I’m sure a fuselage with the prerequisite hypersonic aerodynamics has already been penned. As is propulsion: Turbo-ramjets…like that found on the SR-71 Blackbird…to accelerate from the runway to Mach 3+, a Scramjet to take over thrust from Mach 3 to Mach 10.
Then what? Something gotta keep the craft speeding up to Mach 20+ for orbital velocity.
I figure this craft would be fueled with Hythane, a cryogenic mixture of Hydrogen and Methane…the precise ratio to be determined.:rolleyes:
As fuel is used up inside the multi-cell fuel tank, it is refilled with air pumped in directly from the high-pressure area inside the jet intake.
Say you have a fuel tank with 12 fluid-tight compartments…as one compartment is exhausted, hot compressed air is piped in, refilling it, and chilled by the freezing cold hythane fuel via heat-exchanger, until it becomes liquified itself, thus taking up less space inside the craft.
As the hypersonic craft transitions the atmosphere (the Karman Line…100 km up), the craft transitions the jets to close off the intakes, feed fuel and liquified air directly into the combustion zone, and the jets now operate as rockets.
I suppose there may be a way to refine liquified air onboard to be almost pure oxygen…but that would be a more advanced mission later on.
If you’re tapping compressed air from within the jet intake, you’re only utilizing what’s already there…no additional drag.
I figure only a mere fraction of the massive volume of air being inhaled by the jets need to be tapped, channeled, liquified, and stored in order to work.
Extracting gases from the jet intake will necessarily cause some drag, because you are interrupting the flow.
The big energy cost in your scheme would be liquefying the air - you need to balance the mass of the compression and liquefaction system against the mass of stored oxygen you would need to take with you otherwise. If and only if the second mass (stored oxygen) is greater than the mass of the compression system might this be worth doing. And without some sort of refinery, the liquified air would be only a fifth as powerful an oxidiser as pure oxygen. so you’d need five times as much.
That’s the air that needs to go through the engine, to be expelled out the back yielding thrust - if you steal it, you’ve harmed your engine’s performance. So you’ll need to enlarge the intake, which will increase drag.
I’m picturing the industrial plant that would be needed to produce large volumes of liquified air in a couple of minutes, then trying to imagine the weight and power burden this implies for a single-stage-to-orbit vehicle. Among your problems would be the need to reject enormous amounts of heat to a rapidly thinning atmosphere.
So, there’s this large mass of air that once was stationary and is now contained in the vehicle, moving at the same speed as the vehicle. And it’s only 21% oxygen. So 79% of this mass you’ve expended energy on accelerating to vehicle velocity does you no good at all. In addition it reduces the efficiency of the fuel you later burn. And you continue to accelerating all this wasteful crap as you burn through it.
This is helpful how?
Rocket design 101: Try to have everything on board be fuel, oxidizer, structural or payload.
Enola Straight - Cool thread. I ran some numbers from a Stoichiometry/Thermodynamics perspective. Here are three different cases :
1> 100% Hydrogen
Lets take 10 gallons of pure liquid Hydrogen at its saturation temperature i.e. -422F. This corresponds to approximately 189 lb-moles. By stoichiometry this will need about 451 lb-moles of air. The temperature of the liquid air will be -309 F and it will take up 31 gallons. So there is space deficit of approximately 21 gallons or twice the volume of Liquid hydrogen.
Also, lets look at Energy - consider 70F to be a baseline temperature. So cooling duty provided by 10 gallons of Hydrogen when it goes from -422F to 70F = 12,368 kJ. And the cooling duty needed by 31 gallons of liquid air when it starts from 70F and goes to -309 F = 39,975 kJ. We have a huge deficit here, more than 3 times the required duty.
