What are your ultimate commercial rocket designs?

[Hyperion5]

I was inspired after reading Spacex's page on the Falcon 9 and then seeing some interesting comments & claims by Spacex, Musk & ULA to create this thread. 

Spacex points out that their design for the Falcon 9 was shaped by a report on launch failures:
http://www.spacex.com/falcon9.php

"The vast majority of launch vehicle failures in the past two decades can be attributed to three causes: engine, stage separation and, to a much lesser degree, avionics failures. An analysis (p. 23) of launch failure history between 1980 and 1999 by Aerospace Corporation showed that 91% of known failures can be attributed to those subsystems."

To create the ultimate commercial rocket and counter those issues, Spacex brags about their design features:

Triple-redundant avionics with GPS overlay
Engine-out capability & reliability on SI
(They claim the Falcon 9 features "an improved" version of the design architecture of Saturn rocket series (Saturn I, IB, & V)
Same propellant tanks used in both stages (+ cost savings)
Road-transportable stages (only 3.66 meters in diameter)
Superior engine design eliminates hydraulic failures
Then there's the simple propellants to deal with, namely non-cryogenic RP-1 & modestly cryogenic Lox

Musk believes the Falcon 9 series is the best and cheapest commercial launcher on the market today.  He openly scoffs at one of his major competitor's rockets:  "Ariane 5 has no chance," he told BBC News.  "I don't say that with a sense of bravado but there's really no way for that vehicle to compete with Falcon 9 and Falcon Heavy. If I were in the position of Ariane, I would really push for an Ariane 6."
http://www.bbc.co.uk/news/science-environment-20389148

Of course ULA doesn't agree totally with Spacex' premise even if they aren't launching many commercial payloads: “Not to minimize SpaceX’s impressive achievements, but ULA’s customers want to see a track record of success, repeatably delivering complex payloads to orbit, safely and on time.  In the launch business, price is never the sole consideration for the buyer. That’s because launch price is a small percentage of the total program value (which can exceed replacement cost when there’s no money to replace, like the Glory spacecraft).” http://www.nasaspaceflight.com/2012/11/ula-customers-class-reliability-main-consideration/

Elon Musk clearly believes he's got the best commercial rocket design on the market today.  Some of his competitors would beg to differ.  I'm not convinced there is a really ideal commercial rocket design out there though.  All of the top contenders have some crucial flaws in their design that makes me hesitate to say, "this is as good as commercial rocket design gets".  So that brings me up to this question set and the ultimate challenge for everyone. 

1) What would your ideal commercial rocket designs look like?  Explain why your rocket designs would be superior to all competitors. 

There's a couple of markets I think you can divide this into.  Commercial launchers tend to send things to one of two areas currently. 
Low-earth orbit
Geosynchronous transfer orbit

There are several other places you might eventually send commercial payloads. 
Lunar orbit
Lunar surface
Earth-Moon Lagrange points
Beyond-Earth Orbit (probably to asteroids (Planetary Resources) or Mars) 

I would urge you to pick a couple of target markets or less & if possible, and refine the design by selecting a payload range or the amount of crew you hope to send there.   

The challenge is to create the absolute best, most competitive commercial rocket design possible that would be able to dethrone all current & near-future competitors (like possibly the Long March 5 & Angara series).  The best part is there are a huge number of design features to consider and plenty of different possible payload markets to target.  You can base your rocket company anywhere reasonable in the world (yes to Brazil, Russia, Europe, US, China, India, etc--no to places like Rwanda) in case that makes a key difference to your business case.

The categories to consider would include things like payload range, thrust range, target market, fairing sizes, engine number & type, number of stages, transportability/core size, man-rating (yes/no), propellants, modularity, engine-out capability, avionics, and so on.  This thread is flexible. 

Edit--Concern duly noted and acted upon. 

[Silmfeanor]

This topic is too broad, there are too many factors. You might want to add the word "current" in there somewhere. Or you're gonna get weird answers:

I'll go with a few examples:

fully reusable SSTO / TSTO.  High flight rate. Lifting body or with wings, aircraft like operations. Think Skylon. Doesn't work at all if you havent got the flight rate, or not enough money to finish design + building.

Hydrogen gun. Only works on propellant launches. can fire quite a few times per day. Not done before.

Simple RP-1 LOX 2 stage to orbit, common engine design. non-reusable. Basically falcon 9. Works in current environment, if it can be made reliably. RP can be changed to methane.
Perhaps it can be made reusable. Perhaps making it reusable is not economical, or not workable.

As I said - too many options, too many unknowns.

[Hyperion5]

 I thought about what my own ultimate commercial rocket design would look like.  Looking at what Musk & ULA have said, I think the two keys to success are low cost per kg to orbit & reliability. 

To get the cost per kg down, my design would be done in-house like Spacex does.  My initial choice of home country was the US, even if ITAR is not so helpful at times.  For the design, I wanted to choose a propellant mix that would be easy to handle, extremely cheap, and benefit the design.  My propellant choice was made when I saw Strangequark talking up the advantages of staged combustion methane engines.   

Centrifugal pumps are basically constant volumetric flow rate devices. A methox engine’s optimum O/F (~3) is pretty close to the equal volume ratio between O2 and CH4 (2.75). This means the pumps will have nearly identical requirements, and running them off a single spindle should be easy. In addition your turbine pressure drop will be low, so you have a simple 1-stage turbine. Then, your preburner is fuel-rich, running at the sweet spot temp for nickel superalloys (which might as well be brass in this industry), with a “non-coking” fuel. It could be a beautifully simple engine, with excellent sea-level and vacuum characteristics, a very respectable density Isp, and killer T/W.

After deciding on an entirely metholox rocket powered by Staged Combustion-cycle engines I made a few more choices to up reliability.  I wound up targeting my rocket at the large geosynchronous comsat market where it faces competition from the Ariane 5, Proton M, Zenit 3SL & soon from the Falcon Heavy, Long March 5 & Angara rocket families.  A lot of competition means there should also be quite a few payloads and my rocket will be able to prove itself against plenty of competitors. 

Here's what I came up with:

Rocket Name: Neptune V
Length: 60 meters
Diameter: 5.2 meters
Mass: 636 mt (1,400,000 lbs) (Imperial numbers rounded)
Thrust at liftoff: 909 mt (2,000,000 lbf)
Estimated LEO Payload: 19 mt (with margin to spare)
Estimated GTO Payload: 8 mt
SI: Five Triton I engines (400klbf, 2500 Psi, 35 expansion ratio, SC engines) arranged in Saturn V-like S-IC format
Estimated performance: 320 sec SL Isp
SII: Five Triton II engines: (50klbf, 2500 Psi, high expansion ratio, SC engines) arranged in Saturn V-like S-IC format
Estimated performance: 394 sec Vac Isp

The rocket has transportability issues, hence the factory has sea-side access and all launch sites would be sea-accessible.  This also is part of the reason for calling it "Neptune V" (the other being methane).  It's designed with blowout panels and kevlar lining to protect the other engines and propellant tank against bad engine failures.  Unlike current launchers, it would have engine-out capability on both stages at T+.1, giving potentially very good reliability.  Given how many times single-engine stages failed, I made sure this could handle an engine problem or two.  The rocket itself would be aluminum-lithium alloy, with common tanks for CH4 & Lox & also a common bulkhead between them to save weight and potentially expense. 

This launcher would be able to lift the biggest commercial comsats in existence.  It would be initially capable of 5 restarts with its SII engines and then 10 restarts on later versions for direct GEO insertion.  I'd equip it with fairings that would allow multiple missions for maximum flexibility.  It would be man-rated to enable potential future revenue from companies like Bigelow or their customers.  The eventual long-term goal would be reusability (hence the engine layout), and given the size of this thing, it'd be capable of carrying much more useful payloads to orbit than a fully reusable Falcon 9.  Eventually, I'd fly a Neptune V Heavy (3 cores) to compete against the Falcon Heavy straight on.  For anything heavier than that there's the 5-core version, the Neptune V Ultra. 

[neilh]

What's the ultimate commercial aircraft design?

[Rabidpanda]

Estimated performance: 394 sec Vac Isp

That seems way too high for methane.

[Hyperion5]

Estimated performance: 394 sec Vac Isp

That seems way too high for methane.

Well that's at 98% efficiency.  Anything lower and you'll be seeing your engine Isp dropping into the 380s. 

This topic is too broad, there are too many factors. You might want to add the word "current" in there somewhere. Or you're gonna get weird answers:

I'll go with a few examples:

fully reusable SSTO / TSTO.  High flight rate. Lifting body or with wings, aircraft like operations. Think Skylon. Doesn't work at all if you havent got the flight rate, or not enough money to finish design + building.

I've heard it repeated often that SSTOs are less than ideal launchers and really hard to pull off.  What are the odds you'd put on any fully reusable SSTO emerging as a commercial competitor in the launch market? 

Hydrogen gun. Only works on propellant launches. can fire quite a few times per day. Not done before.

You're right, that is a strange possibility.  Not something I'd ordinarily associate with being a possibility for a top commercial launcher. 

Simple RP-1 LOX 2 stage to orbit, common engine design. non-reusable. Basically falcon 9. Works in current environment, if it can be made reliably. RP can be changed to methane.
Perhaps it can be made reusable. Perhaps making it reusable is not economical, or not workable.

If you were designing something with today's tech, would you go this route with your launcher/launch vehicle? 

What's the ultimate commercial aircraft design?

I was busy driving for three hours and didn't get a chance to correct anything till now.  I've taken care of that issue.  It's now up to people to put forward what they think are the excellent designs meant to serve various payload markets (LEO, GTO/GEO, etc). 

[Tcommon]

Fully reusable, fast turnaround, lifting body. TSTO with identical stages - one carries extra tanks, the other carries payload +booster (if necessary). Propellant TBD. Something like this;

[Rabidpanda]

A rocket powered VTOL SSTO RLV fueled by liquid propane or LNG that can put 5-10 tons into LEO.

[Rabidpanda]

Estimated performance: 394 sec Vac Isp

That seems way too high for methane.

Well that's at 98% efficiency.  Anything lower and you'll be seeing your engine Isp dropping into the 380s. 

What's your source for that?  The highest number I've seen people give methane/lox is around 380 seconds.

[Hyperion5]

Estimated performance: 394 sec Vac Isp

That seems way too high for methane.

Well that's at 98% efficiency.  Anything lower and you'll be seeing your engine Isp dropping into the 380s. 

What's your source for that?  The highest number I've seen people give methane/lox is around 380 seconds.

The only reason you haven't seen better than 380 seconds vac Isp is the fact that you've never seen someone try out a metholox engine with 2500 Psi/172.37 Bar chamber pressure before.  380 seconds is on the low end of what staged combustion upper stages engines can do.  All you have to do is up the chamber pressure and have an adequate expansion ratio. 

http://www.astronautix.com/engines/rd160.htm

RD-160

Engine: 129 kg (284 lb). Chamber Pressure: 118.00 bar. Area Ratio: 352. Thrust to Weight Ratio: 15.5. Oxidizer to Fuel Ratio: 3.69.

Status: Developed 1993-.
Unfuelled mass: 129 kg (284 lb).
Height: 1.70 m (5.50 ft).
Diameter: 0.76 m (2.51 ft).
Thrust: 19.60 kN (4,406 lbf).
Specific impulse: 381 s.
Burn time: 900 s.
First Launch: 1993-.

The Russians hit 381 seconds of Vac Isp with this 1711 Psi metholox engine.  The ones I'd be using would have some 46% more chamber pressure than the RD-160.  I did the numbers on the jump in performance between the 2103 Psi NK-33 to the 3746 Psi RD-191, and they come out to a 16.5% chamber pressure increase per 1% jump in Isp.  Mind you, since we're coming up from an even lower pressure, the curve will actually be better than that. 

I figure a conservative estimate of 16% chamber pressure increase yielding a 1% Isp vac increase would see our engines' Isp jump 2.875%.  That yields 391 seconds of Isp, a number that stands right next to my figures.  If you use more realistic ratios of chamber pressure increase:Vac Isp increase, you should have no problems hitting that number.  I had modemeagle run over the numbers with his engine simulator to be sure, and the "Triton IIs" hit these figures with a lower expansion ratio than that featured on the RD-160.  I could have hit 400 seconds Vac Isp had I upped the pressure further. 

[Andrew_W]

If you'd asked me this twenty years ago I'd have shown you a sketch for a parallel burn LH2/LOX launcher, lift-off weight 360 tonnes, boosters and core 5 meters in diameter, the core stage powered by 1 SSME performance with each of 2 boosters powered by a pair of similar engines, I figured about 15-18 tonnes to LEO. The boosters lowered for recovery under paragliders, the core main engine housed in a up-side-down reentry capsule for recovery, the core stage tank going into orbit to be of future use.

Now I'd advocate subsonic air-launch with various orbital upper stages:
 1. A 140 tonne LH2/LOX space-plane with a 5 tonne, 8 person emergency re-enty capsule sitting semi-recessed into the LOX tank as with a fighter plane cockpit, powered by a pair of J2X or similar.
 2. An unmanned version to act as a fuel truck.
 3. An unmanned version with the capsule replaced with a PL fairing.

I see a natural progression to larger faster carrier aircraft, with the system evolving to using an orbital tether to catch the space-plane.

[Hyperion5]

If you'd asked me this twenty years ago I'd have shown you a sketch for a parallel burn LH2/LOX launcher, lift-off weight 360 tonnes, boosters and core 5 meters in diameter, the core stage powered by 1 SSME performance with each of 2 boosters powered by a pair of similar engines, I figured about 15-18 tonnes to LEO. The boosters lowered for recovery under paragliders, the core main engine housed in a up-side-down reentry capsule for recovery, the core stage tank going into orbit to be of future use.

Now I'd advocate subsonic air-launch with various orbital upper stages:
 1. A 140 tonne LH2/LOX space-plane with a 5 tonne, 8 person emergency re-enty capsule sitting semi-recessed into the LOX tank as with a fighter plane cockpit.
 2. An unmanned version to act as a fuel truck.
 3. An unmanned version with the capsule replaced with a PL fairing.

I see a natural progression to larger faster carrier aircraft, with the system evolving to using an orbital tether to catch the space-plane.

Ah, so a bit like the Raptor-powered spaceplane I once proposed but with liquid hydrogen.  This sounds sort of like a cross between the successor to the Lynx and a next-generation Virgin Galactic spaceplane.  XCOR I've heard is looking into a more powerful LH2/Lox engine for a possible follow-up to the Lynx.  Your design certainly would look interesting to the people at XCOR & Virgin Galactic. 

-----

I know my launcher looks rather conventional against everything else proposed so far, but it'd also be an unconventional LV.  No one, to my knowledge, has ever built an entirely metholox carrier rocket of truly commercial scale.  Also to my knowledge, no one's copied the engine number and layout of a 2-stage Saturn V since the original, nor had engine-out capability on all stages since then. 

