New Mars Forums

Perhaps someone can see what's wrong with this proposal of a SSTO. It's too obvious to not have been thought of before:

High performance engines that you would need to get high thrust at sea level should have high chamber pressure. Such engines need to be turbopump driven to generate these high pressures. Such engines though cost hundreds of millions in development costs, an expensive prospect for something as theoretical and controversial as an SSTO.

Pressure-fed engines on the other hand are easy and low cost to produce. In fact amateurs or small commercial concerns such as Armadillo Aerospace, Masten Space, and Garvey Space have made such engines and rocket stages. These are always low chamber pressure engines. You could make a high chamber pressure pressure-fed engine but then the propellant tanks would also require such high pressures, and this would result in the tanks being too heavy for a SSTO.

BUT how about this method to generate the high pressures in the engine for a pressure-fed system that at the same time has the  relatively low pressure tanks of turbopump systems: the long cylindrical propellant tanks produce additional pressure at their bottoms from the weight of the propellant. Suppose we then have a pipe leading down from the bottom at 1/10th the diameter of the tank, so 1/100th the cross-sectional area. Shouldn't this result in the pressure within the pipe being 100 times that of the pressure within the tank?

So even if the pressure in the tank is only, say, 2 bar, typical of turbopump systems, the pressures in the pipes leading into the engines can be 200 bar. Correct?

But what about when the propellant becomes more and more burned off resulting in smaller and smaller weight to cause the pressure at the bottom? As the propellant gets burned off the weight decreases so the T/W goes up and so does the acceleration. Then the effective "weight" of the propellant also goes up, maintaining the same pressure at the bottom. You may want to throttle the engine though to limit the structural loads on the craft. But another consideration is that high up at near vacuum conditions high chamber pressure and thrust are not so important. What is important is high Isp, which can be done by using altitude compensation such as an aerospike.

A couple of problems though. Shouldn't this already be happening with the rockets used now? In that case why are turbopumps being used at all?

Secondly, the propellant tanks in rockets are pressurized aside from the increased pressures observed at the bottom due to gravity. So shouldn't this in itself without any consideration of the propellant weight or thrust or acceleration allow the reduced pipe diameter to have increased pressure inside the pipes?

Thirdly, there is the issue of  Pascal's Principle:

http://hyperphysics.phy-astr.gsu.edu/hbase/pasc.html

This says that a pressurized liquid equalizes pressure throughout the vessel whether at wider or narrower parts of the vessel. In that case shouldn't the pressure be the same even in the narrower pipe?

  Bob Clark

I've been thinking of ways we can get such nanocraft to link up through self-assembly and form larger structures that can do more detailed observations and experiments. This could work even for visits to far off destinations still in the Solar System such as Kuiper belt objects like Pluto or the Oort cloud.

The main problem is getting the many objects flying independently and getting further apart the further out they go to gradually be drawn to each other and link up. Once they link up, I don't it would be too difficult to then get them to do self-assembly.

But it's that drawing together step that is the hard part.

Some ideas on how they might be made to link up: perhaps the light sails can be angled so that they would be directed to conglomerate at a common point. It is known that solar/laser sails can do "tacking" to change their direction.

Another possibility is that there will be the ionized solar wind and interplanetary and interstellar dust that the nanoprobes could react against to be directed to a common point.

BTW, I don't think it would too difficult to do the self-assembly. Still I'd like to get some feedback on how it could done. Note the nanoprobes are consider to be about the size, and complexity of, say, a virus, or RNA molecule. This also raises the possibility of alien species using this to seed other star systems with life.

   Bob Clark

Elon Musk has confirmed the F9 first stage can reach orbit as an SSTO:

Elon Musk Verified account @elonmusk @TobiasVdb
The F9 booster can reach low orbit as a single stage if not carrying the upper stage and a heavy satellite.
https://twitter.com/elonmusk/status/669132749500887040

Thank you very much, Mr. Musk. See:

The Coming SSTO's: Falcon 9 v1.1 first stage as SSTO, Page 2.
http://exoscientist.blogspot.com/2013/1 … first.html

Bob Clark

  Yes. That is indeed odd. The entire purpose of the upgrade was to increase the payload to GTO while maintaining reusability, in fact by 33%. But this should also increase the payload to LEO as well.

  Bob Clark

'Red Dragon' Mars Sample-Return Mission Could Launch by 2022.
by Mike Wall, Space.com Senior Writer   |   September 10, 2015 09:00am ET

http://www.space.com/30504-spacex-red-d … eturn.html

  Can be done even more cheaply than this, in fact by a single Falcon 9 launch. Using a fueled Dragon forces a 10 ton lander in the mission. Much smaller propulsive stages exist to reduce the mission size.

   Bob Clark

Falcon Heavy might match first version of SLS.

SpaceX has said the improvement of the performance with the Falcon 9 v1.2 would be about 30%. With the upgraded engines and densified propellants applied to the Falcon Heavy, the max payload of the FH to LEO could then also be increased from 53 metric tons to 70 mT.
If so, then the FH, at a ca. $100 million cost, could match the capability of the initial Block I version of the gigadollar SLS at 70 metric tons to LEO.

First Falcon Heavy Launch Scheduled for Spring.
by Jeff Foust — September 2, 2015
http://spacenews.com/first-falcon-heavy … or-spring/

Actually, I'm dubious of the cited payload of the Block I SLS as 70 mT. I think it's likely to be closer to 90 mT:

SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 5: A 90+ metric ton first launch of the SLS.
http://exoscientist.blogspot.com/2013/0 … -50th.html

   Bob Clark

Just released image of the mysterious Ceres mountain imaged at closer Survey mission distance:

Aug. 25, 2015
Dawn Sends Sharper Scenes from Ceres.

http://www.nasa.gov/jpl/dawn-sends-shar … from-ceres

  Bob Clark

I'm informed by a NASA nuclear rocket engineer that the high number of 148,000 watts/kg really is in the range of the power-to-weight ratio of nuclear rocket engines. But this means when you do the conversion to electrical power you reduce the efficiency by three orders of magnitude! In other words the lack of efficiency really does not have to do with the nuclear space reactor itself.
One problem with the conversion to electric power is poor efficiency methods are used. For instance for the American systems thermoelectric conversion is used, which typically is only 5% to 8% efficient.

But for the power to weight ratio to drop so low it must be the additional weight of the electrical equipment that contributes to it. And indeed the specific power of electric motors is at most in the 10,000 watts/kg range:

Power-to-weight ratio.
2.1.2 Electric motors/Electromotive generators
https://en.wikipedia.org/wiki/Power-to- … generators

Still if we used an electric motor at this efficiency level, run in reverse to generate electricity, then it should give the needed specific power.

  Bob Clark

The Falcon 9 and Ariane 5 core also use common bukhead design.

   Bob Clark

Thanks. The conference was a lot of fun. The formula I used is equivalent to the one you cited. I just wanted to save having to calculate the massflow rate. That the two are equivalent comes from the fact that thrust = (massflow rate)*(exhaust velocity). So:
power = .5*(massflow rate)*(exhaust velocity)*(exhaust velocity) = .5*(thrust)*(exhaust velocity).

  Bob Clark

The Mars Society convention in Washington, DC went great. The two presentations I gave were well attended. These track talks were not livestreamed as were the primary invited talks. However, they will be put up on the net at a later point. I'll post the links to the videos when they are available. The topics of my presentations are described in these blog posts:

http://exoscientist.blogspot.com/2015/0 … etary.html

http://exoscientist.blogspot.com/2014/0 … sible.html

  Bob Clark