New Mars Forums

About the cooling issue, this site is a great resource for discussions of spaceflight by experts in the industry back in the day: https://yarchive.net/space/

This page discusses that Pratt & Whitney actually tested a “telescoping nozzle” while the engine was firing and found that it worked:
 
https://yarchive.net/space/rocket/teles … ozzle.html

Also an engine with telescoping nozzle to extend while the engine was firing had been planned for a spaceplane back in the 60’s, but the projected was not completed:

A bat outta Hell: the ISINGLASS Mach 22 follow-on to OXCART
by Dwayne Day
Monday, April 12, 2010

https://www.thespacereview.com/article/1602/1

  Robert Clark

I like that idea. There are several methods of doing variable nozzles. For instance they have been used on jet engine afterburners for decades:

  Bob Clark

GW, in post #14 I showed a graphic that showed how much better an adaptive nozzle could do in ISP over the fixed nozzle applied to the Vulcain engine. I imagine there would also be a great improvement with a variable nozzle over 312s vacuum Isp of the fixed nozzle Merlin 1D. For instance a Russian upper stage RP1 engine was able to get 358s vacuum Isp.

So my question is could your engine analysis program described here, https://exrocketman.blogspot.com/2024/0 … mator.html, do the calculations of the Isp with altitude of the Merlin 1D when given a variable nozzle that matched the exit area to the ambient pressure by altitude?

  Bob Clark

Here’s that discussion by Tim Chen on that LinkedIn group:

https://www.linkedin.com/feed/update/ur … hGKm6TyGa4

He refers to this graphic:

But that graphic takes a far too small value for RP1(kerosene) engines ISP of only 200s. This is worse than the Merlin 1D at 312s vacuum ISP. But the high efficiency Russian engines such as the RD-180 do better than this at 338s.

At a propellant fraction of .95 possible for RP-1, this is a mass ratio of 20 to 1. Using a 338s ISP gives a delta-v of:

338*9.8Ln(20) = 9,923.0636, well above the 9,200 to 9,400 needed for orbit.

Also, a variable nozzle could get even better performance.

   Bob Clark

I started a group on LinkedIn on developing SSTO’s:

SSTO - Single Stage to Orbit.
https://www.linkedin.com/groups/13205030

My first post in that group was on developing an expendable SSTO. The reason is the opinion against a SSTO being feasible is so strong it’s important to break that mindset by just getting an expendable one with significant payload. 

  The point of the matter is getting an expendable SSTO is easy with minor modification to existing engines: add an extensible nozzle to get high vacuum ISP at vacuum, but which can be retracted at sea level. Such extensible nozzles have been in use since the 90’s with the RL10 upper stage engine. The problem is they haven’t been used on first stage engines.

The surprising conclusion you get is any existing (liquid fueled) stage can be made SSTO by addition of an extensible nozzle, though for an upper stage you would have to cut fuel load or add additional engines since commonly they don’t have enough thrust for liftoff from ground since they don’t need it.

But there is a problem here as far as estimating payload capability. Commonly with fixed nozzle stages you use a fixed value for the ISP to estimate the delta-v possible. This is convenient since it is only a single line calculation using the rocket equation. But that probably is not accurate for a variable nozzle since you are changing the characteristics of the nozzle through out the flight. You could you use, say, the vacuum isp for this case but that would probably over estimate it.

What really needs to be done is accurate trajectory simulations of the type NASA uses. Then you could see how the variable nozzle deviates from what a fixed nozzle would do. I’ve only seen one paper actually do that. I’ll see if I can find it.

  Bob Clark

Bad news on the latest Starship static fire test:

@Blobifi
Starship gazer has released a video of the static fire on Facebook showing what looks to be an Rvac destabilizing before the the other engiens shut down.
https://x.com/blobifie/status/191824849 … QWCAYS9AQw

Some reports are a vacuum Raptor actually experienced a RUD.

  Bob Clark

Yes, it’s Archimedes. By the way when the question is asked who were the greatest intellects humanity has ever produced the answer commonly given is Archimedes, Newton, and Gauss.

The basic principle discovered by Archimedes, startling in its scope, is the principle of “mechanical advantage”. The idea is moving a small mass a large distance or high speed is equivalent to moving a large mass a small distance or low speed. Archimedes first enunciated this principle in regards to the lever. He famously said, “Give me a place to stand and I shall move the Earth!”

Nowadays, we consider it “trivial” and “obvious” because the energy is the same in both cases. You see it in many forms, levers, pulleys, gears, hydraulics, etc. But note the principle goes both ways: if you want to achieve high speed you can do it by moving a large mass low speed.

