Exoscientist

Peter Diamandis, progenitor of the X-prize has proposed an answer:

Thanks for the response. I think we can get a better payload as an expendable considering the high average Isp of the SSME’s, approx. 4,400 m/s. This page takes the delta-v to orbit as only 9,000 m/s , about right for a base 200 km orbit and launching eastward to take advantage of the Earths 400+ m/s rotation speed near the equator: Towards Reusable Launchers - A Widening Perspective. https://www.esa.int/esapub/bulletin/bullet87/pfeffe87.htm Then using your numbers: 4400Ln(1 +306/(36 + 9.4)) = 9,000 m/s, for a 9.4 ton payload as expendable. Robert Clark

Can you say what is the bare dry mass and propellant load is for the expendable rocket (no heat shield or landing legs)? Then what is the expendable payload to a 200 km basic orbit. Then there are various materials we can use to get light weight heat shield and landing legs. Robert Clark

I would love to do my own Realism Overhaul simulation, but I have never learned Kerbal. (What would be the suggested approach to take to first learn Kerbal, and then proceed to Realism Overhaul?) Rather than your digging into the details of my proposed Delta IV Medium's first stage modified with two SSME's and lightweighted tanks, you could just assume some hydrolox stage with a 10 to 1 mass ratio, since the existing Centaurs get that, and then give the stage SSME engines. I think you'll be surprised at how high the payload you could get as an expendable. Given the high payload fraction, you could then estimate how much mass is needed for thermal protection, landing legs, and propellant needed to be kept on reserve for a vertical, powered landing for the reusable case. Alternatively, you could estimate the thermal protection, landing legs and wing mass needed for the horizontal landing method of reusability. In this regard, if you're using a non-lifting ascent to orbit, the wings only have to be sized to support the dry mass of the craft on return. This will significantly reduce their mass. I also advise using short, stubby wings, not the oversized, heavy wings used on the shuttle. As in these examples:

Thanks for the calculation. As I math guy I love to see the calculations. We're probably pretty close in our estimates if you're taking, say, 350+ s for the Merlin vacuum Isp with "altitude adaptation"( below I'll give a link to a paper that uses this terminology.) The only difference is our interpretation of that 9,720 m/s delta-v you calculated. For just going to LEO, that's actually pretty good. Such a vehicle could carry quite a bit of payload if you take the required delta-v to LEO as only 9,000 m/s. This is the value taken here: Towards Reusable Launchers - A Widening Perspective. https://www.esa.int/esapub/bulletin/bullet87/pfeffe87.htm Here's an article that examines various types of "altitude adapting" nozzles: Advanced Rocket Nozzles. September 1998 Journal of Propulsion and Power 14(1998):620-633 Authors: Gerald Hagemann, Ariane Group Hans Immich, Aerospace Consultant T. van Nguyen, Aerojet GenCorp, Inc. Gennady Dumnov, Siemens https://www.researchgate.net/publication/224796963_Advanced_Rocket_Nozzles The image shows how much higher an ideal altitude adapting nozzle can get in vacuum Isp compared to a standard bell nozzle. This is ideal but the point of the matter is rocket engines get quite high efficiency in the range of 95%+ Robert Clark

Thanks for that. But what I want first is an estimate of the payload as a simple expendable rocket of the Delta IV Medium equipped with SSME's. Then you can try various landing modes such vertical propulsive or winged horizontal. Not having wings for the expendable version should simplify the aerodynamics. Robert Clark

The expander cycle is among the least efficient engines for sea level engines, due to its limited thrust. That is why it is used for upper stage engines where high thrust is not required. But a gas generator and staged combustion engine used for sea level launch can also get high vacuum Isp by using an extended nozzle. Robert Clark

This part in the passage you quoted from Sutton that you did not highlight is why expander cycle engines are not used for sea level engines and gas generator and staged combustion engines are used instead, because these types of engines can provide more power, which is needed for launch from the ground: Liquid Rocket Propulsion. Engine Cycles – Expander • Relies on a turbopump to force propellants from tanks to the combustor • Tanks kept at lower pressures • Fuel heated via regenerative cooling process and passed through turbine to drive pumps • Thrust-limited due to square- cube rule (heat transfer) • RL10 (Delta IV, Atlas V), LE-5B (H-IIA, H-IIB) http://rocket.gtorg.gatech.edu/files/slides/Liquid_Rocket_Propulsion.pdf In any case, it is still the case that ANY engine of any combustion cycle can get high vacuum Isp by using long nozzles. Some further references: Rocket Engine Nozzle https://en.wikipedia.org/wiki/Rocket_engine_nozzle Rocket Propulsion Elements, by Sutton and Biblarz http://mae-nas.eng.usu.edu/MAE_5540_Web/propulsion_systems/subpages/Rocket_Propulsion_Elements.pdf [full text] Rocket Propulsion. http://www.braeunig.us/space/propuls.htm Robert Clark

