sevenperforce

Coward!

Sure thing. I'll detail them out in the dedicated thread.

Yep. Could be higher, actually. I need to re-run my optimization algorithm with the correct vacuum specific impulse at various H:O ratios instead of SL specific impulse. But it's going to be hella high.

Well, I realized I was using SL ISP for H2/LOX rather than vacuum ISP. The actual ISP of the dual-mono engine is going to be closer to 440 s. Advantages: higher T/W ratio; lower tankage volume and weight. Dramatically.

What if you go even smaller? No payload at all, minimal avionics, smallest possible terminal stage.

A cursory examination of rocket engineering rapidly identifies LH2/LOX rocket engines as the gold standard for efficiency. It's hard to beat 450+ seconds of vacuum specific impulse, even though hydrolox rockets don't always have the greatest T/W ratios and require very large fuel tanks which eat into mass fraction rapidly. I think I can beat that, though. Hydrogen peroxide is not a very good monopropellant. It has only 161 s of impulse. Hydrazine is a bit better, as far as monopropellants go; it boasts upwards of 220 s. And together, they're not much improved; a hydrazine/peroxide bipropellant rocket can't even break 300 s in a vacuum. Put them together in the right way, though, and I think I might be on to something. This is a fairly basic, no-nonsense linear aerospike engine. There's just a single difference: instead of using small bipropellant combustion chambers, it uses staggered monopropellant combustion chambers, alternating between hydrazine and high-test peroxide. The peroxide thrusters and hydrazine thrusters produce flows which, after beginning to expand against the aerospike, are already moving very fast. High-test peroxide's 161 seconds of specific impulse corresponds to an exhaust velocity of around 1.58 km/s while hydrazine's 220 seconds corresponds to an exhaust velocity of around 2.16 km/s. At a molar mass ratio of 3:2 (peroxide:hydrazine), the mutual flow is traveling down the aerospike at an average velocity of 1.83 km/s. But it doesn't stay that way. As the two compressed supersonic flows mix, they ignite with each other: Decomposed hydrazine contains 4 grams of diatomic hydrogen per mole; decomposed peroxide contains 16 grams of diatomic oxygen per mole. At the previously-mentioned 3:2 molar ratio, the oxygen and hydrogen will burn with a vacuum specific impulse of 430-450 seconds. Of course, the reactants compose only one third of the mass of the flow, so the net increase in propellant flow speed will be about 2.49 km/s. However, because that increase takes place in a flow which is already moving at 1.83 km/s, the speeds stack. This staged combustion results in a total exhaust velocity of 4.32 km/s, for a specific impulse of 441 seconds. Because monopropellant thrusters are being used, the thrust-to-weight ratio will be fantastic, a major advantage over other linear aerospike designs. Moreover, both fuels are dense and liquid at room temperature, allowing small tank volume and a smaller launch vehicle.

Huh. One of the hydrazine thrusters I saw for sale online (used in various spacecraft) weighed in at 650 grams but only had a peak thrust of 24.6 N.

This is more of a simple question than a topic, per se, but what's a ballpark T/W ratio for a monopropellant hydrazine thruster? How about for a monopropellant hydrogen peroxide thruster? I can see the T/W ratio for something like the Merlin 1D quoted online but comparable figures for monoprop thrusters aren't forthcoming.

The Imperial system was based around powers of 2 and 3 rather than powers of 10 in order to streamline mental arithmetic and mental fractions. There are 8 fluid ounces in a cup and 6 teaspoons in a fluid ounce...which means if you're changing the proportions of a recipe, you can subdivide to pretty much any fraction you want without needing to use a calculator. And yes, it's wonky to have exactly 5280 feet in a mile, but if you're surveying land without fancy laser rangefinders then it's nice to be able to split a mile into 2, 3, 4, 5, 6, 8, 10, 11, 12, 15, 16, 20, 22, 24, 30, 32, 33, 40, 44, 48, 55, 60, 66, 80, or 88 equal segments without having to get into fractions of feet.

As you scale down rockets, the atmosphere becomes "thicker" from the rocket's point of view, to the point that it would eventually become impossible to achieve orbit. If some hobbyists got together and pooled their resources, what's the smallest possible rocket that could actually reach orbit? Staging allowed; no payload required other than aeroshell and engines.

Obligatory....

Hush, you're ruining my pet theory with science

The Falcon 9 first stage can do SSTO with about 400 m/s to spare, if I recall correctly. Other than the aerospike engine, what advantage does this have over a scaled Falcon 9 stage 1?

Yeah, ascent is the relatively easy part; tanks lend themselves well to handling axial stress. Controlled separation shouldn't be a problem either. Anaxial stress on re-entry is going to be a poodle. I confess I'm still not sure what your form factor and details are going to look like. Are you thinking of an SSTO that can function as an asparagus strap-on booster and achieves a higher dV than the Falcon 9 first stage merely because it can survive re-entry?