Basically the Neptune V family would be like a cross between the Angara family, a 2-stage Saturn V, and the mini-Saturn V that results would burn methane. The Angara series inspired the overall family:

Neptune V
LEO: 19 mt (est.)
GTO: 8 mt (est.)

Neptune V Heavy (3 core)
LEO: 64.6 mt (est.)
GTO: 20 mt (est.)

Neptune V Super Heavy (4 core)
LEO: 84 mt (est.)
GTO: 28 mt (est.)

Neptune V Ultra (5 core)
LEO: 100 mt (est.)
GTO: 35 mt (est.)

Bear in mind my GTO numbers are very conservative. 

The family could compete with just about every major launcher bigger than a Vega.  The Neptune V would have multi-payload capability to take on the Soyuz and Falcon 9, while it could also compete with the Zenit 3SL, Proton M, the Angara A3, several Long March 5 variants, and the Ariane 5 for big geosynchronous satellites.  To compete with the Angara A7V & Falcon Heavy, the Neptune V Heavy would be more than adequate.  It should be able to put a Bigelow BA 330 into geosynchronous orbit, enabling the first GEO space station.  By the time you get to the 25-engine Neptune V Ultra, you could enable missions to the moon.  That's basically one launcher family capable of competing for payloads ranging from 8 mt to 100 mt.  Even the Angara family won't be that wide-ranging. 

[Robotbeat]

Reusable and cheap enough to fly every day. It'd become ultra reliable and cheap.

[sdsds]

I second Robotbeat's suggestion regarding the flight rate. At a minimum, it needs to be a question of "how many flights per month" rather than "how many months between flights."

I think that requires highly automated inspection between flights; perhaps running a "cold flow" diagnostics check and maybe taking a boroscopic peek at likely trouble spots such as engine throats. Reflight without disassembly should be the goal.

[strangequark]

Reusable and cheap enough to fly every day. It'd become ultra reliable and cheap.

Indeed, though I'd add a change to partially reusable. Give me a few billion, and I think I would pursue a TSTO with a reusable methane first stage that does as much of the delta V as is practical, and then a disposable, mass-produced pressure fed second stage.

[guckyfan]

Reusable and cheap enough to fly every day. It'd become ultra reliable and cheap.

Indeed, though I'd add a change to partially reusable. Give me a few billion, and I think I would pursue a TSTO with a reusable methane first stage that does as much of the delta V as is practical, and then a disposable, mass-produced pressure fed second stage.

I like that approach, as it seems certainly achieavable and can reduce cost a lot. But as higher ISP is very efficient on an upper stage, a turbo pump may be worth the extra cost and give less $/kg.

[beb]

I'm not a rocket scientist so my thoughts will have to remain vague.

The "ultimate commercial" rocket design would be one that gets you into space the cheapest so I think what you're seeing coming from SpaceX is about as close to the Ultimate Commercial rocket design as one can get. Still there is roo0m for some thought.

1.    I'd call for a single engine that can be used across a wide spectrum of applications. By that I mean one engine, kerolox, of, say, a half-million pounds thrust. A single engine would give you Delta II performance, two engines would equal EELV performance and four engines would get you into EELV-Heavy performance. Then forming a heavy version of this four engine rocket would get you a 12-engine, 6 million pound HLV launcher.  And I'd go with a gas generator engine sacrificing some ISP for a simpler, cheaper design.

2. I'd call for a moduler first stage tank design. I liked the ACES concept of adding babrrel sections to the tanks so that one line can produce small, medium or large stages. I'm not sure whether a common stage for 1 and 2 engine vehicles makes more sense or a common stage for 2 and 4 engine rockets. Maybe it's possible to remove enough  barrel sections to size the stage for 1, 2 and 4 engines. That would make the single engine version very stubby but I've often wondered how stubby a rocket could be before it becomes impractical to launch.

3. The upper stage would be an ACES hydrolox unit. As again it would use one engine in different numbers for the different mission needs and the same tankage, in different length.

4.    Standardized attachment points so that one could use the single engine version as a booster to the 2 or 4 engine launch vehicles to more closely match launch needs with launch capacities.

Most of this is just Atlas V Phase II. It always struck me as a very sound idea. The chief difference is replacing the Russian made engine with a domestic variant

[Jim]

What's the ultimate commercial aircraft design?

bingo

[simonbp]

What's the ultimate commercial aircraft design?

bingo

Long cylindrical fuselage, swept wings with complex control surfaces and winglets, high-bypass turbofan engines in nacelles either on the wing or tail, dense seating in the cabin, all facing forward, APU in the tail to provide ground power. Next time you are at the airport, I dare you to find a commercial passenger aircraft that doesn't follow this description.

So, the point is not "ultimate" as in final forever, but rather the style of design that many companies will converge on after many years of operation.

[Silmfeanor]

So, the point is not "ultimate" as in final forever, but rather the style of design that many companies will converge on after many years of operation.

I think that's the point to be made in that discussion. Rocketry is nowehere near this state.

Also, how about a F-22 raptor? An UAV? Weather/spying high altitude balloons? VTOLS? B-2?

[Hyperion5]

So, the point is not "ultimate" as in final forever, but rather the style of design that many companies will converge on after many years of operation.

I think that's the point to be made in that discussion. Rocketry is nowehere near this state.

Also, how about a F-22 raptor? An UAV? Weather/spying high altitude balloons? VTOLS? B-2?

F-22 Raptor is a military vehicle, as are many UAVs, weather and spy balloons rarely have the kind of commercial utility of airliners, VTOL aircraft so far are mostly the province of the military, and we all know there's little commercial utility to a B-2. 

Rocketry is beginning to move towards a certain number of more standard features.  Take the number of stages for example.  When the USSR built the Sputnik-PS from the R-7, it had only one stage.  Within a few short years they were building Proton rockets with three and four stages.  The N-1 rocket had no fewer than five stages.  Atlas rockets were famous for their relatively unique staging.  In the early years there was relatively little consensus on the number of stages a rocket should have. 

Look at the situation today.  The incoming Angara & Long March 5 rocket families have a maximum of 3 serial stages, and will often fly with just two.  Adding in boosters makes them 2.5 & 3.5 stage rockets.  The Ariane 5 is a 2.5 stage rocket, as will be the Falcon Heavy, the Space Launch System, and if the Russians built the Sudrozhestvo (a 5-Zenit core HLV), that as well.  Both the Delta IV & Atlas V are 2 stage rockets in their base format, and 2.5 stage rockets with added cores as boosters.  If I had a guess as to why we're settling on 2-3.5 stage rockets as our common range, I'd guess it would be to minimize serial staging risk.  All one has to do is compare the reliability of the Soyuz family vs the Proton family over the years to see some of this risk in action.

[Andrew_W]

What's the ultimate commercial aircraft design?

bingo

Long cylindrical fuselage, swept wings with complex control surfaces and winglets, high-bypass turbofan engines in nacelles either on the wing or tail, dense seating in the cabin, all facing forward, APU in the tail to provide ground power. Next time you are at the airport, I dare you to find a commercial passenger aircraft that doesn't follow this description.

So, the point is not "ultimate" as in final forever, but rather the style of design that many companies will converge on after many years of operation.

Well said.

[Rugoz]

You should be more specific, like

The cheapest/most reliable launcher for...

"20 6t comsats into GTO orbit per year"

or

"4 20t payloads into LEO per year"

The ultimate commercial rocket design depends on market conditions, just like the "ultimate commercial airplane" design.

[watermod]

Tongue totally in cheek and wearing an evil Ming the Merciless hat...
 
It starts with a business plan that removes competition...
so...
Start with spinoff's of the original Project Orion launching as near as possible to your main competitor's factories. I should probably go up on one of the Orion launches to avoid the lawsuits and arrest.

After the radiation takes care of the  competition ... any kind of rocket for normal launches I wanted...
 
Sorry - this question was just too wide open.

[RanulfC]

As one of the folks whom Hyperion5 ran this thread-concept by for "checking" before he posted I apologize (a lot!) for missing the fact it could be seen as too "broad" a subject

However I'll take a "stab" at answering both the questions of:
What would your ideal commercial rocket designs look like?
What's the ultimate commercial aircraft design?

Since I've recently run across a company that is trying for both
(Ok, technically TransPower is ONLY the lead on one segment and an overall "management" group for the project. However their "partner" and "lead" company for "Application Analysis" including research, design and evaluation of concepts and systems is Boeing

How about a "Nuclear Hybrid Space Plane"?
http://www.transpowerusa.com/wordpress/space-transportation/commercial-space-plane/

I'm kind of "impressed" that their Operational Goals for the vehicle extend all the way out to Jupiter:
http://www.transpowerusa.com/wordpress/space-transportation/commercial-space-plane/operating-goals/

The idea is to use a "core" of Fusion Plasma (Gasdynamic Mirror, GDM) to produce neutrons to bombard a "blanket" of Thorium-232 which transmutes to Uranium-233 and undergoes fission. The thermal output would then be applied to a reaction-mass and used in Combined Cycle Engines in both Air-Breathing and Rocket modes to achieve space flight.

Nice work, if you can get it

Randy

[Elmar Moelzer]

At a very brief first look, I like the concept. I will look at it some more when I have the time. One of the problems with conventional nuclear thermal propulsion is that small amounts of nuclear fuel and fission products would escape with the exhaust gases. Not sure whether that is a problem with their system too, but it might give some people reason for concern.

[Hyperion5]

As one of the folks whom Hyperion5 ran this thread-concept by for "checking" before he posted I apologize (a lot!) for missing the fact it could be seen as too "broad" a subject

That's alright, Randy.  It's kind of hard to foresee everything, as people in my field know all too well. 

However I'll take a "stab" at answering both the questions of:
What would your ideal commercial rocket designs look like?
What's the ultimate commercial aircraft design?

Since I've recently run across a company that is trying for both
(Ok, technically TransPower is ONLY the lead on one segment and an overall "management" group for the project. However their "partner" and "lead" company for "Application Analysis" including research, design and evaluation of concepts and systems is Boeing

How about a "Nuclear Hybrid Space Plane"?
http://www.transpowerusa.com/wordpress/space-transportation/commercial-space-plane/

I'm kind of "impressed" that their Operational Goals for the vehicle extend all the way out to Jupiter:
http://www.transpowerusa.com/wordpress/space-transportation/commercial-space-plane/operating-goals/

The idea is to use a "core" of Fusion Plasma (Gasdynamic Mirror, GDM) to produce neutrons to bombard a "blanket" of Thorium-232 which transmutes to Uranium-233 and undergoes fission. The thermal output would then be applied to a reaction-mass and used in Combined Cycle Engines in both Air-Breathing and Rocket modes to achieve space flight.

Nice work, if you can get it

Randy

This design concept would almost be like suddenly going from Apollo to launching into orbit like the Millennium Falcon does in Star Wars.  If they could make it work it'd revolutionize the field and make the colonization of the moon and Mars a relative cake-walk to what it is today. 

I'll say this for my design concept of the Neptune V.  It may not be as revolutionary as that, but it's basically taking current design trends and pushing them to the extreme.  It's got common tanks on both stages like a Falcon 9, it improves its dry mass fraction by using common bulkheads on both stages (like on Saturn V's S-II), and this is all thanks to its uses a cheap easy-to-handle propellant mix (not dissimilar to Soyuz & Falcon 9).  That cuts construction and handling costs considerably. 

To max out reliability I decided it needed multi-stage engine-out capability, which is enabled by five engines per stage.  That means you're using the same quantity of engines per launch as a Falcon 9 (though half of them are much smaller) but since there are only five in action at a time, the stage failure rate is likely to be quite low if you can get the engines reliable.  To maximize reliability and minimize serial staging risk it only has two stages (like Falcon 9/Atlas V/Delta IV).  Also, due to the engine count per core, you can make this thing extremely modular like the Angara family.  1, 3, 4 and 5 core versions are all possibilities, with even the 25 engines lifting the Neptune V Ultra 2 fewer than the Falcon Heavy. 

To maximize performance it would use lithium-aluminum alloys to save weight.  Its engines are also basically cousins of each other, with the same chamber pressures (2900 Psi), and both engine types are high-efficiency staged combustion metholox engines.  They're a bit overkill for rockets of this size, but the reason for that is I want to still be able to send up useful payloads on fully reusable versions.  That's enabled by the x layout of the five engines on each stage, so once you can get a Grasshopper-style program running, these things ought to be able to match Elon's Falcon 9. 

The only issue right now is I'm trying to figure out what size Neptune V to consider.  I'm currently considering 3 versions varying in power weighing 478 mt, 510 mt and 636 mt respectively.  The smallest would be an excellent Falcon 9/Zenit 3 competitor, while the largest would be taking on the Ariane 5, Long March 5, Delta IV Heavy & Angara A5. 

[RanulfC]

At a very brief first look, I like the concept. I will look at it some more when I have the time. One of the problems with conventional nuclear thermal propulsion is that small amounts of nuclear fuel and fission products would escape with the exhaust gases. Not sure whether that is a problem with their system too, but it might give some people reason for concern.

Actually by the end of the NERVA program there was almost no measurable product escape even after being run for hours. The system described in the link would either use heat-exchanger system or run liquid thorium in a heat exchanger to move the heat to the exhaust gas.

Randy

[Hyperion5]

At a very brief first look, I like the concept. I will look at it some more when I have the time. One of the problems with conventional nuclear thermal propulsion is that small amounts of nuclear fuel and fission products would escape with the exhaust gases. Not sure whether that is a problem with their system too, but it might give some people reason for concern.

Actually by the end of the NERVA program there was almost no measurable product escape even after being run for hours. The system described in the link would either use heat-exchanger system or run liquid thorium in a heat exchanger to move the heat to the exhaust gas.

Randy

So the question then becomes, can you make a nuclear thermal rocket-powered LV politically acceptable?  If you can realize the design, you're only halfway to commercial rocket heaven.  Imagine you're the program head of this hybrid nuclear rocket spaceplane project and the design is finally getting ready for powered flight.  Your major problem is the Sierra Club and a huge crowd of anti-nuclear activists from across the world have descended on Cape Canaveral to try to stop you.  How do you win the PR war with these people?  Security will stop them from preventing the flight, but how do you deal with the Congressman demanding this sort of vehicle be outlawed from the states?  I would think you'd want to be able to fly from the major continents, not out of isolated islands. 

[RanulfC]

So the question then becomes, can you make a nuclear thermal rocket-powered LV politically acceptable?  If you can realize the design, you're only halfway to commercial rocket heaven.  Imagine you're the program head of this hybrid nuclear rocket spaceplane project and the design is finally getting ready for powered flight.  Your major problem is the Sierra Club and a huge crowd of anti-nuclear activists from across the world have descended on Cape Canaveral to try to stop you.  How do you win the PR war with these people?  Security will stop them from preventing the flight, but how do you deal with the Congressman demanding this sort of vehicle be outlawed from the states?  I would think you'd want to be able to fly from the major continents, not out of isolated islands. 