2> 100% Methane
Since the liquid temperature of Methane is -267F , it will not be able to liquefy air which needs -309F , so this case is not explored further
3> Hythane (50% moles of Hydrogen and 50% methane, by mass it is about 11% Hydrogen)
Lets take 10 gallons of this mixture at its saturation temperature i.e. -450 F. This corresponds to 188 lb-moles. By stoichiometry it will need about 1120 lb-moles of air. The temperature of the liquid air will be -309 F and it will take up 76 gallons. So you have a volume deficit of about 7 times the volume of liquid Hythane.
Also, lets look at Energy - consider 70F to be a baseline temperature. So cooling duty provided by 10 gallons of Hythane when it goes from -450F to 70F = 18,430 kJ. And the cooling duty needed by 76 gallons of liquid air when it starts from 70F and goes to -309 F = 99,275 kJ. Once again, we have a big deficit of around 4 times the available duty.
Also, in my experience cryogenic heat exchangers are finicky and take a long time (hours) to reach steady state. You can certainly overcome these problems by adding compression power to the streams or other methods.
Everyday Astronaut explained the difficulties facing SSTO whether air-breathing or pure rocket: https://youtu.be/Sfc2Jg1gkKA
The ultimate purpose of a hypersonic air-breathing SSTO is getting payload to orbit reliably at a lower cost than competing methods. But as SpaceX has demonstrated, it is possible to achieve great cost savings simply using reusable rockets. The economics are discussed in this science blog: Fully reusable Spacex Rockets would be lower cost than Skylon spaceplanes | NextBigFuture.com
Airbreathing SSTOs may be like a cool idea, but you don’t get extra points for getting to orbit “the hard way”, or getting to orbit the “coolest way”. The goal is getting useful payload to orbit the simplest, safest and cheapest way. So far no SSTO of any type (air-breathing or pure rocket) has ever achieved earth orbit, even as a sub-scale prototype with zero payload. Due to thermal, complexity and structural mass issues, getting payload to orbit via single-stage airbreathing propulsion is like trying to run a four-minute mile while wearing 10-pound ankle weights.
As a thought experiment so show how difficult this is, there is little difference between a single-stage hybrid air-breathing/rocket SSTO vs two-stage reusable vehicle with an air-breathing hypersonic first stage and a rocket-powered second stage. In some ways that is even easier since the dead weight of the air-breathing apparatus need not be lugged to orbit and each stage can be optimized for the particular regime. Yet even this has never been done.
We need to buy Venezuela while it’s cheap. We can build a space port well over 10,000 feet altitude nearly on the equator. Launching any SSTO from there will reduce the fuel requirements.
I think some kind of rail gun like launcher would be preferred for take off. They should be able to accelerate a space plane to several times the speed of sound without the plane carrying the weight of the fuel for takeoff.
Apparently, my idea had been anticipated as far back as 1957.
It occurred to me, if oxygen was NOT loaded up on the runway, but instead, scooped up, liquified, and stored onboard in space previously containing fuel, you would not have to make additional allowances for volume and mass for the oxidizer.
Gather up oxygen in the minute or so transversing the atmosphere, then using it in the vaccuum of space.
The Aerospace Plane plans were shelved during the Mercury astronaut program…perhaps with todays technology, it could work?
From your Wikipedia link : “ . It will be seen that heat-exchanger limitations always cause this system to run with a hydro-gen/air ratio much richer than stoichiometric with a consequent penalty in performance[1] and thus some hydrogen is dumped overboard.“
This is consistent with my calcs above. You’ll be wasting a lot of hydrogen which will be used just as coolant and would not have enough oxygen to combust. I think the author is taking scientific liberty when saying “some hydrogen is dumped overboard”.
Ordinary hydrogen rocket engines run much richer than stochiometric to start with–the Space Shuttle Main Engines ran at 6:1 mass ratio, instead of the expected 8:1 (oxidizer:fuel). That probably could have increased further, though maybe not to the 2.7:1 that’s required assuming your numbers are correct.
That said, I think the situation is much worse than you indicate. The starting temperature isn’t 70 F; it’s 1800 F due to the ram compression of the gases (according to the page on the SABRE engine)