Ooh, this is a fun one. Let's see here. Black holes. Stellar collapse causes an exponential density gradient once neutron degeneracy pressure is exceeded. As a result, all black holes start at about 3/4 of the Planck mass and exist as a gravitationally bound photon pair orbiting at 3/2 of the Planck length. They tunnel out and collapse on the order of the Planck time, but return to a bound state as soon as they encounter another photon. Black holes are thus a gravitationally-bound photon gas cloud with a radius just greater than their own Schwarzschild radius, and all their behavior (Hawking radiation, etc) can be characterized using this quantization metric. Dark matter. Black hole quanta have a very low but nonzero probability of escaping intact rather than decaying. Those which escape intact are metastable and persist in empty space as dark matter; their spin axis causes preferential sorting and a thin disk halo within galaxies. This thin disk destabilizes the Oort Cloud whenever we pass through the galactic plane, resulting in a higher incidence of cometary impacts and periodic extinction events. Venus. There is no life on Venus anymore. The microbial life that used to exist on Venus created a runaway greenhouse effect and suffocated itself. Moon landing. The United States was dead-set on faking the moon landing until they realized it would be easier to actually land on the moon than it would be to successfully fake it.

It's fine to start with a target performance spec and build a vehicle around it. You just have to justify that you can fit the components of your vehicle into those specs. The onus is on the designer to show that. In this case, I'd be suspicious of structural integrity and re-entry handling. Trying to put those tanks into a blended lifting body requires that the structure handle highly asymmetric stresses where it would typically only take axial stresses. Weight cost of that additional structural support plus heat resistant shielding is really going to crowd those specs. But using target specs as a starting point is entirely valid as long as you're willing to take those things into the balance. If anything, that's a more useful approach than just trying to design a rocket piece by piece, because it tells you much more readily whether you'll be able to get the performance you need.

We could move Vesta wherever we wanted it. Makes for a promising low-gravity world. Titan is in a neat place simply due to the high access to raw materials just off-world. Also, geothermal energy galore. Shouldn't be hard to terraform. Of the ones you listed, Mars is good if you can make it profitable. But there's a whole thread about that.

Peroxide and hydrazine pumped into a single combustion chamber have a crappy ISP, true. But staged/induced supersonic shockwave combustion carries a lot more promise; I did a little iterative optimization by hand and got a vacuum ISP of 414s at a 2:3 molar ratio (N2H4/H2O2). Even if the supersonic combustion mode can't manage efficiencies quite as high as, say, the SSMEs, the exhaust velocity is still going to be on the order of 3700-4000 m/s, with T/W comparable to the SRBs and extraordinarily compact, lightweight tankage. Plus, monoprop fuels make pre-preburning turbopumps super simple. You'd probably want an open aerospike of some kind...

404 seconds is the vacuum specific impulse of open-staged hydrazine/peroxide at a 1:1 molar ratio. I'll have to run a numerical analysis to see how far that can be improved by varying the ratio. Sea level specific impulse would be dramatically greater, and T/W is also higher than conventional rockets. Performance that outstrips hydrolox while using dense, liquid-at-room-temp fuels is...promising.

The weight cost of making something like the Falcon 9 stage 1 into a lifting body would probably be prohibitive. I'm surprised no one reacted to the projection of 404s peak impulse for a hydrazine-peroxide dual-staged-combustion rocket.

Hydrogen peroxide and hydrazine in an open staged-combustion rocket would have a peak vaccuum specific impulse of 404s and a T/W ratio that makes a Merlin 1D look like a turbofan. And that's not even with any sort of optimization analysis; just straight stochiometric. Ah, to get back from higher/faster than the Falcon 9 first stages can manage.

Hybrid-fueled, maybe. Solid-fueled motors can't restarted or throttled. And a hypergolic liquid-fueled option has much better Isp than a hybrid rocket. Certainly the former. Though N2O4 isn't a monoprop...at least, not as far as I know. I was thinking of something more like H2O2 and hydrazine. Rockets are Carnot engines; they convert heat to velocity, with diminishing returns. If, however, two monopropellants were fired into each other at already-supersonic flow speed, their remaining chemical potential could be utilized at much higher efficiency. Basically a two-stage combustion process...an internally-burning scramjet. This is basically what I had as my first rendered concept earlier in this thread.

Oh yeah. Of course, it's not really a lifting body at all. There's a bit of induced hypersonic lift but nothing subsonic because it's vertically landed. Yeah, for crew return, it's hard to imagine anything simpler, safer, or more reliable than a propulsively-landed capsule a la Dragon V2. Only tangentially related...but had anyone ever proposed pairing two chemically reactive monopropellants for an SSTO design?

The horizontal-attitude vertical takeoff is just a pipe dream, but the landing is pretty necessary. I don't like wings. Do ye not with wings what ye can accomplish with a short retro burn, the good book saith. Cargo is probably not necessary either. On the other hand, an air-augmented SERV design would quite possibly be the smallest crew transfer shuttle we could reasonably build.

Madness?! This is SPARTA! *Single-Propulsion Atmospheric Ram Terminal Ascender