Then in regards fusion and superluminal speeds you translate the known techniques of producing large mass at low speed into moving small mass at (extremely) high speed.

I just thought it stunning such a simple principle discovered millennia ago could solve the deepest questions of 21st century physics.

  Robert Clark

Thanks for that, GW. I wanted to ask in regards to the “hot metal” TPS, the heat would rapidly propagate to the upper side of the vehicle. This means the radiative surface area would double, thus doubling the heat emission.

Shouldn’t this improve the survivability?

  Bob Clark

I discussed before I think SpaceX is not taking the best approach to developing the Superheavy/Starship. They were spectacularly successful with the Falcon 9 by first getting the expendable version, then proceeding to reusability. If they had taken that approach with the SH/SS they would already be flying the expendable version and perhaps even also the partially reusable one, i.e., reusing the booster only, a la the Falcon 9.

Note then for the expendable version the dry mass of the Starship might have been as low as 40 tons:

Elon Musk @ElonMusk
Probably no fairing either & just 3 Raptor Vacuum engines. Mass ratio of ~30 (1200 tons full, 40 tons empty) with Isp of 380. Then drop a few dozen modified Starlink satellites from empty engine bays with ~1600 Isp, MR 2. Spread out, see what’s there. Not impossible.
https://x.com/elonmusk/status/1111798912141017089?s=61

Then with the ca. 20 ton fairing, the needed ‘wing’ area need to be added might be ca. 1,800 sq.m. But this wouldn’t be as heavy as regular wings with their thickness to generate aerodynamic lift. It would only have the character of a thin plate since it is meant only to be a drag decelerator.

  Bob Clark

Keep in mind how huge Starship is. Jet fuel, kerosene, is a bit less than water’s density of a ton per cubic meter. Then the Antonov’s wing fuel load would be about 300 tons and the Concorde’s about 100 tons. The Starship carries 1,200 tons of propellant.

But it’s a bit like comparing apples to oranges. In this scenario the “wing” would not be carrying fuel. It would only be acting as a drag decelerator, a la the ‘parashield’.

Still, I’d like to see what could done with true wings using high lift/drag aerodynamics at hypersonic speed. This would undoubtedly make the wings smaller.

Optimum Mach 25 waverider [from Bowcutt, Anderson and Capriotti, 1987

  Bob Clark

SpaceX is having difficulty finding an effective thermal protection system for Starship. I had earlier speculated that if it were giving sufficiently large wings, it could go completely without TPS. The possibility was occasioned by this article:

Wings in space.
by James C. McLane III
Monday, July 11, 2011
Wing loading (the vehicle’s weight divided by its wing surface area) is a prime parameter affecting flight. The antique aluminum Douglas DC-3 airliner had a big wing with a low loading of about 25 psf (pounds per square foot of wing surface). At the other end of the spectrum, the Space Shuttle orbiter has a high wing loading of about 120 psf. This loading, combined with an inefficient delta-shaped wing, makes the orbiter glide like a brick. A little Cessna 152 private plane features a wing loading of about 11 psf and modern gliders operate down around 7 psf. A space plane with huge lifting surfaces and a very low wing loading might not require any external thermal insulation at all. Building a space plane with a wing loading of, say, 10 psf should not be an impossible proposition. Perhaps some day it will be done.
http://www.thespacereview.com/article/1880/1

I found a further article that allows this possibility to be further constrained:

SpaceOps 2010 Conference
25-30 April 2010, Huntsville, Alabama AIAA 2010-1928
Applications of Ultra-Low Ballistic Coefficient Entry Vehicles to Existing and Future Space Missions
David L. Akin∗
Space Systems Laboratory, University of Maryland, College Park, MD 20742
https://spacecraft.ssl.umd.edu/publicat … hieldx.pdf

As discussed there the relevant parameter is actually ‘ballistic coefficient’, (mass)/(drag coefficient*drag area), β = m/CD*A, given in metric units kg/m^2, where the drag area is by cross-section.

The author uses a slightly variant definition given in units of pascals where up in the numerator is given the weight in Newtons. But it’s easy to convert to the more commonly used version by dividing by g, 9.81 m/s^2, i.e., about 10. He estimates when ballistic coeffcient is below 200 Pa, or about 20 kg/m^2 in the more common units, the max temperature during reentry would be ca. 800°C. This should be a temperature stainless-steel is able to withstand.

I estimated an additional wing area of 36m*50m = 1,800m^2 would allow the max temperature to stay below the 800°C point:

Reentry of orbital stages without thermal protection, Page 2.
https://exoscientist.blogspot.com/2025/ … thout.html

  Bob Clark