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GW, about reusable orbital stages, SSTO's or upper stages, I was surprised in most discussions of them it is assumed they are able to reenter to achieve an only 100 m/s terminal velocity. This is true whether horizontal or vertical propulsive landing is being discussed. See for example here:
http://yarchive.net/space/launchers/hor … nding.html
As the example of the space shuttle shows, this is not so surprising for winged, horizontal landing. But the thing is this is assumed even for vertical propulsive landing. This is surprising because this means for a cylindrical rocket stage, without wings, reentering broadside or a conical stage entering base forward, they can cancel out almost all of the ca. 7,800 m/s orbital velocity to get down to only 100 m/s terminal velocity from aerodynamic drag alone.
Actually, I haven't seen the argument for this. Anyone have a reference for the idea you can get down to 100 m/s terminal velocity even without wings?
But if so why all the big debate over vertical vs. horizontal landing? For either method nearly all the 7.8 km/s orbital velocity will already be cancelled out before either landing method comes into play, and even then they only have to account for a measly 100 m/s.
That is surprising though that without even needing wings you can cancel out nearly all of orbital velocity to get down to terminal velocity. Given that, on Earth reentry for an orbital reusable isn't even particularly hard. So why all the hullabaloo about how difficult it is to get full reusability because you also need to return the orbital stage from orbital velocity?
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  Bob Clark

RGClark wrote:

Quaoar, did you find the Cearun simulator did a good job for the sea level ISP? I’ve been using the free version of the Rocket Propusion Analysis program, http://propulsion-analysis.com/index.htm.

By comparing to known engines, I found it does a good job for the vacuum ISP, but poorly for the sea level value. I’d use Cearun eventhough it doesn’t have a good GUI, it it did do a good job simulating sea level Isp’s.

  Bob Clark

Hi, Bob

I've also tried to simulate the glorious F-1: the vacuum and sea level Isp of the real rocket were about 96% of the virtual one

Bob,

Since you or someone with your screen name appears to have read Hop David's blog post from 21 March 2013, could you explain this concept to Louis in a way he understands?

Maybe I'm just terrible at math today, but could you tell me how you derived a 4% structural mass fraction from the total mass?

1,100t (propellant) + 85t (vehicle) = 1,185t (total mass with no payload)

85 / 1185 = .07172 (structural mass fraction of total mass with no payload)

1,100t (propellant) + 150t (payload) + 85t (vehicle) = 1,335t (total mass with max payload)

85 / 1,335 = .06367 (structural mass fraction of total mass with max payload)

I believe .06367 (6.367%) is as low as it will ever be, given the structural mass and maximum propellant mass figures provided by SpaceX.  I've no clue how they arrived at 85t without building their vehicle, so it must be an estimate of some kind.  If you short load propellant or take away payload, then the structural mass fraction only goes up, not down.  The only way that structural mass fraction can go down is if you can cram more propellant mass into the tanks without increasing structural mass or, of course, if structural mass goes down.

To answer Quaoar in post 7,  yes,  I think your Isp simulator software is doing a decent job,  as far as it goes.  I think you might find a nozzle with an area ratio = 1000 is not something easily built,  and likely rather heavy. 
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GW

Kbd512:
I thought the rampressor was for compressing gases.  The compressible fluid mechanics that makes it work doesn't apply to liquid media,  because the ideal gas equation of state P = rho R T simply does not apply to fluids of such high bulk modulus.  How in the world would a rampressor compress a liquid?  I don't understand.
Lessee,  I'm not sure I remember the names of the two piston-pumped rocket airplanes that XCOR flew.  Those used storable liquids for propellants.  Alcohol and some oxidizer,  I believe that wasn't LOX. 
There was really nothing about the chamber,  injector plate or nozzle that could wear out,  and they had containment in case they cracked.   The pump assembly (and I don't know how it was powered),  seemed to have a life-between-overhauls exceeding 1000 hours. 
What that really means is thousands and thousands of burns,  each a tiny handful of minutes long.  Pretty much about like conventional aircraft engines of most any type,  except for the short flights.
I kinda liked the sound of that long service life.  But it's not a 1:1 substitution for a turbopump assembly on an existing engine.  Doesn't use chamber bleed gas as the drive.  That hot gas bleed is what makes pumps short life.  That's what you want to avoid.
GW

Chinese space official seems unimpressed with NASA’s lunar gateway
China plans to focus its activities on a surface science station.
Eric Berger - 7/17/2018
ARS Technica

This week, the European and Chinese space agencies held a workshop in Amsterdam to discuss cooperation between Europe and China on lunar science missions. The meeting comes as Europe seems increasingly content to work with China on spaceflight programs.

Although the meeting is not being streamed online, space systems designer and lunar exploration enthusiast Angeliki Kapoglou has been providing some coverage of the meeting via Twitter. Among the most interesting things she has shared are slides from a presentation by Pei Zhaoyu, who is deputy director of the Lunar Exploration and Space Program Center of the China National Space Administration.

Overall, Pei does not appear to be a fan of NASA's plan to build a deep space gateway, formally known as the Lunar Orbital Platform-Gateway, at a near-rectilinear halo orbit. Whereas NASA will focus its activities on this gateway away from the Moon, Pei said China will focus on a "lunar scientific research station."
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https://arstechnica.com/science/2018/07 … r-gateway/

GW, that DARPA proposal I wrote was not successful. It was the one we were discussing about new methods of thermal protection in 2013 here:

Index » Interplanetary transportation » Reusable Rockets to Orbit.
2013-10-10 16:47:00
http://newmars.com/forums/viewtopic.php … 19#p117119

This discussion was from 2013. I finally got around to writing the proposal in 2015.

However, what DARPA really wants now in regards to launch access is low cost flights for small payloads, possibly using a reusable booster. The thermal protection issue was not key to that. But what is still a key question and what doomed the X-33 was the inability to get lightweight conformal tanks.

I took a look again at your blog post:

Sunday, October 6, 2013
Building Conformal Propellant Tanks, Etc.

Done successfully,  you have a tank only a few percent heavier than a cylinder of the same volume,  but not heavier by factors.  It will be at least a little bit heavier,  that is inevitable.  That’s simply the price you must pay for the shape you want.  Update 10-7-13:  for the same panel thicknesses and weights as cylindrical construction,  a lower-bound estimate of the weight growth factor is the perimeter length ratio,  computed from cross-section views.

http://exrocketman.blogspot.com/2013/10 … s-etc.html

I'm still struck by your statement that you can get close to the same weight efficiency for lobed tanks as for cylindrical ones using metal tanks. You mentioned the figure of merit that determines the weight growth is perimeter to length...