I thought there were some existing parametrics for the savings in mass fraction for a stage or a lander launched dry (versus wet), but apparently not. Is there anyone out there who whomp up a simple system in AutoCAD or Solidworks? Say 50,000 lbs of LOX/Hydrogen, launch acceleration 6 gees?
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What exactly is the question?
The question is, if the space vehicle (lander or insertion stage) is launched empty, does it reduce the structural requirements, and therefore dry weight?
So if design a lunar lander that intended for say max gee load of a fully fueled lander of 1 gee, can such lander be launched empty from the Earth launch, if max gees of Earth launch is expected to reach 6 gees?
Another aspect is could such lander could be fullly fueled and get earth launch which has 6 gee, and it could withstand such loads because it’s supported while being launched?
It depends,
Quick back of the napkin calculation says yes, especially if you pressurize the “empty” tank with helium or some other gas. But a “Real” answer would require more details about the theoretical spacecraft and launch vehicle. Relative CoG, number of Gs being pulled, aerodynamic stresses at max Q, that sort of thing.
This may be a silly question, but why would you want to launch it dry? Wouldn’t it make more sense to launch fueled, burn to provide the “upper stage”, then arrive on orbit dry?
I want to launch it dry because it allows me to launch it on a much smaller (and perhaps more reliable) launch system. It also may reduce the structural mass, which will have benefits for on-orbit performance, particularly if it’s used for multiple missions.
I’m not qualified to provide an answer, but I’ll take a stab at the route to the answer;
We’re talking a stage delivered to orbit dry; what sort of reduction in dry mass is possible vs. a stage that has to support both itself and its fuel under launch accelerations. Is that right?
IMHO, if the launch accel when launched dry is 6G, and the operating conditions the stage will have to handle during fueled use is 6g, the mass savings will be zero, because the structural margins have to be able to handle the max subjected loads.
If, on the other hand, the launch accel (dry stage) is 6g, while the fueled use is 2G, you’ve got potential mass savings in the tank and load carrying structures. (because launched wet, the stage would need to support the fuel mass plus its own).
However… a complicating factor is structural loading from the payload mass. If it was to be launched wet but sans payload, vs. dry with payload, you’ve got the issue of dealing with the payload mass at 6G instead of the fuel mass.
My guess is we’re talking about potential dry mass savings in departure stages enabled by orbital fuel depots. I think they’d be significant in some cases, but it’d depend on mission specifics. My guess is the vast range of variables this imparts are why there’s no standard set of parametrics.
No Solid Works or FEA but the back of my envelope says the savings will be minimal. Thin walled cylinders are pretty efficient in terms of the compressive load from acceleration, especially if internal pressure is enough to prevent buckling. Bending stress would be dominated by the mass of the payload at the pointy end. The 45,000 lbs of O2, volume about 20 cubic meters, would be supported at the base by the bulkhead above the engines and that might be lightened. the 4,500 lbs of H2 would occupy 32 cubic meters above and contribute little to bending moment.
Dynamic loading from things like shock and vibration might be increased by running the tanks with just pressurizing gas, my envelope isn’t that big.
Maybe a few percent of the mass of the tanks?
Using a denser fuel would likely have a bigger effect from reduction of the size/length of the tank and concomitant reduction in bending moment.
You’ve put your finger on the issue. There is no payload on the pointy end. The stage itself is the payload.
How close to minimum gauge is the tank? The biggest load savings I see launching dry is a reduction in pressure at the bottom of the tank.
What about the structures that support the tanks? Those, too, could be reduced for dry tanks. Likewise, the mounting hardware that secures the stage to the LV.
However… IMHO, it would help greatly if we had an idea of launch G vs. lost launch mission G. If mission G peaks at around what launch G does, IMHO launching dry saves nothing, because the structural demands on the stage would be the same. (though some mass savings via the mounting coupling to the LV would still occur)
Rand,
I don’t see how a bare tank makes a lander unless it’s supported by a surrounding structure, like the Shuttle. Spherical tanks like the LEM are a different case and the supporting structure could be lighter. A lot would depend on specific details.
An empty cylindrical tank for in-orbit connection to engines and payload would still have to withstand launch loads from its own weight but I wouldn’t expect to see really large differences. The stresses from the mass of the propellant/oxidizer itself are very efficiently distributed and concentrated at one end.
An actual number will have to come from someone who has designed more rockets and possibly fewer potato diggers than me.
Arizona CJ,
Cylindrical tanks are their own support as far as loads from axial acceleration/weight. In a fairly compact structure, the difference between 0 and 300,00 lbs isn’t going to be that great. I doubt that something this size, a few meters in diameter, could be designed to stay together during assembly and launch that couldn’t withstand a strictly axial load this size almost as an afterthought. Other loads would depend on the actual configuration.
Inflatable fuel bladders launched in collapsed (empty) state? Many of the comments prior seem to assume rigid structure designs constrained to fixed sizes. And design constrained by application? A low-gee lander (asteroid/moon) v. planetary insertion would give different results.
As is well known most of the compressive loads of a launch vehicle are supported by the pressurization of the propellant tanks. See for example page 9 of the report:
Launch Vehicle Design: Configurations and Structures.
Space System Design, MAE 342, Princeton University
Robert Stengel
http://www.princeton.edu/~stengel/MAE342Lecture4.pdf
This is true even when the tanks are not of the “balloon” tank variety such as the original Atlas, which could not support its own weight on the ground when not pressurized. It had to be pressurized with nitrogen during storage on the ground.
So without the pressurization provided by the propellant, your stage is likely to not be able to survive the same level of acceleration as when fully fueled.
The solution would be to fill your stage with a light inert gas such as helium which would give the stage the desired amount of stiffness without adding too much weight. You could even reduce the weight further by heating the helium so a smaller amount in mass would give you the same level of pressurization.
Bob Clark
The general rule is that the dry weight of propellant tanks scales with the volume, not the weight, of the contents. So, the dry weight of a tank with volume V shouldn’t change whether you load it with LH2, LOX, or dry air.
Looks like Rand’s kickstarter backers are getting their money’s worth.