Not to get to far "off-topic" but it IS the question and the answer is actually pretty easy IF one "assumes" a reasonable world

Public education and outreach is a big factor here because you won't HAVE "huge-crowds" of anti-nuclear activists if you are open and honest in providing the public information because the majority of "anti-nuclear" activists tend to be misinformed at the least. Part of that education and outreach would of course have to target politicians as well, due to their positions. Though in truth all the education and honesty in the world won't overcome some of the "concerns" and those will have to be addressed in design or operations planning. Even then there will always be a "hard-core" segment that will never be convinced that anything "nuclear" can ever be safe in any form or fashion and there will have to be considerations made of their effect as well.

In the end it might JUST be "better" all around to at least start flying from a remote Pacific island until a decent flight record is built up and public fears begin to fade. Part of what makes this more likely in a Thorium scenerio than most others is the non-proliferation aspect of Thorium over (say) Uranium or Plutonium reactors. Thorium would "allow" some face-saving political manuevering to permit many "first-world" nations to support putting thorium reactors, processing and even "spaceport" facilities into some equatorial countries without seeming to support nuclear proliferation or weapons development yet still maintaining "control" and being part of the process for profit and political purposes.
(Of course in the long run such a "policy" and attitude could easily come back to bite the "supporting" nations in the butt given the whole convayed attitude of "better your backyard than mine" but I for one have never noted a reall "long-view" political attitude for most nations in the fist place

In the end "I" feel the PR battle is winnable and in truth it has always seemed to me to be a battle for pro-nuclear advocates to "lose" rather than for anti-nuclear advocates to "win" in the public support arena. YMMV of course

Randy

[Hyperion5]

At a very brief first look, I like the concept. I will look at it some more when I have the time. One of the problems with conventional nuclear thermal propulsion is that small amounts of nuclear fuel and fission products would escape with the exhaust gases. Not sure whether that is a problem with their system too, but it might give some people reason for concern.

Actually by the end of the NERVA program there was almost no measurable product escape even after being run for hours. The system described in the link would either use heat-exchanger system or run liquid thorium in a heat exchanger to move the heat to the exhaust gas.

Randy

So what exactly are the major engineering challenges to turning this into reality?  Which challenge is larger, the engineering or paying for the vehicle's development?  I can see a business case for a nuclear hybrid rocket spaceplane, but how much upfront investment would it require?  If it's going to cost half of Bill Gates' 60 billion plus dollar fortune, then I can see why Boeing isn't exactly flinging resources at the problem. 

[edkyle99]

If I had a guess as to why we're settling on 2-3.5 stage rockets as our common range, I'd guess it would be to minimize serial staging risk.  All one has to do is compare the reliability of the Soyuz family vs the Proton family over the years to see some of this risk in action.

There are a couple of problems with this comparison. 

First, an R-7 rocket consists of six to seven propulsion modules, depending on mission, each of which must separate from something.  Proton only has four propulsion modules (stages) at most.  So R-7, the more reliable of the two, has more separation events than the less-reliable Proton system.

Second, staging accounts for a small percentage of failure causes.  Most failures are propulsion related.  Guidance/flight control also causes losses, as do what is probably the most annoying orbital launch failure - payload shroud separation issues.  I only show one staging related cause among all of the Proton failures over the years.

In my view, launch vehicle design is substantially driven by cost - both development and operating cost.  Minimizing the number of different propulsion system types (think Merlin or YF-22 or RD-107/108) cuts development costs.  High reliability helps to keep operating costs low.  A good design builds in reliability by not pushing margins too hard (see R-7).  A troublesome design pushes margins too much (see the Briz M failures on Proton).

As for this thread's open question, I believe that a good "commercial rocket" is more than just the rocket - it is also the launch site and the testing sites and all of the industrial infrastructure that supports the program. 

Assuming that we are talking about supporting a variety of payload types and orbits (which I believe we must), then I would nominate Kourou, or something like it, as a nearly-ideal launch site.  Kourou can, all by itself, support launches to nearly any orbital plane - something that no other launch site (except Sea Launch Odyssey and air launch Pegasus) can do.  Kourou also gains the velocity advantage of its near-equatorial position for GTO launches.

As for the rocket, or, rather, rocket family, I would nominate China's Chang Zheng, or something like it, as an example of a flexible, reliable option.  CZ comes in numerous variations, but all are based on the same basic propulsion system and launch system and flight control system, etc. China Academy of Launch Vehicle Technology (CALT) builds it all, and tests it all, providing strong engineering control for the program.  The CZ family can fly in just two stages to lift 2 tonnes to polar orbit, or can be augmented with strap on boosters (using nearly identical engines) to lift Shenzhou crew spacecraft weighing more than 8 tonnes to LEO, or can be topped by a high energy upper stage to lift nearly 6 tonnes to GTO.  CZ flies from relatively basic launch sites.  Like R-7, CZ is reliable (0.98 success rate in 115 launches since 2000) because it is generally a conservative design.  No other country in the world has anything like CZ that can cover such a large payload range using essentially one rocket.

 - Ed Kyle

[go4mars]

Well, I'd like one that can significantly expand the launch market through price and capability.  Easily reusable, scalable, and cheapest fuel (methane).  Aviation is something like a 700 billion dollar industry.  I'd aim to capture 1% of that, and would want to be able to provide worthwhile destinations and adventures in space. 

Can fly 1,2,3, or 5 1st stage cores per launch.  5 would be shaped like 8o8
Fully reusable, always crossfeeds when more than 1 core. 
1st stage cores are at least 20 meters in diameter. 

Gaseous combustion plume cushion for atmospheric re-entry, and propulsive landing (liquid combustion). 

It would have optionality to fly with no upper stage (just a payload on parallel staging to orbit), or 1 or 2 upper stages, also an option depending on mission/goals.  Option for the 3rd stage to be hydrogen (or nuclear if I'm allowed to suggest it). 

Upper stages would fly themselves into position in the stacking process, or be emplaced by 'quick-detach' rocket belts that have significant cosign losses and use their own fuel.  Cosign loses so the stage below doesn't get cooked.  Fast and precise.  Maybe neutral buoyancy dirigibles. 
Parallel staging might be generally preferred compared to stacking, because of ground handling challenges. 

There would be a semi-ballistic space plane that goes on as a top stage, and seats hundreds.  It would re-enter above the ocean hypersonically, then switch to electric power, cruising at 80 thousand feet or so.  Then slow down, and glide through the atmosphere toward an exciting high-g vertical propulsive landing, 30 miles or so from city limits.  Passengers then quickly transfer to very high speed vacuum tunnel trains to population hubs on the destination continent.  Refuel, relaunch.  Or if we demand to be more boring than that, a subsonic glide to touchdown at a conventional airport then taxi tens of miles back to a launch site (I might be straying off topic there). 

Anyway, back to the rocket:  It enables colonization of space, and fast(ish) large missions to the outer solar system.

It is sized such that it's largest iteration comfortably gets the ultra-high isp fusion system, (General Fusion's secret project: were such to exist) into LEO. 

Called the Zeus (Ian Douglas novel reference).

Slightly modified versions of the 1st stage cores will be launched mostly empty, as part of the payload, and used as a landing stage on Mars, the moon, or elsewhere; available for re-fuel and re-launch.  The Mars version might also have a big ballute that can double as a greenhouse or pressurized habitation structure (only needed for entry to Mars at interplanetary speeds and not required for point-to-point travel on Mars).

Part of the reason for large diameter is for radial artificial gravity within the transit vehicles and stations.  If first stage core is 20 m diameter, the payload, packed might be 30 meters.  Station, Inflated, might be 50+ meters. 

Ground handling and payload processing would be a considerable challenge at all stages. But not insurmountable. 

The End.

[Hyperion5]

If I had a guess as to why we're settling on 2-3.5 stage rockets as our common range, I'd guess it would be to minimize serial staging risk.  All one has to do is compare the reliability of the Soyuz family vs the Proton family over the years to see some of this risk in action.

There are a couple of problems with this comparison. 

First, an R-7 rocket consists of six to seven propulsion modules, depending on mission, each of which must separate from something.  Proton only has four propulsion modules (stages) at most.  So R-7, the more reliable of the two, has more separation events than the less-reliable Proton system.

A good point, but of those six-seven separation events, four of them are of the less dangerous parallel staging variety.  The overall point I should have clarified was that adding more serial stages, not parallel stages, adds serial staging risk but more importantly ups the chances of upper stage failures on the rocket.  Of the six failures and one partial failure of the R-7 family since 1992, four out of the seven have related to upper stage issues, with 3 out of those 4 relating to issues with the third stage. 

I've looked long and hard, and so far I haven't found an example of a Soyuz suffering a bad set of booster separations since 1992.  There was one LRB explosion in 2002, but that's different than a parallel staging going awry. 

Second, staging accounts for a small percentage of failure causes.  Most failures are propulsion related.  Guidance/flight control also causes losses, as do what is probably the most annoying orbital launch failure - payload shroud separation issues.  I only show one staging related cause among all of the Proton failures over the years.

In my view, launch vehicle design is substantially driven by cost - both development and operating cost.  Minimizing the number of different propulsion system types (think Merlin or YF-22 or RD-107/108) cuts development costs.  High reliability helps to keep operating costs low.  A good design builds in reliability by not pushing margins too hard (see R-7).  A troublesome design pushes margins too much (see the Briz M failures on Proton).

Again, while it's true the actual staging risk is not as bad as I feared, I was thinking more generally as well of upper stage failures. I looked through the Proton's flight record since 1992, and I'm counting some 14 failures and partial failures.  Of those 14 failures, a full 10 out of 14 of them concern something going wrong with the final stage, and those failures include issues with others besides Briz-M stages.  By cutting down to 2-3 serial stages, the Angara rocket family that will succeed the Proton family ought to see failure rates more like that of the R-7 family.  The Russians seemed to have learned the lesson that cutting serial stages improves reliability, and looking at the Proton's record I can see why.   

As for this thread's open question, I believe that a good "commercial rocket" is more than just the rocket - it is also the launch site and the testing sites and all of the industrial infrastructure that supports the program. 

Assuming that we are talking about supporting a variety of payload types and orbits (which I believe we must), then I would nominate Kourou, or something like it, as a nearly-ideal launch site.  Kourou can, all by itself, support launches to nearly any orbital plane - something that no other launch site (except Sea Launch Odyssey and air launch Pegasus) can do.  Kourou also gains the velocity advantage of its near-equatorial position for GTO launches.

Certainly a good location for launching all kinds of payloads to all sorts of orbits.  I certainly can't fault the ESA for setting up at the location. 

As for the rocket, or, rather, rocket family, I would nominate China's Chang Zheng, or something like it, as an example of a flexible, reliable option.  CZ comes in numerous variations, but all are based on the same basic propulsion system and launch system and flight control system, etc. China Academy of Launch Vehicle Technology (CALT) builds it all, and tests it all, providing strong engineering control for the program.  The CZ family can fly in just two stages to lift 2 tonnes to polar orbit, or can be augmented with strap on boosters (using nearly identical engines) to lift Shenzhou crew spacecraft weighing more than 8 tonnes to LEO, or can be topped by a high energy upper stage to lift nearly 6 tonnes to GTO.  CZ flies from relatively basic launch sites.  Like R-7, CZ is reliable (0.98 success rate in 115 launches since 2000) because it is generally a conservative design.  No other country in the world has anything like CZ that can cover such a large payload range using essentially one rocket.

 - Ed Kyle

I'm not a fan of the current CZ family, and it comes down to two reasons.  The first is the fuels that propel current Chang Zheng rockets.  They're toxic, hypergolic propellants that cause huge issues whenever there's a failure and they fall near populated areas.  If a Falcon 9 failed shortly after launch, it would not be leaving a horrific mess and launching many horrific vapors into the air.  The Chinese know first-hand how dangerous this is from the time when a CZ veered off-course and killed 100 people in the 1990s.  Those toxic propellants also add handling costs and complexities, which can hardly help on costs.  Given the current CZ family is about to be replaced by non-hypergolic LV families like the Long March 5 series, I'd say even the Chinese agree hypergolic fueled LVs are not good long-term solutions. 

My second reason has to do with versatility.  Sure the CZ family as it is right now is more flexible and uses a lot of common parts.  However, once Russia starts launching the Angara family, even the upcoming CZ families will pale in versatility & commonality by comparison.  The Angara 1.2 will launch only 3.7 mt, while an Angara 5/KVRB will launch up to 28.5 mt.  There's even talk of Angara 7 versions launching over 40 mt and perhaps as much as 50 mt, with all but the big Angara 7 versions launching from a common pad.  All versions, except perhaps the biggest proposed Angara 5 & 7 versions, would use a common URM, and all of them use a common and proven first stage engine in varying numbers.  They all use, with the exception of the Angara A5P, a common second stage and engine.  The only real variety is in third stage, where the choice is between the Briz-M and varying numbers of RD-0146 LH2 engines (equivalent to RL-10B-2). 

Sure the Angara family is not yet proven, but I would bet long odds on it being more ideal commercial rocket family than the Chang Zheng family as it is or will be in a few years.   

Edit--correcting my faulty memory on fatality count.  The articles still count 80 houses destroyed, which makes a fatality estimate of 100 entirely reasonable by third parties.  This is an authoritarian state we're talking about here, not France or the US.  There is a tendency amongst such states to cover up the full extent of the disaster, the best example being Chernobyl in the USSR.

[edkyle99]

I'm not a fan of the current CZ family, and it comes down to two reasons.  The first is the fuels that propel current Chang Zheng rockets.  They're toxic, hypergolic propellants that cause huge issues whenever there's a failure and they fall near populated areas.  If a Falcon 9 failed shortly after launch, it would not be leaving a horrific mess and launching many horrific vapors into the air.  The Chinese know first-hand how dangerous this is from the time when a CZ veered off-course and killed 400 people in the 1990s. 

There is no basis for that claim, and the number keeps growing as the years pass.  At the time, China's official media reported six dead and 57 injured.  A U.S. Congressional report estimated that the real number might have been 100 "or more", but also provided no basis for the claim. 

The CZ problem isn't the hypergolic propellants, it is the inland launch sites.  A big kerosene/LOX rocket falling on the town would have been just as devastating as a CZ-3B.  Hyper fueled rockets have been safely launched for years from sea-side launch sites adhering to strict launch criteria.  Ariane 4 flew from Kourou, safely and very successfully, for many years, as did Titan from the Cape and Vandenberg.  (Nearly every rocket launched carries some toxics, if not for the rocket almost certainly in the satellite).

Going to RP/LOX in place of hypergolics is a good idea because it simplifies fueling processes, which helps reduce costs.

Re: the serial staging discussion.  The problem with Proton, IMO, isn't that it has a fourth stage, it is that Briz M is being made to perform very long burns.  The propulsion system wasn't initially designed for such burns, which have been facilitated by the addition of a drop tank.  Longer burns mean longer heat-soaking periods for chambers and rotating machinery, which stresses everything.  The long burns are separated by even longer coast periods, pushing mission length to 9 hours or more. 

