOX/RP1 powered rocket, at least half of your propellant has already been expended. At all times, you're feeding LOX/RP1 to the engines at 20psi to 22psi, rather than 45psi for LH2.
If your vehicle's propellant tanks are designed to withstand 45psi LH2 pressurization and 3g acceleration loads, then you're de-facto able to use those same tank designs to withstand higher acceleration loads from heavier LOX or RP1 propellants, at half the internal pressurization, after half of the propellant has been consumed during ascent. Your fuel tank is not subjected to LH2 permeation and embrittlement, and the cold LOX tank is still much warmer than the LH2 tank.
The problems LOX/LCH4 poses for a SSTO are less severe than those imposed by LOX/LH2, yet they're still greater than LOX/RP1, with the end result being that you push less payload into orbit than you do with LOX/RP1 for a given vehicle dry mass. The only way you net an actual benefit for SSTO vehicle dry mass is if LOX/LCH4 is the same density as LOX/RP1, same engine TWR, and you get the Isp improvement of CH4 over RP1. Then and only then do you push more payload into orbit for a given vehicle dry mass. CH4 gives us a modest bump in Isp, the same or slightly worse TWR for CH4-fueled engines, and significantly greater vehicle volume and thus dry mass.
With a 10km/s dV increment as a hard requirement, a practical SSTO becomes a highly constraining problem with an equally constrained practical solution set. If your propellant combo doesn't have a density near that of water, or your engines cannot provide 175:1+ TWR, or your propellant tank material isn't 400ksi+ with stiffness similar to steel and weight similar to Magnesium, then you don't have a practical SSTO, because all your payload mass gets consumed by the vehicle itself. All of those performance characteristics must be combined into a single vehicle design.
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Another thing to consider in regard to tank inert mass: is it, or is it not, also the airframe of the vehicle? If it is, what additional loads besides endways gees and internal pressurization, must it carry? If it is not, its mass is in addition to the airframe mass that surrounds it.
I'm only guessing that wall thickness depends upon internal pressure level and material strength as a rough first approximation, and the volume of material to be weighed is the tank surface area times that thickness times some factor a tad above 1, that accounts for local thickness increases, such as at the head-cylinder joints. That takes care of low-density hydrogen, vs the other fuels.
That simple estimate ignores other very real-world mass additions for things like internal surge baffles, and the plumbing that takes propellant from another tank through this one, to the engines. Sometimes that plumbing must be quite heavily insulated.
As for resisting other airframe loads such as bending, and external pressure on one side but not the other (acting to cause collapse laterally), that requires either external or internal stringers, and internal frames, to act similar to a monocoque fuselage.
And I still do not believe in organic matrix polymers that aren't structural junk at 300+F. Composite tanks will inherently require more heat protection that stainless steel, in ANY entry scenario, even the low velocity entries of 1st stages. 300 F is the total temperature of the air at only Mach 2.2 in the stratosphere. That's only 0.64 km/s velocity through the cold stratospheric air.
Aluminum goes slightly hotter than that at about Mach 2.5-2.6.
GW
Last edited by GW Johnson (2024-06-16 14:12:56)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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P2SI™ 900HT composite prepreg system exhibits a glass transition temperature of 900°F (482°C) and excellent mechanical strength retention above 400°C. Void free composites are easily fabricated into structural components using a variety of reinforcements, including glass, quartz, and carbon. This paper presents the latest processing, performance, and durability characterization of this recently developed composite system. In addition, results on the use of P2SI™ 900HT for high temperature tooling applications, including 371°C autoclave cure of polyimide composites is presented. To date, no other polymer matrix composite has been successfully used as a tooling material in 371°C autoclave processes.
Resins for the Hot Zone, Part I: Polyimides
This article speaks to composites withstanding sustained temperatures of 600F to 1,000F, used in military aerospace applications. Belief is not required, because the high temperature performance is still there, regardless of knowledge or belief. The F-35, in particular, has composite components in its airframe that must withstand temperatures of up to 1,000F in routine service, particularly those near the engine and its exhaust nozzle. The complaints from the US Air Force were that even that level of capability was not enough, so we developed resins with ever-higher service temperature limits. Note that this article is dated from 2009 and it states, even 5 years ago, these materials were still in the incubation stage of development".
Materials tech has moved on, even if the thinking of the older generations of aerospace engineers who used metals has not.
All the materials tech I spoke of is quite real, it's actively used by our military or space contractors in real aircraft or rockets, and it does, in point of fact, withstand the temperature range it's subjected to, else we'd have aircraft literally falling out of the sky following catastrophic structural failures. Since that's not happening, the materials tech is on point.
