For the small rockets Armadillo and Masten have flown, I don't think turbomachinery would be an option.
The difference would be in the materials used to accommodate for the chemical makeup and temperature of the gas.
A steam turbine shouldn't inherently be different than any other kind of turbine for a given size and pressure ratio. The difference would be in the materials used to accommodate for the chemical makeup and temperature of the gas.
Yeah, that's what I meant. You also don't need to design a gas generator. Not show stoppers, I just imagine it make things easier while you're still learning how to do it.
Note that the water used on steam plants is not straight from the tap water. For this application you'd probably want to go with deionized water with no scale forming impurities in it running in a sealed system. Note that the work of the Whitehead group at LLNL indicates turbines scale down badly below about 5000 lb thrust. The turbines are small, which means the clearances are very tight and the surface finish has to be very good, as the boundary layer will be proportionately thinner.OTOH a positive displacement reciprocating pump using say corrugated pistons would be easier to fabricate, eliminate rotating seals and bearings and deliver complete separation of propellant from driving fluid. Also the pump drive fluid is relatively inert (super heated steam should be treated with considerable respect. ).The problem with all separate pump driver fluid concepts is now you need two heat exchangers, the CC wall and whatever you're dumping the waste head the was not used in the pump drives. This might not be the problem people think it is if the engine stays in the atmosphere. Different goals, different trades.In vacuum you're looking at either radiators (which get big at low temperature differences, even with a carbon fibre "carpet" around them to enhance emission) or dumping it to the propellant flow IE a 2nd HX.Historically people have said "No, weight penalty not worth the flexibility of being able to select optimum drive fluid properties"OTOH amateur developers, or people wanting lower maintenance reusability might prioritize things differently.
I still have to ask why? Your list of cons is missing many items.
1. If there was a single type of engine that is better than every other one in every way, then why are there so many different ones with various different cycles currently in use? It would seem that by your logic, there's only one metric of importance.I didn't think I needed to state some of the more obvious pros, but I will:0) No secondary combustion chambers1) Closed cycle2) Higher chamber pressure than a standard expander3) Lower temps and higher pressure drops across turbines, and possibly higher chamber pressures compared to the various staged combustion cycles (not to mention less corrosive than the hot oxidizer rich ones) while being no more complex, and likely simpler. Those sound like some decent pros to me.
1. All the goods ones are in use or have been used. This one is on sideline because it isn't good.0) why is that good?The complexity, extra mass, and losses in the heat exchangers are greater drawbacks. Not to mention the inability to start such an engine.
In early 2006, XCOR and ATK GASL received a DARPA contract to "investigate, develop, and demonstrate a novel configuration for a liquid rocket engine, namely a Third Fluid Cooled (TFC) liquid rocket engine," for which we used the Tea Cart to generate superheated steam suitable for driving a Rankine cycle. This steam cycle will allow a turbopump system to develop high chamber pressure in a more efficient way than staged combustion cycles, thus improving the performance and durability for boost propulsion, as well as orbital transfer applications.
The RL10 starts just fine.
DARPA contracts are not relevant indicators for viability of technology. More like throwing stuff at a wall and see what sticks.
Quote from: msat on 07/16/2015 10:44 pm The RL10 starts just fine.It just needs valves opened to start.
Existing engines don't have a closed system driving the turbo pump so back pressure is not a concern
For most propellants, certain levels of local combustion gas velocity (or mass flux) flowing parallel to the burning surface leads to an increased burning rate. This "augmentation" of burn rate is referred to as erosive burning, with the extent varying with propellant type and chamber pressure. For many propellants, a threshold flow velocity exists. Below this flow level, either no augmentation occurs, or a decrease in burn rate is experienced (negative erosive burning).The effects of erosive burning can be minimized by designing the motor with a sufficiently large port-to-throat area ratio (Aport/At). The port area is the cross-section area of the flow channel in a motor. For a hollow-cylindrical grain, this is the cross-section area of the core. As a rule of thumb, the ratio should be a minimum of 2 for a grain L/D ratio of 6. A greater Aport/At ratio should be used for grains with larger L/D ratios.
7.07 inches squared / 3.14 inches squared = 2.25. Now here comes my question (assuming I've just got everything right).
Going by what I emphasized in the above quote, the grain L/D ratio should be no more than 6. I assume that by "L/D" ratio, the author means grain length to diameter ratio? Does this mean the outside grain diameter? So would the limit on the length of my solid rocket engine be around: 6 inch diameter * 6 = 36 inches?
Quote from: ClaytonBirchenough on 08/06/2015 12:13 am7.07 inches squared / 3.14 inches squared = 2.25. Now here comes my question (assuming I've just got everything right).All good so far but note that you can simplify the math by just squaring the ratio of diameters (or radii, does not matter which). Area is proportional to square of length, that's enough knowledge so no need to calculate the actual areas.(3/2)2 = 2.25Quote Going by what I emphasized in the above quote, the grain L/D ratio should be no more than 6. I assume that by "L/D" ratio, the author means grain length to diameter ratio? Does this mean the outside grain diameter? So would the limit on the length of my solid rocket engine be around: 6 inch diameter * 6 = 36 inches?It's the grain L/D, L is grain length and D is outside diameter. I don't know how the thumb rule 2 Ap/At ratio scales with bigger L/Ds.The combustion area in your simple cylinderical core desing would double at the end of the burn, thrust also. Eventually you want to seek out more complex port geometries for more even thrust.Have you read Richard Nakka's pages? A lot of information for solid rocket hobbyists there.http://www.nakka-rocketry.net/th_grain.html
Please correct me if I'm wrong:Was MMS the last usage of the RL-10A-4-2 Centaur engine, or do they have specific manifested flights for the old engine?
Quote from: longdrivechampion102 on 11/21/2015 11:36 pmPlease correct me if I'm wrong:Was MMS the last usage of the RL-10A-4-2 Centaur engine, or do they have specific manifested flights for the old engine?OSIRIS-REX and possibly all Dual Engine Centaurs.