 - Ed Kyle

[Hyperion5]

I'm not a fan of the current CZ family, and it comes down to two reasons.  The first is the fuels that propel current Chang Zheng rockets.  They're toxic, hypergolic propellants that cause huge issues whenever there's a failure and they fall near populated areas.  If a Falcon 9 failed shortly after launch, it would not be leaving a horrific mess and launching many horrific vapors into the air.  The Chinese know first-hand how dangerous this is from the time when a CZ veered off-course and killed 100 people in the 1990s. 

There is no basis for that claim, and the number keeps growing as the years pass.  At the time, China's official media reported six dead and 57 injured.  A U.S. Congressional report estimated that the real number might have been 100 "or more", but also provided no basis for the claim. 

Faulty memory on that one, sorry.  I've edited it down to 100 people as a reasonable estimate given 80 homes were destroyed.  I still think you might not want to entirely rely upon official Chinese media for your figures on this.  The official press is not exactly noted for always telling the truth when it's awkward for the CCP. 

The CZ problem isn't the hypergolic propellants, it is the inland launch sites.  A big kerosene/LOX rocket falling on the town would have been just as devastating as a CZ-3B.  Hyper fueled rockets have been safely launched for years from sea-side launch sites adhering to strict launch criteria.  Ariane 4 flew from Kourou, safely and very successfully, for many years, as did Titan from the Cape and Vandenberg.  (Nearly every rocket launched carries some toxics, if not for the rocket almost certainly in the satellite).

Going to RP/LOX in place of hypergolics is a good idea because it simplifies fueling processes, which helps reduce costs.

That last line basically sums up my case for not using hypergolic fuels--it costs less and is safer.  This is precisely why the Chang Zheng family is not staying with hypergolic fuels.  They are not good fuels for commercially cost-competitive LV families nor are they fun to handle.  Hypergolic fuels are good for vehicles like ICBMs or if you want an all-weather launcher ready to lift off at any time.  For me, commercially competitive LV families ought to be man-rated and their handling should be relatively easy.  The last man-rated hypergolic LV from the US was the Titan II, and I think there's a good reason for that.  Besides, imagine trying to sell space tourists on hitching a ride on a Chang Zheng versus a Soyuz.  I think they'd be sold on the Soyuz from the moment you described what the propellants in a Chang Zheng could do to them. 

Re: the serial staging discussion.  The problem with Proton, IMO, isn't that it has a fourth stage, it is that Briz M is being made to perform very long burns.  The propulsion system wasn't initially designed for such burns, which have been facilitated by the addition of a drop tank.  Longer burns mean longer heat-soaking periods for chambers and rotating machinery, which stresses everything.  The long burns are separated by even longer coast periods, pushing mission length to 9 hours or more. 

 - Ed Kyle

Alright, but it's not just been the Briz-M that's been failing, or else the Proton's lifetime reliability would not be so much worse than the R-7 family's.  Just counting from 1992, I count a DM-2 upper stage failing in February 1996, a Block D-2 upper stage failing to restart in November 1996, a Proton K/DM3 upper stage failing one second into its second burn in December 1997, and a Block DM upper stage failure in November 2002.  That's just the final stage failures before the Briz-M came into regular service.  There's something wrong about the way the rocket's designed to go about its missions if it has this many problems with the upper stage.  Besides that, we have the example of the Soyuz failures since 1992.  Three-fourths of the upper stage failures happened on the third stage, not the second stage.  I don't think that's necessarily just a coincidence.     


I would bet almost anyone would not design a 4-stage rocket if they were trying to create their ultimate commercial rocket design from scratch.  I seriously doubt you can make a 4-stage rocket as reliable as one with 2-3 stages even if your overall build & design quality is excellent.  More stages generally equals more things that can go wrong.  I'm fairly sure your risk isn't going down when you tack on another stage and add yet another burn to the flight plan. 

[edkyle99]

Alright, but it's not just been the Briz-M that's been failing, or else the Proton's lifetime reliability would not be so much worse than the R-7 family's.  Just counting from 1992, I count a DM-2 upper stage failing in February 1996, a Block D-2 upper stage failing to restart in November 1996, a Proton K/DM3 upper stage failing one second into its second burn in December 1997, and a Block DM upper stage failure in November 2002.
...
I would bet almost anyone would not design a 4-stage rocket if they were trying to create their ultimate commercial rocket design from scratch.

Fewer stages obviously means fewer engines, which should mean less money.  However, I'm still not ready to agree that fewer stages automatically means better reliability.  I think that in real life other factors tend to outweigh the number-of-stages factor.

Consider that Proton M/Briz M currently sports about the same reliability record as Sea Launch Zenit and China's CZ-3 series - rockets that only have three stages.  (It also has a better record than two-stage Falcon 9, but it is too soon to really make any judgements given Falcon 9's low number of flights.)  Also consider that Proton with DM-2 or DM-2M upper stages proved more reliable than Proton with Briz M (and also more reliable than Ariane 5G). 

CZ itself is very reliable.  The CZ-4 series is essentially tied with Soyuz U and Soyuz FG at the top of the reliability list for active launch vehicles. 

 - Ed kyle

[Hyperion5]

Alright, but it's not just been the Briz-M that's been failing, or else the Proton's lifetime reliability would not be so much worse than the R-7 family's.  Just counting from 1992, I count a DM-2 upper stage failing in February 1996, a Block D-2 upper stage failing to restart in November 1996, a Proton K/DM3 upper stage failing one second into its second burn in December 1997, and a Block DM upper stage failure in November 2002.
...
I would bet almost anyone would not design a 4-stage rocket if they were trying to create their ultimate commercial rocket design from scratch.

Fewer stages obviously means fewer engines, which should mean less money.  However, I'm still not ready to agree that fewer stages automatically means better reliability.  I think that in real life other factors tend to outweigh the number-of-stages factor.

Consider that Proton M/Briz M currently sports about the same reliability record as Sea Launch Zenit and China's CZ-3 series - rockets that only have three stages.  (It also has a better record than two-stage Falcon 9, but it is too soon to really make any judgements given Falcon 9's low number of flights.)  Also consider that Proton with DM-2 or DM-2M upper stages proved more reliable than Proton with Briz M (and also more reliable than Ariane 5G). 

It may not be an automatic guarantor of better reliability to have fewer stages, but it surely doesn't help reliability to have more stages either.  Of the four failures & partial failures involving the upper stages of the R-7 family since 1992, three-fourths have involved a third stage, not the second stage.  I also have to point out the reliability record of the Ariane 5G is skewed because it was the initial vehicle of the Ariane 5 family.  They hadn't ironed out the design's initial issues by then, while the Proton by that time was an established design that had been launching for decades.  If you had compared the Proton's early years with that of the Ariane 5G's early years, the comparison would not be nearly as pleasant for the Proton. 

I actually talked to Jim a bit on the Zenit.  At the time I thought the reason why the Zenit's reliability was poor (but improving) was the quad-chamber design of its first stage engine.  I contrasted that with the ever-reliable RD-180, which ULA was considering mounting two per core of for a prospective Atlas V Phase 2.  I thought the reason why ULA wanted two RD-180 engines is because it would be more reliable due to their dual-chamber design than a single quad-chamer RD-170. Jim said the following about that theory:

Wrong.  The reason they use RD-180 for advanced designs is because they have a source for engine.

RD-17X problems have nothing to do with the number of chambers.  It has to do with manufacturing quality.

So basically if the build quality were higher for the RD-170 engines used on the Zenit, you wouldn't be seeing anywhere near the number of first stages exploding, which would improve the Zenit's reliability considerably.  Imagine if an enlarged Atlas V used an RD-170 engine.  I would guess it'd suffer very few first stage failures in contrast to the Zenit because the quality control would be higher.  I still don't think the Zenit is the most ideal LV family, but it is potentially very modular.  RKK Energia's proposed creating a 2.5 stage, 5-core Zenit called the Sodruzhestvo to be Russia's HLV in the future.  http://www.russianspaceweb.com/sodruzhestvo.html
 

CZ itself is very reliable.  The CZ-4 series is essentially tied with Soyuz U and Soyuz FG at the top of the reliability list for active launch vehicles. 

 - Ed kyle

You know, there's probably a way for the R-7 family to top the Chang Zheng family in reliability and have almost as good versatility as well.  My idea is you take the Soyuz 2.1V and Soyuz 3 concept and take the best elements from both for a new R-7 family. 

http://www.russianspaceweb.com/soyuz1_lv.html
http://www.russianspaceweb.com/soyuz3_lv.html

For the base version, you simply add on an advanced ACES-type upper stage with four RD-0146E engines to the Soyuz 2.1V.  You factor a T+.1 engine-out capability into the upper stage payload number and throttle back the stage.  For the next version, you throttle the upper stage up, add on a pair of NK-33-1 boosters and lengthen the upper stage propellant tanks.  The biggest version would feature the normal 4 boosters and have the longest upper stage and still have T+.1 engine-out redundancy factored in on the upper stage.  Something like this ought to be even more reliable than the existing R-7 family and more modular, though you'd want all versions to be able to lift off a single pad like the Angara 1.2-Angara 5/KVRB. 

I factored in engine-out capability because ULA has been pretty adamant what they think the reliability effects of having it: "Significantly, the dual and quad engine ACES variants incorporate engine-out.  Engine-out capability provides the single largest lever to improving system reliability (Figure 4). It is a key basis for ULA’s desire for multiple smaller (25Klb) class engines rather than a single larger upper stage engine."

According to ULA's math, you'd see a better than order of magnitude improvement in 2nd stage failure rates with the Atlas V.  I think given how similar the RD-0146 engine is it wouldn't be too hard to copy this with a more modern and less costly engine for the R-7 family.  A Soyuz family like this might even keep Musk up at night before the Falcon 9 achieves demonstrated reusability. 

Edit--
An alternative cheaper version would be four RD-0146U engines up top burning CH4/Lox (metholox version of RD-0146).  You'd give up some payload lift but cut your handling issues significantly plus not have to worry about boil-off as much prior to additional burns. 

[Hyperion5]

Alright, but it's not just been the Briz-M that's been failing, or else the Proton's lifetime reliability would not be so much worse than the R-7 family's.  Just counting from 1992, I count a DM-2 upper stage failing in February 1996, a Block D-2 upper stage failing to restart in November 1996, a Proton K/DM3 upper stage failing one second into its second burn in December 1997, and a Block DM upper stage failure in November 2002.
...
I would bet almost anyone would not design a 4-stage rocket if they were trying to create their ultimate commercial rocket design from scratch.

Fewer stages obviously means fewer engines, which should mean less money.  However, I'm still not ready to agree that fewer stages automatically means better reliability.  I think that in real life other factors tend to outweigh the number-of-stages factor.

Consider that Proton M/Briz M currently sports about the same reliability record as Sea Launch Zenit and China's CZ-3 series - rockets that only have three stages.  (It also has a better record than two-stage Falcon 9, but it is too soon to really make any judgements given Falcon 9's low number of flights.)  Also consider that Proton with DM-2 or DM-2M upper stages proved more reliable than Proton with Briz M (and also more reliable than Ariane 5G). 

CZ itself is very reliable.  The CZ-4 series is essentially tied with Soyuz U and Soyuz FG at the top of the reliability list for active launch vehicles. 

 - Ed kyle

I just thought of something when we're talking about reliability.  I saw this claim by Spacex on their Falcon 9 page:

"Falcon 9 has nine Merlin engines clustered together. This vehicle will be capable of sustaining an engine failure at any point in flight and still successfully completing its mission. This actually results in an even higher level of reliability than a single engine stage. The SpaceX nine engine architecture is an improved version of the architecture employed by the Saturn V and Saturn I rockets of the Apollo Program, which had flawless flight records despite losing engines on a number of missions."


While I've been impressed with the reliability of the Soyuz & Chang Zheng families, I've always wondered if they might benefit from engine-out capability.  Both the Saturn V & Saturn I/IB have engine-out capability on at least one stage, and both rocket families actually wound up using that capability.  It's true none of the Saturns were really commercial LVs, but I noticed something remarkable about their reliability record.  All of the Saturns that did lift off on flights intended to make orbit accomplished the task, just as Spacex claims, even if one Saturn V made an orbit that wasn't quite in the plans. 

Which begs the question, should an ideal commercial LV have engine-out capability?  I've run across literally dozens of instances where a single engine stage, most often an upper stage, fails, and thus a mission is doomed.  Surely that has a significant cost to launch providers even if they're cutting expense and complexity with single-engine stages.  Are the extra costs and complexity be worth it if the chance of engine failure costing the mission is significantly reduced? 

[Jim]

Falcon 9 can not sustain an engine failure during second stage flight, which is more than 2/3 of the time to orbit.

Also, Soyus can not sustain an engine out as there is only one engine per booster

[Hyperion5]

Falcon 9 can not sustain an engine failure during second stage flight, which is more than 2/3 of the time to orbit.

Also, Soyus can not sustain an engine out as there is only one engine per booster

I guess I should have made my point clearer, as I didn't mean the Soyuz had an engine-out capability.  With regards to Spacex, obviously they're trying their best PR spin on the Falcon 9's design, as it in my opinion is not an "improved version" of the Saturn family's design architecture. 

But what about the broader point?  Is the extra complexity of multiple engines per stage worth it if that stage can tolerate an engine failure?  Whenever I think of engine-out capability, I always think of Johnathan Barr and Bernard Kutter's Atlas V Phase 2 report showing 3 of 4 engine capability would reduce upper stage failures to almost 1/16th that of a single-engine Centaur (page 5 of http://www.ulalaunch.com/site/docs/publications/EELVPhase2_2010.pdf). 

[Jim]

Look at the PR spin surrounding the Falcon-1

[Hyperion5]

Well, I'd like one that can significantly expand the launch market through price and capability.  Easily reusable, scalable, and cheapest fuel (methane).  Aviation is something like a 700 billion dollar industry.  I'd aim to capture 1% of that, and would want to be able to provide worthwhile destinations and adventures in space. 

Can fly 1,2,3, or 5 1st stage cores per launch.  5 would be shaped like 8o8
Fully reusable, always crossfeeds when more than 1 core. 
1st stage cores are at least 20 meters in diameter. 

I've been trying to square this circle for awhile now, but I still can't figure out how you're going to make 20 m diameter rockets easily reusable.  You point out the difficulty of doing this straight off. 

Ground handling and payload processing would be a considerable challenge at all stages. But not insurmountable. 

The End.

I totally understand the colonization market is what you're trying to fill, but how can a rocket this size be both "easily reusable" while giving the ground crews such huge issues in moving it around or payload integration?  I get that you could make it reusable.  I'm just questioning whether anything this size could be "easily reusable".  I mean, the first stage alone on one of these Zeus rockets is going to dwarf an entire Saturn V.  Now imagine you've landed that first stage and now you've got to get it back on a crawler using a movable crane (because there's a chance it'd destroy any fixed crane landing at the boost-back pad).  It sounds like a herculean endeavor, even for someone with big equipment experience. 