There are underwater acoustic systems that can communicate underwater: https://dosits.org/people-and-sound/communication/how-is-sounds-used-to-transmit-data-underwater/ Robert Clark

Thanks for the very interesting news clip. It wouldn’t take a nuclear mission to land on Enceladus. NASA is planning an SLS mission to Europa, and the SLS could also reach Enceladus with a lander. Actually, a lander mission to Europa could be accomplished much more cheaply using the Falcon Heavy: Low Cost Europa Lander Missions. https://exoscientist.blogspot.com/2015/02/low-cost-europa-lander-missions.html Likewise a lander mission to Enceladus, of smaller size, could be mounted using the Falcon Heavy. Robert Clark

A key issue is it continues to be said SSTO’s aren’t technically feasible. Whenever they are spoken about it is like you need nuclear engines or some “unobtainium” materials. That it has been feasible since the 70’s is an important fact. Doing an accurate trajectory simulation is therefore important to establish that fact. When you did the accurate trajectory simulations you would find an optimized rocket that uses both maximally efficient engines and maximally lightweighted tanks would as an SSTO have better payload than current rockets which are not optimized. Moreover such an optimized rocket when used in a TSTO version would double the payload of a current TSTO which is not optimized. So even if you don’t want to use SSTO’s, it is important to confirm this using accurate trajectory simulations to prove you can double the payload of a current TSTO using simple methods already known about for decades. Robert Clark

I’m sorry but this is simply incorrect. Any rocket engine can get high vacuum Isp by using a long extension nozzle, as indicated by the Merlin Vacuum getting 348s Isp. The expander cycle used on the RL-10 is actually one of the least efficient cycles used on rocket engines. It was used on the RL-10 because as an upper stage engine all it needed was a long nozzle, not combustion efficiency, to get high vacuum Isp. Remember this is in regards to vacuum Isp. For sea level engines you need high thrust and gas generator cycles and staged combustion cycles are better at providing that. Robert Clark

We’ve been discussing the advantages or disadvantages of making such advances such as max Isp engines and max lightened stages. This has only been estimated by the rocket equation. What really needs to be done is a Kerbal simulation using the Real Solar System mod or Realism Overhaul to show the Delta IV given SSME’s and lightweighted stages really can get the high payloads as an SSTO and TSTO suggested by the rocket equation estimates. Robert Clark

The designers of the Centaur want the long nozzle because they do want the high Isp in vacuum it provides. They could get this just in being fixed. But by making it extendible they are able to fit it in a shorter interstage. This is a key principle about rocket engines that by using wider nozzles you are able to get higher vacuum Isp. The problem with using it at sea level is such large nozzles cause what is called flow separation. This is a dangerous instability condition that can literally tear an engine apart. This is why such large nozzles are not used at sea level. The idea behind variable area nozzles is they are small at sea level and extend larger at altitude in vacuum. This effect can be emulated also by using an aerospike. If we are to make it be an actual extensible nozzle we can use recent materials using ceramics that can cut the weight by a factor of 3. Robert Clark Actually the grade SpaceX wants to use is much stronger than standard steel and much more expensive. As I mentioned in my blog post there are some metals now that can match and exceed carbon composites. That’s perfectly fine, as long as they reduced the structural mass of the stage. Carbon composites stages are used in some rocket stages, most commonly in solid rockets. SpaceX didn’t use it not because it wouldn’t work, but because the carbon composites are far less heat resistant than steel and would have required more mass for thermal shielding. Robert Clark

We’ve been discussing in this thread the feasibility of creating this as a real SSTO: My argument that it is doable is just by using the rocket equation. Who wants to take up the challenge of doing a Real Solar System mod simulation showing it is possible? Robert Clark

The vacuum Isp of the RS-68 is actually only 412 s. That is why performance would be radically improved by increasing this to 470 s. You can get a high vacuum Isp on an upper stage engine just by using a nozzle extension. So for example the RL-10 engine on the Centaur upper stage gets a ca. 462 to 465.5 s Isp by using an extendable nozzle attachment. Robert Clark