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That kind of stuff is quite expensive. To my knowledge, whole airplanes are not made with it.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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GW,
The F-22, F-35, and B-21 are, even though parts of their airframes don't require it. The weapons bays on the F-22 and F-35 became so hot that the weapons themselves, or at least their electronics, were failing. The airframes were fine, but the electronics failed.
Now that I've thought about that, messing up the electronics, even if the airframe can tolerate the heat without any real issues, might prove far more problematic. That would obviously be very bad since this vehicle has no windows and you're trying to land it using instruments only. We might have to rethink protection levels in certain areas to prevent the electronics from getting fried. We're relying on DAS and Ku-band radar to fly all-weather precision approaches. In the F-35, they used fuel as a thermal sink to cool the electronics. We'll need something similar.
Ever wonder why we're burning out F-135 hot sections in our F-35s within 3,000hrs?
They're demanding so much bleed air to cool the electronics that their hot sections are getting BBQ'd for lack of cooling air flow over the blades. The technical solution has been to make the hot sections even more tolerant of high temperatures, rather than reducing the cooling air demand for the electronics. More modern electronics with the F-35's latest tech refresh haven't helped. They produce less waste heat, but now there's simply more total computing power crammed into the same airframe and no space or bleed air reserve for more cooling capacity, because the Air Force changed F-35 compute requirements after the engine design was finalized and tested. The airframes themselves are built to readily withstand the heat coming from the engine(s). The electronics? That's another story. That's why the F-22s and F-35s open the weapons bays while on the ground after they start the engine(s).
Incidentally, prices for high temperature resins have been improving. They're now used in exhaust systems for sports cars and motorcycles. It's still very pricey, I'll grant you that. An enormous amount of development work has been done on high temperature resins, though. We've had these resins for at least 15 years now, and they do work. Whatever the manufacturer states, I still subtract at least 150F off of whatever they claim in the way of temperature tolerance, even if there is ample test data proving that it works. I think F-22, F-35, and B-21 count as "ample test data". The airframes of those jets would all fail catastrophically if they couldn't take the heat.
We're using expensive fiber, expensive resin, and expensive TPS materials. We're paying up front for a high quality airframe, rather than paying later with lives lost. I expect the bill of materials for the fiber and resin to amount to about $10M. All told, we're looking at around 27,000kg worth of specialty materials, primarily fiber and resin with small bits of very costly steel. This includes about 16,200kg, at $250/kg, which is what T1100 fiber tow / roving costs in bulk, so $4.05M for the fiber. I expect the resin to cost at least as much per kg. The TPS materials will cost around $1,000/kg. In terms of airframe cost per kg, it's very steep, but you have to remember that it doesn't weigh very much and a lot of the weight is in the engines, which don't cost nearly as much to produce. I would estimate a $75M airframe cost (materials, energy, and labor), which is quite a lot ($1,250/kg)- easily more than the F-22 or F-35 on a per-weight basis, but over 1,000 flights it's only $75,000/kg, and per-passenger-flight cost only amounts to $150 of the ticket price. Almost all of the real cost comes in the form of propellants / personnel / facilities, just like a normal airline service.
If you were to show the airframe bill of materials to a manager at Lockheed-Martin, they would choke on the cost. Northrop-Grumman might find it more reasonable if we're talking about B-21 airframe cost, but even those airframes are far less costly per unit weight, if the electronics and RAM specialty materials were excluded. That's why LM put so many Titanium forgings in the F-22. In bulk, it was far less expensive than CFRP, and they were pretty fanatical about keeping the actual airframe cost within the realm of sanity. The compromise with the Air Force was making every 4th spar from Titanium because it was less than half the cost of CFRP. The end result was absurd wing spar inspection and maintenance costs to monitor cracks in the Titanium, but hey, the government saved a little bit of money up front on the airframe itself. There's a nice presentation on this special brand of silliness.
I figure that we're charging about $150 per ticket sold to pay off the airframe, $6,000 to $7,000 for the propellants, plus some additional cost for facilities, permits / paperwork / legal expenses, and everything else. If our per-ticket cost amounts to $10,000 per passenger, then this is easily the cheapest ride to orbit by a lot. For a family of 4, we need to figure out how to get the total cost down to about $100,000 to move mom, dad, and the kids to Mars. Doing that will require using multiple complementary launch technologies with exceptionally high reliability and launch rates, low marginal / per-flight costs, and streamlined bureaucratic nonsense related to FAA type certification. That last part will be the real "trick" to making all of this work.
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