It would have optionality to fly with no upper stage (just a payload on parallel staging to orbit), or 1 or 2 upper stages, also an option depending on mission/goals.  Option for the 3rd stage to be hydrogen (or nuclear if I'm allowed to suggest it). 

Part of the reason for large diameter is for radial artificial gravity within the transit vehicles and stations.  If first stage core is 20 m diameter, the payload, packed might be 30 meters.  Station, Inflated, might be 50+ meters.   

Alright, well I've got no problem with the 1.5 staging option.  In fact, given the Angara A5P is supposed to be a 1.5 stage vehicle, I'd say this is even more feasible with methalox engines.  The real issue with vehicles of this size is their size and power.  There was something like a 4.5 mile exclusion zone around the Saturn V rockets launching Apollo missions for very good reasons (see N-1 explosions).  So while you may be able to convince NASA to launch at Cape Canaveral or the ESA to let you launch at Kourou, that'd probably only be the single-stick versions.  Anything bigger than that you'd probably get pushback from the local community.  Just consider for a moment that the 300 decibels of sound from the Saturn V lifting off would be enough to kill you at distances of up to around 800 feet. 

That means you need to find a relatively isolated but still sizable island somewhere to launch.  Midway Atoll comes to mind (once housed 2000 residents), as do the islands just off the eastern coast of Puerto Rico.  Btw, fun speculation, but if you were to launch a 5-core Zeus rocket from Puerto Rico, I would guess you should be able to see it from as far away as Trinidad & Tobago in the south and the Bahamas and southern Florida in the north.  It'd be the first pan-Caribbean viewing experience of an artificial object's launch.   

[go4mars]

I totally understand the colonization market is what you're trying to fill

I would ideally strive for a business model that I felt had a chance of paying for it's development assuming only the terrestrial hypersonic passenger travel market.  Again, if you could get 1% of the global airline business, that's over $7 billion revenue each year.  The Martian wannabe's like me would be gravy.  Part of the reason for concentrating on effective "first" stage capability is because most of the work done by this thing will be throwing large passenger compartments across the world.  Initially ballistic, then zero g for a while, then a high and fast cruise phase, then relatively low and subsonic for final approach.  These passenger compartments could take many different forms, but I suspect would have electric battery powered systems for extremely fast cruising at 80000 feet or so, and before you got over populated areas, you'd get lower and subsonic.  There would be rockets with relatively small tanks for terminal descent vertical landing (or maybe horizontal landing).           

But yes.  These would also enable meaningful space colonization and tourism programs. 

how can a rocket this size be both "easily reusable" while giving the ground crews such huge issues in moving it around or payload integration?

Methane plays a part (no coking, cheap, etc.), but there are plenty of ways it could be conceivably devised.  I like the idea of landing with long legs, or on submerged risers, within a deep pool of fresh water (height adjusted so that the nozzles don't get quenched on landing (takeoff submerged "sea-bee style" is fine)).  This is also where you would take off from.  Payload integration would be a challenge, and would need special/new approaches.  The most common payloads might need to be constructed nearby (or possibly on barges).  It is possible that the payloads would fly themselves to the top of the rocket in a refuellable "propulsively landing powered clamshell", which is topped off just before launch.  An alternative is to have a "rocket belt" position the payload (or upper stages) on the rocket, then detach itself and fly up and away.  Either of these would need either significant shielding on top of the lower stage, or more likely would have significant cosine losses (a la superdraco orientation) so that the rocket plumes don't torch the stage directly below. 

Consider for a moment that the 300 decibels of sound from the Saturn V lifting off would be enough to kill you at distances of up to around 800 feet.

I think you mean 220 decibels at the source(correct me if I'm wrong), and diminishing pretty fast.  The last shuttle launch I saw (215 decibels at source?) was from KSC visitor's center (6 miles apparently) and it was not loud.  I could hardly hear it over the noise of the crowd.  I read somewhere that at 6 miles it's down to 90 decibels (I suspect measured when the wind was supremely calm).     I know it's a logarithmic scale, with 12-gauge shotgun blasts at 165.  Have a look at this chart to get a better idea of how it diminishes with distance (exponentially).  I would suggest that you are overstating the concern, though yes, it would be a factor to examine.  I missed Saturn V, but intend to be as close as practicable when FH lights up.  Guys who saw shuttle and Saturn V (similar lift-off thrust) generally seem to prefer the "throatier sound" from kerolox combustion.  I wonder how methane compares.

http://www.makeitlouder.com/Decibel%20Level%20Chart.txt 


The key idea here is creating something so capable and cheap, that it very meaningfully expands the market.  This is clearly NOT for just serving the current market a little better. 

[Hyperion5]

I totally understand the colonization market is what you're trying to fill

I would ideally strive for a business model that I felt had a chance of paying for it's development assuming only the terrestrial hypersonic passenger travel market.  Again, if you could get 1% of the global airline business, that's over $7 billion revenue each year.  The Martian wannabe's like me would be gravy.  Part of the reason for concentrating on effective "first" stage capability is because most of the work done by this thing will be throwing large passenger compartments across the world.  Initially ballistic, then zero g for a while, then a high and fast cruise phase, then relatively low and subsonic for final approach.  These passenger compartments could take many different forms, but I suspect would have electric battery powered systems for extremely fast cruising at 80000 feet or so, and before you got over populated areas, you'd get lower and subsonic.  There would be rockets with relatively small tanks for terminal descent vertical landing (or maybe horizontal landing).           

But yes.  These would also enable meaningful space colonization and tourism programs. 

Alright, but where are you going to get all these pads built?  The exclusion zones on rockets the size are going to be at least five miles around the pad in all directions.  That and noise zoning ordinances will eliminate launching near most cities.  Of course if you're launching from NYC to Singapore you probably won't mind a few hours getting there, but at lesser distances this would be a factor versus something like a supersonic jet. 

how can a rocket this size be both "easily reusable" while giving the ground crews such huge issues in moving it around or payload integration?

Methane plays a part (no coking, cheap, etc.), but there are plenty of ways it could be conceivably devised.  I like the idea of landing with long legs, or on submerged risers, within a deep pool of fresh water (height adjusted so that the nozzles don't get quenched on landing (takeoff submerged "sea-bee style" is fine)).  This is also where you would take off from.  Payload integration would be a challenge, and would need special/new approaches.  The most common payloads might need to be constructed nearby (or possibly on barges).  It is possible that the payloads would fly themselves to the top of the rocket in a refuellable "propulsively landing powered clamshell", which is topped off just before launch.  An alternative is to have a "rocket belt" position the payload (or upper stages) on the rocket, then detach itself and fly up and away.  Either of these would need either significant shielding on top of the lower stage, or more likely would have significant cosine losses (a la superdraco orientation) so that the rocket plumes don't torch the stage directly below.

I'm not sure I'd recommend the clamshell trick.  That sounds like it'd introduce more risk into this than you'd want.  You'll probably want as little risk in payload integration as possible.  Have you considered using a giant airship like the envisioned Aeroscraft ML86X?   

Consider for a moment that the 300 decibels of sound from the Saturn V lifting off would be enough to kill you at distances of up to around 800 feet.

I think you mean 220 decibels at the source(correct me if I'm wrong), and diminishing pretty fast.  The last shuttle launch I saw (215 decibels at source?) was from KSC visitor's center (6 miles apparently) and it was not loud.  I could hardly hear it over the noise of the crowd.  I read somewhere that at 6 miles it's down to 90 decibels (I suspect measured when the wind was supremely calm).     I know it's a logarithmic scale, with 12-gauge shotgun blasts at 165.  Have a look at this chart to get a better idea of how it diminishes with distance (exponentially).  I would suggest that you are overstating the concern, though yes, it would be a factor to examine.  I missed Saturn V, but intend to be as close as practicable when FH lights up.  Guys who saw shuttle and Saturn V (similar lift-off thrust) generally seem to prefer the "throatier sound" from kerolox combustion.  I wonder how methane compares.

http://www.makeitlouder.com/Decibel%20Level%20Chart.txt 


The key idea here is creating something so capable and cheap, that it very meaningfully expands the market.  This is clearly NOT for just serving the current market a little better. 

I'd always wondered if I'd remembered that old source correctly.  Still those kinds of decibels and the thrust involved suggest you're going to have to build some supremely sturdy launchpads, particularly because a launch failure, even with engine-out reliability and good quality control, would probably only be a matter of time.  So how do you make sure your pads could survive such an explosion?  I expect you could save any crew with a launch abort system at least. 

You were wondering what methane sounded like, so I found a video of a small methalox engine firing:



Methane engines have without doubt some of the most beautiful exhaust flames out there, with enough carbon that it's much more visible than a hydrolox engine's exhaust but only enough so that the flame burns a beautiful blue. 

[Kaputnik]

A question which I think might fit into this thread.

If we look at rocket motors/stages currently produced, which type would give the best impulse per $?

I would assume that some already-available unit produced for perhaps a military application would fit the bill. Maybe a missile, RATO, etc.

Any suggestions?

[Hyperion5]

A question which I think might fit into this thread.

If we look at rocket motors/stages currently produced, which type would give the best impulse per $?

I would assume that some already-available unit produced for perhaps a military application would fit the bill. Maybe a missile, RATO, etc.

Any suggestions?

Well Ed Kyle keeps talking up the Trident II D5 missile, which has had 143 successes in a row now.  Surely one of the ballistic missiles currently in use by Russia, US, China or the other major powers fits the bill.  I believe he said you could get about a metric ton into orbit with the Trident.  If you could super-size that missile to match the Soyuz you might be on the right track for commercial payloads.  The insurance premiums would be tiny compared to launching on a Zenit or Proton. 

You mentioned impulse/$.  I'm assuming you mean specific impulse/$?  There's also impulse density/$, which might be more important.  That seemed to be the approach that Spacex took.  They didn't aim to smash records on Isp/$, but they did do phenomenal on thrust/$ and quite well on impulse density/$.  I wonder how well any of those correlates with the cost per kg to orbit. 

[Robotbeat]

Impulse is different from specific impulse. Impulse is thrust times time and has units of Newton-seconds.

Also, Trident missiles aren't exactly cheap.

[Hyperion5]

Impulse is different from specific impulse. Impulse is thrust times time and has units of Newton-seconds.

Also, Trident missiles aren't exactly cheap.

Alright, but would you have some suggestions for him instead of the Tridents then? 

[Kaputnik]

I'm assuming you mean specific impulse/$?

No, I meant impulse/$.
i.e. the total propulsive power from the stage, divided by cost.

My thinking is that, if there is some off-the-shelf unit (I'm expecting it to be a small solid, I would be very surprised if it wasn't) then that could form an important part of a commercial MCD vehicle.

[R7]

Always liked the Rocket Company design a lot. High KISS factor. (But I think they needlessly reduced it by adding pilot to first stage)

http://www.amazon.com/exec/obidos/ASIN/1563476967/qid=1119462002/sr=11-1/ref=sr_11_1#reader_1563476967

[Hyperion5]

Always liked the Rocket Company design a lot. High KISS factor. (But I think they needlessly reduced it by adding pilot to first stage)

http://www.amazon.com/exec/obidos/ASIN/1563476967/qid=1119462002/sr=11-1/ref=sr_11_1#reader_1563476967

KISS factor?  That's a new one to me.  What does that stand for?

[kch]

Always liked the Rocket Company design a lot. High KISS factor. (But I think they needlessly reduced it by adding pilot to first stage)

http://www.amazon.com/exec/obidos/ASIN/1563476967/qid=1119462002/sr=11-1/ref=sr_11_1#reader_1563476967

KISS factor?  That's a new one to me.  What does that stand for?

http://en.wikipedia.org/wiki/KISS_principle

[Hyperion5]

Always liked the Rocket Company design a lot. High KISS factor. (But I think they needlessly reduced it by adding pilot to first stage)

http://www.amazon.com/exec/obidos/ASIN/1563476967/qid=1119462002/sr=11-1/ref=sr_11_1#reader_1563476967

KISS factor?  That's a new one to me.  What does that stand for?

http://en.wikipedia.org/wiki/KISS_principle

Ah, so it's about the principle of keeping things simple.  I've been wondering if that principle can sometimes run into problems when faced by things solved by built-in redundancy (extra complexity).  For instance, the Falcon 9 has triple-redundant avionics.  Is this an instance where the complexity that comes with redundancy and the reliability it can bring trumps simplicity? 

Or take engine-out capability.  The EELVs are quite reliable, but they've also suffered the occasional upper stage engine hiccup that's led to problems.  I remember seeing a ULA study that they'd see better than an order of magnitude drop in failure rate on a 4-engine upper stage compared to the current Centaur stage on the Atlas.  Wouldn't that be a case of where more complexity and redundancy would bring about better reliability?  I know it'd be extra engine expense, but the economies of scale would improve considerably if an Atlas were using 4X the upper stage engines it did previously. 

[R7]

Ah, so it's about the principle of keeping things simple.  I've been wondering if that principle can sometimes run into problems when faced by things solved by built-in redundancy (extra complexity).

Not necessarily, they can compliment each other. Design a system with KISS in mind, that usually yields better reliability alone. Then add redundancy, even better.

KISS in the Rocket Company vehicle showed for instance in first stage, IIRC it was pressure fed.

[Robotbeat]

Ah, so it's about the principle of keeping things simple.  I've been wondering if that principle can sometimes run into problems when faced by things solved by built-in redundancy (extra complexity).

Not necessarily, they can compliment each other. Design a system with KISS in mind, that usually yields better reliability alone. Then add redundancy, even better.

KISS in the Rocket Company vehicle showed for instance in first stage, IIRC it was pressure fed.

I think pressure fed makes sense a lot of times, but developing a pump has a LOT of advantages and shouldn't be discounted.

[go4mars]

Alright, but where are you going to get all these pads built?  The exclusion zones on rockets the size are going to be at least five miles around the pad in all directions.

A lot of the world's great cities are near bodies of water.  That's one option.  Another is high-speed transport, whether train, hyperloop, etc., to somewhere feasible and relatively nearby.  My friend's gps showed 347 km/hour on the Eurostar at one point for example.  Calgary might use Nose Hill Park. 

I'm not sure I'd recommend the clamshell trick... You'll probably want as little risk in payload integration as possible.  Have you considered using a giant airship like the envisioned Aeroscraft ML86X?   

If a giant airship existed and was available, feasible, and easily affordable, then it could be a great option.  But if you are building a powered clamshell anyways for other operational reasons, including perhaps abort contingencies and complete system reusability, then I think it makes sense to use it for stacking.  As to risk, the system would be tested many times first, in various wind conditions, etc.  If grasshopper and dragon, the Masten stuff, and other guys can control accurately enough...    then why not?

you're going to have to build some supremely sturdy launchpads, particularly because a launch failure, even with engine-out reliability and good quality control, would probably only be a matter of time.  So how do you make sure your pads could survive such an explosion?  I expect you could save any crew with a launch abort system at least. 

