Copyright 2012 Robert Clark
SASSTO
SASSTO - Saturn-derived SSTO Launch Vehicle
Credit: © Mark Wade
I showed in the post
Low Cost HLV, page 2: Comparison to the S-IC Stage that the S-IC first stage of the Saturn V could give a nearly 20-to-1 mass ratio using a lighter thrust structure and using four
RD-171 engines instead of five F-1 engines. But in fact we can do better than this. The S-IC of the 1960's did not have available the aluminum-lithium alloy used for example on the Falcon 9 and shuttle ET. Here I calculate a lighter structure using this lighter alloy and some mass reducing structural changes.
The tank mass of the S-IC stage and of some other rocket stages is discussed in this key report:
Single Stage To Orbit Mass Budgets Derived From Propellant Density and Specific Impulse.
John C. Whitehead
32nd AIAA/ASME/SAE/ASEE Joint Propulsion ConferenceLake Buena Vista, FLJuly 1-3, 1996
http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/379977.pdf
On page 8 in Fig. 6 is given a comparison of the tank weights of the Saturn S-IC, Atlas II, and shuttle ET:
We've already reduced the mass of the stage down to 113,000 kg by using a lighter thrust structure in the prior post. Fins are not really needed for large rockets with computer guidance and control, so remove these to reduce the mass again to 112,000 kg. The four 4 meter diameter RD-171 engines can fit under the 10 meter diameter S-IC tank, so remove the engine fairings to bring the mass down to 108,000 kg.
Now use common bulkhead design to reduce the tank mass further. The question of using common bulkhead design for the S-IC arose during the Apollo design period:
SP-4206 Stages to Saturn.
7. The Lower Stages: S-IC and S-II.
THE S-IC AND THE HUNTSVILLE CONNECTION.
The main configuration of the S-IC had already been established by MSFC, including the decision to use RP-1, as opposed to the LH2 fuel used in the upper stages. Although LH2 promised greater power, some quick figuring indicated that it would not work for the first stage booster.Liquid hydrogen was only one half as dense as kerosene. This density ratio indicated that, for the necessary propellant, an LH2 tank design would require a far larger tank volume than required for RP-1. The size would create unacceptable penalties in tank weight and aerodynamic design. So, RP-1 became the fuel. In addition, because both the fuel and oxidant were relatively dense, engineers chose a separate, rather than integral, container configuration with a common bulkhead. The leading issue prior to the contract awards related to the number of engines the first stage would mount.
http://history.nasa.gov/SP-4206/ch7.htm#197
This could be interpreted to mean the density of the propellant made it unfeasible, but I think the relatively smaller tankage mass using the dense propellants made the more difficult common bulkhead design unnecessary. For instance as you see in that Fig. 6 from Whitehead's report, in the shuttle ET tank the intertank weighs more than the entire oxygen tank. The relative weight of the intertank is not as bad for the kerosene S-IC. Still, common bulkhead design is used for the large kerosene first stage on the Falcon 9 to help save weight.
So to minimize stage weight we will remove the interstage and one of the bulkheads. Assuming top and bottom bulkheads weigh the same for each of the LOX and kerosene tanks on the S-IC, then from the information in Fig. 6, the LOX bulkheads weigh 4 mT each and the kerosene, 3.3 mT. Conservatively, let's say we remove one of the kerosene bulkheads instead of a LOX bulkhead since we may need the larger LOX bulkhead for strength. Then also removing the 6 mT intertank, we bring the dry mass down to 99 mT.
Now estimate the weight saving using the lighter aluminum-lithium alloy. From the
Wikipedia page on the shuttle ET, the tank weight reduced from 35,000 kg using aluminum alloy 2219, the same alloy used for the S-IC tanks, to 26,500 kg using aluminum-lithium alloy, a reduction of 24%.
After the structural changes, the tanks now weigh 25.5 mT. Subtracting off 24% from this is a reduction in mass by 6 mT. This brings the stage mass down to 93 mT.
Keep in mind though, the plan is to use a shuttle ET size tank to save cost on tooling. The ca. 720 mT hydrolox of the shuttle ET becomes ca. 2,100 mT with the 3 times denser kerolox. This turns out to be about the same kerolox carried by the S-IC. So the purpose here was just to get an idea of a lightweight stage you can get using modern materials.
Now notice you get significant payload as a SSTO using the RD-171 engines at 338 s vacuum Isp. Taking the required delta-v to orbit as 9,150 m/s for kerolox, you can get 48 mT to orbit:
338*9.81ln(1 + 2,100/(93 + 48)) = 9,170 m/s.
Note though that if we are to use the shuttle ET as a stage then the pointed end of the LOX tank would need to be removed. We could take the equivalent cylindrical LOX tank of the same volume. It would have the same dry weight, so the stage dry mass stays the same.
However, if you take the full length of a cylindrical tank now as 46.9 m and the diameter as 8.4 m, per the
specifications of the SLWT version of the ET, and the density of kerolox as about 1,030 kg/m^3, then we get about 2,600 mT kerolox. The tank weight would increase somewhat without the pointed end, but not by much compared to the entire stage weight. Then you could loft 82 mT to orbit:
338*9.81ln(1 + 2,600/(93 + 82)) = 9,160 m/s.
A propellant load of 2,600 mT at dry mass of 93 mT corresponds to a mass ratio close to 29 to 1, rather high. But SpaceX has said with their side boosters on the Falcon Heavy they expect to achieve a mass ratio of 30 to 1, and mass ratio does get better as you scale up a stage,with this shuttle ET size stage being much larger.
This payload of 82 mT is better than the 70 mT to be carried by the interim SLS. Remember our HLV is to be developed using the SpaceX-style commercial approach. Then based on a $2,000/kg price of the Falcon Heavy, the full two stage version of our HLV as comparably priced might only cost ca. $200 million per launch at a 100 mT payload to orbit.
So the SSTO version would even cost less than this, perhaps only ca. $100 million per launch for the 82 mT payload to orbit.
The dry weight could be lightened further by using composites. Estimates put the weight savings in the structural mass in the
40% range for a fully composite structure. In that case the payload could exceed 100 mT for this SSTO.
An increase of the Isp could be possible by using an
aerospike or plug nozzle, up to the range of 360 s. The multi-nozzle format of the RD-171 engines makes this feasible. The four nozzles of each engine would be shortened and arranged around a central aerospike. This was the idea behind the aerospikes planned for the X-33 and VentureStar. It was also used earlier in the planned
Beta SSTO of Dietrich Koelle and the
SASSTO SSTO of Phillip Bono.
An argument against this was that the aerospike nozzle would make the propulsion system too heavy. For instance for the aerospike on the X-33 the thrust/weight ratio was only 40 to 1, compared to a 70 to 1 ratio for the SSME's for example. However, the lightweight, high temperature ceramics and composites available now should make the T/W comparable to bell nozzle engines:
Ceramic Materials for Reusable Liquid Fueled Rocket Engine Combustion Devices.
http://ammtiac.alionscience.com/pdf/AMPQ8_1ART06.pdf
Bob Clark
Note: The SASSTO SSTO shown at the top and discussed near the end was derived from the Saturn S-IVB stage, not the S-IC, and was hydrogen fueled. The Beta SSTO discussed was also hydrogen fueled. However, a key result of the cited report of John C. Whitehead is that it is actually easier to make a kerosene-fueled SSTO. This is because the large and heavy hydrogen fuel tanks swamps out the advantage of its higher Isp. - B.C., 6/22/2012.