Yes. But for an upper stage getting high vacuum Isp is largely an effect of using a large expansion ratio. For instance the Merlín Vacuum at 348 s vacuum Isp has a expansion ratio of 164 to 1 using a quite large nozzle attachment: But the Russian RD-58s gets a vacuum Isp of 361 s by using an expansion ratio of 189 to 1: https://web.archive.org/web/20160407043338/http://www.friends-partners.org/partners/mwade/engines/rd58s.htm The importance of altitude compensation is it allows you to get the very high vacuum Isp of a upper stage engine while being being able to launch from the ground, as a sea level engine. It does this by using a variable area nozzle, or a nozzle such as the aerospike that acts as a variable area nozzle. Actually the 348 s Isp of the Merlin Vacuum is quite good. Even if you used a variable area nozzle on the sea level Merlins that just reached the 348 s value, that would also greatly increase the TSTO payload, and allow an SSTO with significant payload. Robert Clark

Actually, not. This shows why both these advances should be undertaken. By using both, we can reduce the cost of space and improve on the payloads of rockets we have now. Robert Clark

Again, when I say an “SSTO can match or beat a TSTO in payload”, that is shorthand for this: Current rockets do not use maximally efficient engines or weight optimized structures on their first stages. But high Isp and low weight are extremely important for maximizing a rockets payload to orbit. So when you do use both of these, it radically improves payload to LEO. The result is the SSTO can match or exceed the payload of a current TSTO. And the two stage version of such an optimized rocket can double the payload of a current rocket. Robert Clark No, I’m not. I am literally saying an SSTO that uses both maximum Isp engines and weight optimized structures can exceed in payload fraction that of a current TSTO, which do not use either of these. So the same size optimized SSTO can exceed the payload of an unoptimized TSTO. Robert Clark

A fair point. One way would be to shorten the nozzles then place them around a central spike, as is planned for the Firefly rocket: Robert Clark

For those interested in the topic of SSTO’s a nice book to read is: Halfway to Anywhere: Achieving America's Destiny In Space. https://www.amazon.com/Halfway-Anywhere-Achieving-Americas-Destiny/dp/0871318059/ There, Stine also argues a SSTO can match the payload of a TSTO. Robert Clark

You missed the point I’m making. The claim is that SSTO’s can’t be competitive to current rockets. But current rockets do not use max efficiency engines or max weight optimized stages! The importance of having a much higher Isp and weight optimized stages is so important that the SSTO will match or exceed the payload of the TSTO that do not use them. Yes, the payload of the TSTO that also does use them will be higher than the SSTO. But still the SSTO will be more efficient than the current rockets that do not. Robert Clark

Everyone knows the great importance of the Isp on the payload a rocket can carry to orbit. Imagine the Merlins used on the first stage having their vacuum Isp being raised from 312 s to 360+ s by using altitude compensation. Now also consider the weight saved on the aluminum-lithium tanks by using carbon fiber is about 30%. Not as good as the weight saved over standard grade aluminum of 50% but still pretty good. Try now the rocket equation with a vacuum Isp of 360+ s to see the payload the first stage can get orbit. You can use the dry mass and propellant mass numbers for the Falcon 9 v1.1 here: https://www.spacelaunchreport.com/falcon9ft.html#components You’ll be surprised by how high the payload is as an SSTO. Unfortunately, this isn’t quite accurate. The commonly used method of estimating the payload by the rocket equation is uncertain for the altitude compensation case. The reason is the estimate is known to be reasonably accurate for using fixed nozzle engines. But its accuracy is unknown for the altitude compensation case. A real trajectory simulation with an altitude compensating nozzle needs to be done to calculate this. Robert Clark

These articles discuss the issue of detecting nighttime light signatures in exoplanets: Proxima Centauri b: Artificial Illumination as a Technosignature by PAUL GILSTER on MAY 21, 2021 https://www.centauri-dreams.org/2021/05/21/proxima-centauri-b-artificial-illumination-as-a-technosignature/ Lights of the Nightside City by PAUL GILSTER on MAY 24, 2021 https://www.centauri-dreams.org/2021/05/24/lights-of-the-nightside-city/ They link also to research articles on the topic. A proposed upcoming mission LUVOIR may be able to detect nighttime lighting at perhaps 10 times higher density than on Earth on an exoplanet of Próxima Centauri: Large Ultraviolet Optical Infrared Surveyor. https://en.m.wikipedia.org/wiki/Large_Ultraviolet_Optical_Infrared_Surveyor Robert Clark

Rather than the commonly made statement that SSTO’s are not technically feasible or that they can’t carry significant payload, the actually truth of the matter is that if you use both high performance engines and lightweighted tanks, then SSTO’s can carry just as much or more payload in terms of payload fraction as do the current TSTO’s for expendable launchers. Moreover, because of the greatly reduced payload of the TSTO on boostback to launch site, the SSTO can actually carry more payload as a fully reusable launcher. That’s a stunning fact. The exact opposite of what is said about SSTO’s is the case. Robert Clark