If it is necessary, build longish legs, and innundate the area with perhaps a meter of fresh water for launch/landing operations.  Though it might not be necessary.  A merlin doesn't seem to cut into the concrete.  I'm not sure how much larger this could be extrapolated to.  In any case, I suspect it will be trickier to deal with this issue on Mars or the moon for very big rockets. 

[R7]

developing a pump has a LOT of advantages and shouldn't be discounted.

And the RComp design didn't, but followed the minimum cost design ideology that lower the stage less need to aim for maximum efficiency in mass/Isp etc. Upper stage used ...  . .. . . .. .. either stock RL-10s or assumed RL-60, can't remember which. Don't have the book, read the story online back in 2005 or so. Unfortunately it isn't anymore available from web archive.

First stage recovery was solved very simply. It just went up and down, provided only enough vertical speed for the second stage to have time to horizontally accelerate into orbit.

[Robotbeat]

Pop-up first stage booster is a good idea.

[go4mars]

Pop-up first stage booster is a good idea.

Totally.  I suspect that's especially true for destinations above LEO. 

[Patchouli]

Pop-up first stage booster is a good idea.

It was one of the more valid points of the Kistler K1.


One thing I'm surprised is that no one has tried is resurrect the concept of the SASSTO.
http://www.astronautix.com/lvs/sassto.htm

It might be possible to build such a vehicle using a single SSME.

Even the ISP numbers are fairly close the SSME has better but slightly better sea level and slightly worse vacuum ISP numbers then the target set for the plug nozzle engine on the SASSTO.

359 sl and 459 vac for the plug nozzle vs 363 sl and 453 vac for the SSME.

The big short coming though it's payload is pretty small at only 3,600kg.
A modern version might be able to do 4,500kg IMLEO.
Though a few SRBs plus going expendable may double that.
Put it on something like a Saturn IB first stage and you get something with lift capacity similar to an EELV heavy.
The big short coming the SSME cannot air start easily and it's expensive which would make an expendable version very expensive.
Though if used with as an upper stage the SSME could be replaced with RL-10s or a J2 derivative.

DC-X though was fairly similar in concept.

[R7]

Pop-up first stage booster is a good idea.

It was one of the more valid points of the Kistler K1.

K1 didn't pop-up, it was supposed to pop-back exactly like F9R plans to (save the parachute landing vs powered).

[Hyperion5]

you're going to have to build some supremely sturdy launchpads, particularly because a launch failure, even with engine-out reliability and good quality control, would probably only be a matter of time.  So how do you make sure your pads could survive such an explosion?  I expect you could save any crew with a launch abort system at least. 

If it is necessary, build longish legs, and innundate the area with perhaps a meter of fresh water for launch/landing operations.  Though it might not be necessary.  A merlin doesn't seem to cut into the concrete.  I'm not sure how much larger this could be extrapolated to.  In any case, I suspect it will be trickier to deal with this issue on Mars or the moon for very big rockets. 

Actually a Merlin will burn straight through concrete, the trick being it has to be fixed to a test stand to do that.  Clearly judging from the Grasshopper tests it won't destroy concrete by landing on it (limited exhaust contact with concrete), but that's with one 147,000 lbf engine that's throttled back.  I'm not as confident that scaling up to an F-1 class engine will see this function as well.  That's why I think your idea about the water at the landing site is on the money. 

I believe you've already seen my PM on your Zeus probably producing 36,000,000 lbf of thrust from five 7.2 mlbf engines (assuming they fit on a 20 m core).  If those babies didn't burn through concrete I'd be surprised.  At the very least though I think the Zeus might actually benefit from a less efficient design.  I don't mean lowering chamber pressure, as that would not help its length/width ratio problem (somewhere around 2.8-3:1 without PLF).  However, if you ditched common bulkheads on both stages, I'd bet you could get the ratio up to 3.5-4 without a PLF, which is way more manageable.  You'd take a hit to payload, but hey, your Zeus will be topping 500 mt to LEO anyways.  It can afford to take a payload hit compared to most rockets.   

[R7]

All >100mt behemoths have common problem to be ultimate commercial rockets now or in near future. There are no payloads for them.

[Hyperion5]

All >100mt behemoths have common problem to be ultimate commercial rockets now or in near future. There are no payloads for them.

Yet.  I think the key would be getting more firms like Bigelow up and running.  If Bigelow could get some space stations into LEO, they might consider launching much bigger versions.  See attached for a 100 mt (65 mt empty) space station module almost twice the volume of the ISS . 

[go4mars]

All >100mt behemoths have common problem to be ultimate commercial rockets now or in near future. There are no payloads for them.

I agree with now.  But "near future" is undefined. If it existed, and was a lot cheaper per launch than the relatively tiny rockets of today, then my belief is that payloads would appear.  Especially payloads related to hypersonic terrestrial transportation (which would need to be planned in from the start of development), space tourism, and space colonization.  The only way to disprove my assumption is to build an inexpensive-to-operate reusable BFR fleet, and watch it fail to attract payloads. 

[Robotbeat]

All >100mt behemoths have common problem to be ultimate commercial rockets now or in near future. There are no payloads for them.

Yet.  I think the key would be getting more firms like Bigelow up and running.  If Bigelow could get some space stations into LEO, they might consider launching much bigger versions.  See attached for a 100 mt (65 mt empty) space station module almost twice the volume of the ISS . 

So you've launched one of them. Great. Now what? You've built all that for a couple launches?

[Patchouli]

All >100mt behemoths have common problem to be ultimate commercial rockets now or in near future. There are no payloads for them.

I agree with now.  But "near future" is undefined. If it existed, and was a lot cheaper per launch than the relatively tiny rockets of today, then my belief is that payloads would appear.  Especially payloads related to hypersonic terrestrial transportation (which would need to be planned in from the start of development), space tourism, and space colonization.  The only way to disprove my assumption is to build an inexpensive-to-operate reusable BFR fleet, and watch it fail to attract payloads. 

The best route to a commercial HLV would be to go with a modular system.
Something that first services EELV class payloads but can be clustered for heavier payloads.

[Robotbeat]

All >100mt behemoths have common problem to be ultimate commercial rockets now or in near future. There are no payloads for them.

I agree with now.  But "near future" is undefined. If it existed, and was a lot cheaper per launch than the relatively tiny rockets of today, then my belief is that payloads would appear.  Especially payloads related to hypersonic terrestrial transportation (which would need to be planned in from the start of development), space tourism, and space colonization.  The only way to disprove my assumption is to build an inexpensive-to-operate reusable BFR fleet, and watch it fail to attract payloads. 

The best route to a commercial HLV would be to go with a modular system.
Something that first services EELV class payloads but can be clustered for heavier payloads.

Agreed with this. This is essentially what SpaceX is doing, and it's essentially much of what people mean by "Atlas V Phase 2" (if I'm remembering correctly).

[Hyperion5]

All >100mt behemoths have common problem to be ultimate commercial rockets now or in near future. There are no payloads for them.

Yet.  I think the key would be getting more firms like Bigelow up and running.  If Bigelow could get some space stations into LEO, they might consider launching much bigger versions.  See attached for a 100 mt (65 mt empty) space station module almost twice the volume of the ISS . 

So you've launched one of them. Great. Now what? You've built all that for a couple launches?

Well you've got to remember Go4Mars' strategy and mine on HLVs are quite different.  He loves the idea of a monolithic HLV, while I've hedged my bets on a modular approach.  If you read back a few pages you'll find my Neptune proposal.  The basic LV is a 2-stage, 638 mt, 5-engine per stage (with T+.1 engine-out capability on both stages) methalox rocket developing 2,000,000 lbf of thrust at liftoff.  Just like an Atlas V Phase 2, the biggest Neptune would use additional CCBs (up to 4) to become a Saturn V-class rocket. 

The family was aimed at the Proton M in basic version but given I was using methane and staged combustion cycle engines on both stages, it turns out it will outlift an 800 mt Chang Zheng 5 (with margin to spare) in the simulations.  I hadn't quite counted on it performing as well as it has, so perhaps it's a tad oversized, but the basic premise is sound.  Just like the Atlas V Phase 2, it provides an easy way to tailor one LV family to a wide range of possible payloads.  However, unlike the Atlas Phase 2, with 5 engines per stage in x-layout it has both potential reusability and better engine-out capability baked in. 

[RocketmanUS]

Atlas V and F9 v1.1 For crew to LEO.

Heavy launch vehicle wide body to LEO.
Launch multiply rockets in a short time during good weather to LEO depot.
Store items at depot and send items to their destinations later as needed or when windows open up for BEO. Payloads what long periods of time to get a ride up anyway so there is no problem having them what at a LEO depot above the weather.

Should lower cost by having multiple launches ( HLV ) in a short period and give the standing army of workers time off between these launches.

Can use SEP tugs to deliver items to higher orbits or EML1/2, LLO from a LEO depot.

Ultimate commercial rocket design would be to have three and a half types. All should be liquid launchers with no solids except maybe the first one listed.
1 ) Pegasus class ( up to 2,000 lb to LEO )
2 ) Atlas V ( no SRB ) and Falcon 9
3 ) Delta IV Heavy and Falcon Heavy ( common boosters to core making this the half type launcher )
4 ) The Heavy wide body launcher ( this class to be launched multiple time in a given period with 1 to 3 months in between these multiple launches )

[Hyperion5]

Atlas V and F9 v1.1 For crew to LEO.

Heavy launch vehicle wide body to LEO.
Launch multiply rockets in a short time during good weather to LEO depot.
Store items at depot and send items to their destinations later as needed or when windows open up for BEO. Payloads what long periods of time to get a ride up anyway so there is no problem having them what at a LEO depot above the weather.

Should lower cost by having multiple launches ( HLV ) in a short period and give the standing army of workers time off between these launches.

Can use SEP tugs to deliver items to higher orbits or EML1/2, LLO from a LEO depot.

Well so long as you're not depending on hydrolox stages you should be able to carry that off.  I still think a modular family than can do it all (Falcon 9-class, Falcon Heavy-class, Energia-class) would save more by eliminating some of the flight risk.  It could still take advantage of SEP tugs but launch far larger models. 

Ultimate commercial rocket design would be to have three and a half types. All should be liquid launchers with no solids except maybe the first one listed.
1 ) Pegasus class ( up to 2,000 lb to LEO )
2 ) Atlas V ( no SRB ) and Falcon 9
3 ) Delta IV Heavy and Falcon Heavy ( common boosters to core making this the half type launcher )
4 ) The Heavy wide body launcher ( this class to be launched multiple time in a given period with 1 to 3 months in between these multiple launches )

I have to disagree with type #1.  Witness the Russians killing off the smallest Angara variant, or the fact that the most launched rocket family in history, the R-7 family, is much more capable than that.  You can even see just how much more competitive the bigger Falcon 9 is than the Falcon 1.  Cost per kg plunged when Spacex went bigger, where the demand was.  It'll go lower still when the 1.1 and Falcon Heavy start flying. 

I think the Angara family would be a better model for what you're proposing.  You've got the Angara 1-5, which will handle everything from Delta II class payloads to stuff beyond the LEO capability of a Delta IV Heavy (28.5 mt vs 23 mt).  You add a wide-body core for the Angara 7 variants (4.1 m--kerolox or 5.7 m hydrolox with RD-0120 engine) for your HLV (35-50 mt).  Otherwise if you're proposing 3 different cores for your commercial payload line that's a lot of overhead.  Angara family can get you down to two cores tops unless you're wanting even more capability than 50 mt. 

For that may I suggest a modular approach using a Zenit/Falcon 9-class LV with up to 4 CCBs?  That'd get you down to one core and possibly as few as one or two engines for almost anything you could want. 

[RocketmanUS]

Atlas V and F9 v1.1 For crew to LEO.

Heavy launch vehicle wide body to LEO.
Launch multiply rockets in a short time during good weather to LEO depot.
Store items at depot and send items to their destinations later as needed or when windows open up for BEO. Payloads what long periods of time to get a ride up anyway so there is no problem having them what at a LEO depot above the weather.

Should lower cost by having multiple launches ( HLV ) in a short period and give the standing army of workers time off between these launches.

Can use SEP tugs to deliver items to higher orbits or EML1/2, LLO from a LEO depot.

Well so long as you're not depending on hydrolox stages you should be able to carry that off.  I still think a modular family than can do it all (Falcon 9-class, Falcon Heavy-class, Energia-class) would save more by eliminating some of the flight risk.  It could still take advantage of SEP tugs but launch far larger models. 

Why not hydrolox stages?

Ultimate commercial rocket design would be to have three and a half types. All should be liquid launchers with no solids except maybe the first one listed.
1 ) Pegasus class ( up to 2,000 lb to LEO )
2 ) Atlas V ( no SRB ) and Falcon 9
3 ) Delta IV Heavy and Falcon Heavy ( common boosters to core making this the half type launcher )
4 ) The Heavy wide body launcher ( this class to be launched multiple time in a given period with 1 to 3 months in between these multiple launches )

I have to disagree with type #1.  Witness the Russians killing off the smallest Angara variant, or the fact that the most launched rocket family in history, the R-7 family, is much more capable than that.  You can even see just how much more competitive the bigger Falcon 9 is than the Falcon 1.  Cost per kg plunged when Spacex went bigger, where the demand was.  It'll go lower still when the 1.1 and Falcon Heavy start flying. 

I think the Angara family would be a better model for what you're proposing.  You've got the Angara 1-5, which will handle everything from Delta II class payloads to stuff beyond the LEO capability of a Delta IV Heavy (28.5 mt vs 23 mt).  You add a wide-body core for the Angara 7 variants (4.1 m--kerolox or 5.7 m hydrolox with RD-0120 engine) for your HLV (35-50 mt).  Otherwise if you're proposing 3 different cores for your commercial payload line that's a lot of overhead.  Angara family can get you down to two cores tops unless you're wanting even more capability than 50 mt. 

For that may I suggest a modular approach using a Zenit/Falcon 9-class LV with up to 4 CCBs?  That'd get you down to one core and possibly as few as one or two engines for almost anything you could want. 

#1 could be replace with at least one dedicated F9 v1.1 launch in a calender year with a 4 month launch window. Another company wants to place small sat to LEO for $10M. So if F9 were to carry at least 6 of those sat then the small launcher might not be needed. The company needs a given amount of launches per year for that price.

If it were a F1 ( Merlin 1D ) I don't think that it would be that big of a problem even if it were just 1 launch a year. Say priced at $10M. With the Merlin 1D it should be able to use a 40 inch inside wide fairing with at least 2,000lb to LEO. I believe that there is a smart way that SpaceX could launch this on either coast and make the stage with the personnel they have. Flooring and tooling should not be a big problem.

Edit:
The second stage could be CH4/LOX. A way for them to test new designs for that type of engine and stage at a lower cost.

For the HLV the Atlas V could be a booster to a wider body core or a single engine or triple engine Falcon booster to a wide body core ( 9 engines on each core would be to much ). Core would use the same engines as the boosters.

The other option would be for the HLV to be a two stage vehicle. 1st and second stage use same type engine with 2nd stage engine vac rated. Minimum of six launches a year with at least three launches at any launch period. This I think would be the better of the two options. HLV for cargo only.

[Hyperion5]

Atlas V and F9 v1.1 For crew to LEO.

Heavy launch vehicle wide body to LEO.
Launch multiply rockets in a short time during good weather to LEO depot.
Store items at depot and send items to their destinations later as needed or when windows open up for BEO. Payloads what long periods of time to get a ride up anyway so there is no problem having them what at a LEO depot above the weather.

Should lower cost by having multiple launches ( HLV ) in a short period and give the standing army of workers time off between these launches.

Can use SEP tugs to deliver items to higher orbits or EML1/2, LLO from a LEO depot.

Well so long as you're not depending on hydrolox stages you should be able to carry that off.  I still think a modular family than can do it all (Falcon 9-class, Falcon Heavy-class, Energia-class) would save more by eliminating some of the flight risk.  It could still take advantage of SEP tugs but launch far larger models. 

Why not hydrolox stages?

Well besides the fact that I've yet to hear of a "cheap" hydrolox engine, you've got to worry about boil-off.  Unless you're willing to spend the cash for a cryo-cooler and ACES-like upgrade, you're going to have to launch everything within a few days.  Depots will help, but only if you've spent the cash on sun shades or cryo-coolers to enable refueling.  Otherwise that doesn't leave a lot of margin for error or schedule slips, which do happen in this business.  That's the whole reason why, if I were doing things, I'd do it all on a single HLV launch with its own dedicated EDS.  Before this is brought up again, that's a modular, 5-core HLV.

[Patchouli]

Atlas V and F9 v1.1 For crew to LEO.

Heavy launch vehicle wide body to LEO.
Launch multiply rockets in a short time during good weather to LEO depot.
Store items at depot and send items to their destinations later as needed or when windows open up for BEO. Payloads what long periods of time to get a ride up anyway so there is no problem having them what at a LEO depot above the weather.

Should lower cost by having multiple launches ( HLV ) in a short period and give the standing army of workers time off between these launches.

Can use SEP tugs to deliver items to higher orbits or EML1/2, LLO from a LEO depot.

Well so long as you're not depending on hydrolox stages you should be able to carry that off.  I still think a modular family than can do it all (Falcon 9-class, Falcon Heavy-class, Energia-class) would save more by eliminating some of the flight risk.  It could still take advantage of SEP tugs but launch far larger models. 

Ultimate commercial rocket design would be to have three and a half types. All should be liquid launchers with no solids except maybe the first one listed.
1 ) Pegasus class ( up to 2,000 lb to LEO )
2 ) Atlas V ( no SRB ) and Falcon 9
3 ) Delta IV Heavy and Falcon Heavy ( common boosters to core making this the half type launcher )
4 ) The Heavy wide body launcher ( this class to be launched multiple time in a given period with 1 to 3 months in between these multiple launches )

I have to disagree with type #1.  Witness the Russians killing off the smallest Angara variant, or the fact that the most launched rocket family in history, the R-7 family, is much more capable than that.  You can even see just how much more competitive the bigger Falcon 9 is than the Falcon 1.  Cost per kg plunged when Spacex went bigger, where the demand was.  It'll go lower still when the 1.1 and Falcon Heavy start flying. 

I think the Angara family would be a better model for what you're proposing.  You've got the Angara 1-5, which will handle everything from Delta II class payloads to stuff beyond the LEO capability of a Delta IV Heavy (28.5 mt vs 23 mt).  You add a wide-body core for the Angara 7 variants (4.1 m--kerolox or 5.7 m hydrolox with RD-0120 engine) for your HLV (35-50 mt).  Otherwise if you're proposing 3 different cores for your commercial payload line that's a lot of overhead.  Angara family can get you down to two cores tops unless you're wanting even more capability than 50 mt. 

For that may I suggest a modular approach using a Zenit/Falcon 9-class LV with up to 4 CCBs?  That'd get you down to one core and possibly as few as one or two engines for almost anything you could want. 

Something like sodruzhestvo.
http://www.russianspaceweb.com/sodruzhestvo.html

A US equivalent could use a single TR-107,RS-84, or F-1A or similar on each core.

The single F-1A variant could be an extremely reliable LV.

The RS-84 could allow upgrades like flyback boosters.

[RocketmanUS]

Atlas V and F9 v1.1 For crew to LEO.

Heavy launch vehicle wide body to LEO.
Launch multiply rockets in a short time during good weather to LEO depot.
Store items at depot and send items to their destinations later as needed or when windows open up for BEO. Payloads what long periods of time to get a ride up anyway so there is no problem having them what at a LEO depot above the weather.

Should lower cost by having multiple launches ( HLV ) in a short period and give the standing army of workers time off between these launches.

Can use SEP tugs to deliver items to higher orbits or EML1/2, LLO from a LEO depot.

Well so long as you're not depending on hydrolox stages you should be able to carry that off.  I still think a modular family than can do it all (Falcon 9-class, Falcon Heavy-class, Energia-class) would save more by eliminating some of the flight risk.  It could still take advantage of SEP tugs but launch far larger models. 

Why not hydrolox stages?

Well besides the fact that I've yet to hear of a "cheap" hydrolox engine, you've got to worry about boil-off.  Unless you're willing to spend the cash for a cryo-cooler and ACES-like upgrade, you're going to have to launch everything within a few days.  Depots will help, but only if you've spent the cash on sun shades or cryo-coolers to enable refueling.  Otherwise that doesn't leave a lot of margin for error or schedule slips, which do happen in this business.  That's the whole reason why, if I were doing things, I'd do it all on a single HLV launch with its own dedicated EDS.  Before this is brought up again, that's a modular, 5-core HLV.

The depot is mostly for hardware. SEP would most likely use Argon.
Lunar lander would use hypergolics till Lunar ISRU was available.
So depot would store Argon and hypergolic propellants along with the hardware. Depot would be unmanned most of the time.

As for HLV an Atlas V core as boosters around a wider core powered by one or more RD-180's could be used. RD-180's are sold outside of Russia So this could be made into a global launcher. The booster and core would need to be made from tech that each of the launch countries would already have. So options might be U.S., Russia , ESA, China. Multiple launch sites and could keep prices down.

[Hyperion5]

The depot is mostly for hardware. SEP would most likely use Argon.
Lunar lander would use hypergolics till Lunar ISRU was available.
So depot would store Argon and hypergolic propellants along with the hardware. Depot would be unmanned most of the time.

I know hypergolic propellants offer major storing advantages, but are liquid methane and oxygen really that hard to keep from boiling off in space?  I was always under the impression it was liquid hydrogen that had the major issues in that area.  If I recall right, the Lox on a Saturn V's S-II was 70 degrees Celsius warmer than the hydrogen being stored in the tank next to it.  In contrast liquid methane would be 20 degrees warmer than Lox.  Or perhaps I've screwed up my units of measurement and that was the difference in Fahrenheit. 

As for HLV an Atlas V core as boosters around a wider core powered by one or more RD-180's could be used. RD-180's are sold outside of Russia So this could be made into a global launcher. The booster and core would need to be made from tech that each of the launch countries would already have. So options might be U.S., Russia , ESA, China. Multiple launch sites and could keep prices down.

That sounds a lot like the cancelled Rus-M rocket family.  I'm sure  TsSKB-Progress wouldn't mind your idea of using their cores in the slightest.  At least it'd get them some business at the expense of Khrunichev. 

[RocketmanUS]

The depot is mostly for hardware. SEP would most likely use Argon.
Lunar lander would use hypergolics till Lunar ISRU was available.
So depot would store Argon and hypergolic propellants along with the hardware. Depot would be unmanned most of the time.

I know hypergolic propellants offer major storing advantages, but are liquid methane and oxygen really that hard to keep from boiling off in space?  I was always under the impression it was liquid hydrogen that had the major issues in that area.  If I recall right, the Lox on a Saturn V's S-II was 70 degrees Celsius warmer than the hydrogen being stored in the tank next to it.  In contrast liquid methane would be 20 degrees warmer than Lox.  Or perhaps I've screwed up my units of measurement and that was the difference in Fahrenheit. 

As for HLV an Atlas V core as boosters around a wider core powered by one or more RD-180's could be used. RD-180's are sold outside of Russia So this could be made into a global launcher. The booster and core would need to be made from tech that each of the launch countries would already have. So options might be U.S., Russia , ESA, China. Multiple launch sites and could keep prices down.

That sounds a lot like the cancelled Rus-M rocket family.  I'm sure  TsSKB-Progress wouldn't mind your idea of using their cores in the slightest.  At least it'd get them some business at the expense of Khrunichev. 

SEP takes months to go from LEO to EML1/2 or LLO. So non hypergolics could boil off. If it's not a problem and there is the tech to transfer them in space plus the Lunar lander to use them , then go with the higher ISP.

[Robotbeat]

Soft cryogenics are space-storable (passively). It's hydrogen you've got to be worried about, and even that can be reduced to very low boil-off.

[A_M_Swallow]

Soft cryogenics are space-storable (passively). It's hydrogen you've got to be worried about, and even that can be reduced to very low boil-off.

Just accept that SEP tankers need the same protection as depots.  A 10% increase in the solar panels will give you lots of electricity for things like refrigeration equipment.  There is also plenty of room for sun shields, possibly attached to the solar arrays.

[RocketmanUS]

Soft cryogenics are space-storable (passively). It's hydrogen you've got to be worried about, and even that can be reduced to very low boil-off.

Just accept that SEP tankers need the same protection as depots.  A 10% increase in the solar panels will give you lots of electricity for things like refrigeration equipment.  There is also plenty of room for sun shields, possibly attached to the solar arrays.

Added mass.
Either make a larger SEP , increased time to destination, or decrease payload. So don't use hydrogen, it take up to much volume anyway.

Or Atlas V and F9 would be good for delivering propellants to LEO. Hydrogen gets good ISP with high thrust with LOX but need very large tanks plus special cooling and handling.

[go4mars]

Well you've got to remember Go4Mars' strategy and mine on HLVs are quite different.  He loves the idea of a monolithic HLV, while I've hedged my bets on a modular approach.

I like modular too, but on a bigger scale. 

Can fly 1,2,3, or 5 1st stage cores per launch.  5 would be shaped like 8o8
Fully reusable, always crossfeeds when more than 1 core. 

[Hyperion5]

Well you've got to remember Go4Mars' strategy and mine on HLVs are quite different.  He loves the idea of a monolithic HLV, while I've hedged my bets on a modular approach.

I like modular too, but on a bigger scale. 

Can fly 1,2,3, or 5 1st stage cores per launch.  5 would be shaped like 8o8
Fully reusable, always crossfeeds when more than 1 core. 

Yes, it's modular, but how often could you possibly ever need to fling 2000+ tonnes of payload at a time into LEO?  I'm sure that despite what some say you could make a fairly substantial LV commercially viable if you charged a low enough price per kg, but surely there's a limit.  My guess is you could at most make something pushing 120 mt viable if you got the costs down, but above that you're running out of payloads.  To develop a payload that even a Saturn V couldn't launch is neither cheap nor a run-of-the-mill engineering exercise.  I suppose we all better hope that Bigelow Aerospace starts getting some orders for its massive BA-2100 space station modules.  The problem is, just one of those has roughly double the interior volume of the ISS, so supplying such large space stations would be a feat.  Though at least ULA, Orbital, Spacex, Arianespace, Khrunichev or RSC Energia would not be complaining about the new demand. 

[Robotbeat]

Besides, the acoustic energy (etc) means you may need to build a new launch pad every time you launch the 2000 mT monster.

[oldAtlas_Eguy]

My ultimate rocket is driven by what would be needed to make SPS commercially viable. A LEO cost rate of <$500/kg (1/4 of that of FH) and a payload capacity of 400-600mt. This would be a triple core MCT like rocket (single core able to do 130-200mt). Some reusability would be used in order to get flight rates up to 25-50 per year, giving a new vehicle build rate of 2-5/yr. Flight hardware retired after 10-20 flights. Full reusability could get the LEO cost rates much lower, but anything less than $500/kg will do (BTW that’s a per launch price of <($200M for the smaller capacity or $300M for the larger capacity) per launch, a revenue from launches of $5B-$15B/yr. Smaller vehicles will not work because in order to support such large scale SPS you would need 200-2500 launches/yr of either a 50mt or 13mt fully reusable vehicle.

The triple core configuration would be cargo only but the single core could be used for personnel transport of as many as 100 persons per launch. In this configuration the US and personnel transport could be an integrated vehicle. Depends on what will work out for safety and cost.

If you are to believe SpaceX such vehicles will be available in 10 years, but using hindsight such a large triple core 400-600mt vehicle would not be available until 2028 with the single or 130-200mt class by 2023.

[Hyperion5]

My ultimate rocket is driven by what would be needed to make SPS commercially viable. A LEO cost rate of <$500/kg (1/4 of that of FH) and a payload capacity of 400-600mt. This would be a triple core MCT like rocket (single core able to do 130-200mt). Some reusability would be used in order to get flight rates up to 25-50 per year, giving a new vehicle build rate of 2-5/yr. Flight hardware retired after 10-20 flights. Full reusability could get the LEO cost rates much lower, but anything less than $500/kg will do (BTW that’s a per launch price of <($200M for the smaller capacity or $300M for the larger capacity) per launch, a revenue from launches of $5B-$15B/yr. Smaller vehicles will not work because in order to support such large scale SPS you would need 200-2500 launches/yr of either a 50mt or 13mt fully reusable vehicle.

The triple core configuration would be cargo only but the single core could be used for personnel transport of as many as 100 persons per launch. In this configuration the US and personnel transport could be an integrated vehicle. Depends on what will work out for safety and cost.

If you are to believe SpaceX such vehicles will be available in 10 years, but using hindsight such a large triple core 400-600mt vehicle would not be available until 2028 with the single or 130-200mt class by 2023.

Would your MCT-equivalent core be using existing engines like the F-1A or a totally new design?  Any ideas on core size and engine count? 

[oldAtlas_Eguy]

My ultimate rocket is driven by what would be needed to make SPS commercially viable. A LEO cost rate of <$500/kg (1/4 of that of FH) and a payload capacity of 400-600mt. This would be a triple core MCT like rocket (single core able to do 130-200mt). Some reusability would be used in order to get flight rates up to 25-50 per year, giving a new vehicle build rate of 2-5/yr. Flight hardware retired after 10-20 flights. Full reusability could get the LEO cost rates much lower, but anything less than $500/kg will do (BTW that’s a per launch price of <($200M for the smaller capacity or $300M for the larger capacity) per launch, a revenue from launches of $5B-$15B/yr. Smaller vehicles will not work because in order to support such large scale SPS you would need 200-2500 launches/yr of either a 50mt or 13mt fully reusable vehicle.

The triple core configuration would be cargo only but the single core could be used for personnel transport of as many as 100 persons per launch. In this configuration the US and personnel transport could be an integrated vehicle. Depends on what will work out for safety and cost.

If you are to believe SpaceX such vehicles will be available in 10 years, but using hindsight such a large triple core 400-600mt vehicle would not be available until 2028 with the single or 130-200mt class by 2023.

Would your MCT-equivalent core be using existing engines like the F-1A or a totally new design?  Any ideas on core size and engine count? 

See the SpaceX MCT thread for some of my thoughts on the engines and cores.

Basiclly a new 1.7mlbf methane engine used in a configuration of 6 on a single core giving ~150mt LEO capability and in a triple core configuration 600mt with crossfeed. Tank diameters of 8-10m. More likely 10m since anlthing over 3.6m requires use of barges so going to 10m vs 5m does not cost any more to transport. Plus a non-hamerhead single core 10m faring is appealing and its height is also manageble for use with methane prop.

[go4mars]

Yes, it's modular, but how often could you possibly ever need to fling 2000+ tonnes of payload at a time into LEO? 

Depends.   Does your cycler need a pusher plate?   Do your transport ships to the Mars colonies each house 30?  Or 150 people?   How much methane and oxygen do you expect to need for your successful mining/solar smelting operation?  Does the BA-21,000 go up with all it's water/shielding inside, or does that go on subsequent flights?  How much do you expect the lead-vortex General Fusion ultimate ISP solution to mass (needs a certain minimum size)?     And perhaps most importantly, how many people do you plan to have in the cabin segment of your hypersonic terrestrial transportation modules? 

[savuporo]

It would be one of these
http://www.scorpius.com/products.htm

Sprite, with 1000 lbs to LEO, more specifically.

Small enough parts to be transportable, assembly line manufacturing, with established mass production quality control methods, 95% reliable, only good for launching bulk materials with low payload value.

[Hyperion5]

It would be one of these
http://www.scorpius.com/products.htm

Sprite, with 1000 lbs to LEO, more specifically.

Small enough parts to be transportable, assembly line manufacturing, with established mass production quality control methods, 95% reliable, only good for launching bulk materials with low payload value.

Technically even the 5 meter core on an Ariane 5 or Delta IV is "transportable", they just need water transport to get to their pads.  Now I realize you're gunning for the Sprite, but those other rockets I see on offer, beginning with the Liberty, seem a lot more promising commercially.  They offer bigger payload fairings and more mass to LEO.  Given the most oft-launched launcher in history is considerably more capable than the Sprite (see R-7 family), wouldn't it make sense to up the scale and recommend something like the Exodus (19,700 lbs to LEO) or Space Freighter (33,700 lbs to LEO)? 

--------------

Since we're advocating rockets, here's what that Neptune V I envisioned looked like in its early simulations.  The design since these simulations has grown from 5.21 meters in diameter to 5.79 meters to improve the engine expansion ratios but has seen its length trimmed to just under 58 meters.  We also hedged in more margin with higher dry masses for both stages and cut down the combustion efficiency from 98 to 96%.  Yet despite all that I am fairly confident it should still out-lift an Angara 5 to LEO despite weighing less all while losing an engine at T+0.1 on both stages.  If you could get the engines and avionics quality up, this thing should crush many rockets in reliability given its engine-out margins. 

[Hyperion5]

Baldusi has been talking up this RD-162 engine so many times that I think it's high time we flung it onto a rocket geared for this thread.  So I dual-chambered it into a theoretical equivalent of the RD-180, added on a quartet of RL-10 equivalents, and the all-Russian counterpart to the Atlas V looked like this:


Novgorod 1
2 or 3-stage methalox/hydrolox rocket
Gross mass w/2 stages: 326.4 tonnes
SI gross mass: 281.4 tonnes
SII gross mass: 45 tonnes
Liftoff thrust: 407.8 tonnes
Diameter: 3.8 meters
Standard PLF: 4.1 meters
SI Engine: RD-163 dual-chamber fuel-rich SC methalox engine
SII Engines: quartet of RD-0146 hydrolox engines
Optional 3rd stage: Fregat SB or KVTK w/RD-0146D engine

Payload to LEO: 14,500 kg (estimate)
Payload to GTO: 6,500 kg (estimate)

Novgorod 3 (3 core)
Payload to LEO: 37,500 kg (estimate)
Payload to GTO: 16,000 kg (estimate)

Novgorod 5 (5 core)
Payload to LEO: 67,500 kg (estimate)
Payload to GTO: 28,000 kg (estimate)

I think this design might actually be more promising the the Angaras, because it could also replace the Soyuz & Zenit families in its single-core variant.  You also wouldn't need a unique core to get more than 35 mt of LEO capability.  It should also be quite reliable thanks to that upper stage setup.  Too bad the Russians aren't building something like it, because it looks really promising. 

[Falcon H]

My ultimate rocket design is a falcon heavy, but with seven cores instead of three, the cores could be recovered useing parachutes, and this beast would need a larger upper stage and fairing, maybe 5 meters wide.  

[Falcon H]

My ultimate rocket design is a falcon heavy, but with seven cores instead of three, the cores could be recovered useing parachutes, and this beast would need a larger upper stage and fairing, maybe 5 meters wide.  

Here is a sketch I made.

[Hyperion5]

My ultimate rocket design is a falcon heavy, but with seven cores instead of three, the cores could be recovered useing parachutes, and this beast would need a larger upper stage and fairing, maybe 5 meters wide.  

First off, I'm pretty sure that structurally the Falcon 9 is not built to handle six booster cores strapped to it, and even it if it is, that is an otherworldly 63 rocket engines.  I know Lobo thinks he can get up to six Falcon Heavy boosters strapped to the SLS and make it work, but I'm skeptical.  I know Lobo's a smart guy with an education in engineering to back that up, but I've yet to see people in the aerospace engineering field like Strangequark or Jim embrace using FH boosters in that quantity.  I know Ed Kyle is openly skeptical of just the 27 engines on the Falcon Heavy, so you can imagine his reaction to your proposal. 

Recovery of the cores is likely to be done via propulsive landings on terra firma, following the Grasshopper model.  I forgot to mention that Elon's already granted that last wish of yours.  The standard fairing on both the Falcon 9 & Falcon Heavy is 5.2 meters. 

[gin455res]

How about a falcon 1, with an xcor designed upperstage to open up the smallsat market?

[strangequark]

I know Lobo's a smart guy with an education in engineering to back that up, but I've yet to see people in the aerospace engineering field like Strangequark or Jim embrace using FH boosters in that quantity.  I know Ed Kyle is openly skeptical of just the 27 engines on the Falcon Heavy, so you can imagine his reaction to your proposal. 

So, yeah, that's a crazy number of boosters. However, I have 5 years of experience in this industry, and have been corrected and educated many times on this forum by "non-experts". Jim has 30-ish years experience, and an encyclopedic knowledge of spaceflight practice and history. I do not belong anywhere near that particular pedestal.

[Hyperion5]

I know Lobo's a smart guy with an education in engineering to back that up, but I've yet to see people in the aerospace engineering field like Strangequark or Jim embrace using FH boosters in that quantity.  I know Ed Kyle is openly skeptical of just the 27 engines on the Falcon Heavy, so you can imagine his reaction to your proposal. 

So, yeah, that's a crazy number of boosters. However, I have 5 years of experience in this industry, and have been corrected and educated many times on this forum by "non-experts". Jim has 30-ish years experience, and an encyclopedic knowledge of spaceflight practice and history. I do not belong anywhere near that particular pedestal.

I don't put Jim on a pedestal; FYI I'm just using you and Jim as examples of aerospace guys I know who aren't exactly jumping up and down saying "63 engines firing at once?  Let's do this!"  I mean, let's just say Falcon H is your boss, and he comes to you with this plan to mount not two but six Falcon Heavy boosters around the Falcon 9 core.  What kind of issues would you point out with this approach? 

I would bet you would have almost immediate reservations about this proposed vehicle, and not only because of there being some 63 rocket engines firing at once.  Just think about the payload masses we start being able to lift with this amount of liftoff thrust.  Supposing you could get past the booster issues, that still leaves you with an absurdly underpowered second stage.  The fairing sizes you could mount would also be inadequate for most 100 mt payloads.  Finally, the structural loads placed on the central core by the boosters and huge payloads would probably force engineers to reinforce the core a lot, which in turn would curb the payload mass you could launch.  Then there's the launchpad for a monster like this; the only ones I can think of that might handle it would be the old Shuttle/Saturn pads.  That means either higher operating costs or hundreds of millions or more for an all-new pad. 

[ClaytonBirchenough]

My ultimate commercial rocket design is not really a design, it's a philosophy; mass production, flying a rocket a day, no matter if payload is available. The rocket would be two stages, have the simplest engines burning LOX/RP-1. One engine in each stage. Rocket could probably launch 10 MT into LEO. If no payload is available everyday, launch extra propellant to a propellant depot. Once a week, you launch a similar rocket, but human-rated carrying a Dragon like capsule.

Payload delivered to LEO: 3650 MT (Excluding manned rocket)
Passengers delivered to LEO: ~364

MASS PRODUCTION MY FRIENDS

[Elmar Moelzer]

My ultimate commercial rocket design would be something designed by Gary Hudson, if someone just came and finally gave the man the funding he deserves.
I don't know exactly what Gary would make it look like with his increased experience and new materials and other options available these days, but I would guess it would be somewhat resembling the Aerospike Test Vehicle, the DC-X and/or the Phoenix VTOL.
I know, I can dream, alright. Not much else I can do anyway.

[QuantumG]

My ultimate commercial rocket design would be something designed by Gary Hudson, if someone just came and finally gave the man the funding he deserves.
I don't know exactly what Gary would make it look like with his increased experience and new materials and other options available these days, but I would guess it would be somewhat resembling the Aerospike Test Vehicle, the DC-X and/or the Phoenix VTOL.
I know, I can dream, alright. Not much else I can do anyway.

Good answer!

[Oli]

Skylon.

Otherwise a reusable kerosene booster with an orbital hydrogen lifting body on top (x-33 on a stick, if you will )

[Lobo]

Since Hyperion pointed me to this thread and ask for my normally incredibly verbose opinion, I'll share a thought here:

First, I don’t know a ton about the pros/cons of methalox vs. kerolox.  I do know that they seem to have pretty similar performance, and are relatively close in density/volume.
But by switching from kerolox to methalox, you’ve traded up one of the most easily handled, transported, and stored fuel’s in RP-1 for a mild cryogenic that’s similar to LOX, with those associated difficulties.

But it seems methalox upper stags can get better isp than kerolox.  Raptor is supposed to be about 380s, while RD-0124 (one of the most efficient kerolox upper stages I think) is about 359s.  Hyperion has pointed out some metallurgical advantages of methalox, and seems to like it, so we’ll pull on that thread some.

From our discussion about Titan and Saturn 1 over on the Apollo not being cancelled thread, it seems that core lengths are very easy, but core diameter changes are much harder.  So I think we should pick one common core diameter that can do it all, and then have different core lengths for different options and applications.

There are customers for payloads in the less than 10mt range.  But I think that F9, Antares, Delta IV, Atlas V, etc all have a similar capacity of 10mt.  I think that’s a good place to have your basic building block.  If the LV is cheap, then it doesn’t really matter if it’s a bit oversized.  I think this is why SpaceX ended up bypassing Falcon 1 and Falcon 5, and went to Falcon 9.   There is still Delta II, but I think if you are looking for a “magic bulluet” LV, I’d probably start with that happy medium of about 10mt.

On the upper end, how big do we need to go?  Well, I think if you had a cheap, 30-40mt launcher, you are good for any payload but a NASA HSF payload for the foreseeable future.  And having a wide payload diameter capability is probably almost more important than pure mass.  But, if you could get this LV system up over 100mt, then I think you’d be in good shape for some wild hypothetical of NASA cancelling SLS and coming to us for our dirt-cheap common commercial launcher.  (Although DCCCL is a bit unwieldy as a name.)

I think a 5m diameter core is a good diameter to have a 10mt to 100mt+ range.  Maybe 5.5m.  Build them someplace where they can be loaded onto a transport ship and shipped the Cape for launching.  The Atlantic or Gulf coasts are the most likely.    They can be shipped to VAFB through the Panama Canal when necessary for polar launches.

So I say focus on that.  I think one larger core is cheaper than 3 smaller cores, and fewer cores have better mass fractions that big clusters of cores.  So there can be a balance.  I don’t think the core itself costs much more or less because of size (well, within reason).  I don’t think a 3m core, a 4m core, or a 5m core will really cost much more or less to build once the tooling is in place.  Unless you plan to use existing tooling, you will make the tooling new for your LV regardless of diameter.

Then design the engine like the RD-170/180/190.  1-chamber, 2-chamber, or 4-chamber, depending on the need, with common turbo pumps.  That makes for a modular and flexible engine system too.
Make the MPS the same for all, just a matter of if you are installing 1, 2, or 4 chambers of the engine.  (Except the long core will need a different MPS, but I’ll get to that in a moment.)

Make the single chamber the equal to the RD-190 series (or AJ26-500), about 500 klbs.  A two chamber will be basically an AJ-1-E6 or RD-180 at 1Mlbs.  The 4-chamber version will be basically a 2Mlbs RD-171.

Your basic 5m, “short” core is going to be the rough equivalent of Atlas V, Delta IV, Antares, and Falcon 9.  About 10mt-ish to LEO.  Like I said, that seems to be a pretty useful workhorse size.  It’s probably be about ½ the height of an Atlas Phase 2 core about 15m tall?  I think Atlas Phase 2 is about 30mt tall, and Delta IV is about 40m tall…all in 15mt.  ULA has a short Atlas phase 2 concept with a single RD-180.  The “medium core” could be basically an Atlas Phase 2, but with a 4-chamber single engine rather than two 2-chamber RD-180’s.  The short and medium cores will have the same MPS, so it’s just a matter of installing 2 or 4 thrust chambers into the engine mount. 

The short core would probably be about 1/3 the height of Delta IV, it’s mathalox, and it’d have a common bulkhead.  The “medium” core would be about 2/3 the height of the Delta IV, and the “tall” core would be about the same height as Delta IV or a little taller.

However, the “Tall” core will need two of the quad-chamber engines.  It will need a new MPS with two engines mounts.  It would look very similar to the Dynetics booster, but with two quad-chamber engines rather than two single chamber engines like the F-1’s.   However, a tri-core medium core will be about the equal to Atlas Phase 2-Heavy, or around 70mt.  Only if NASA chooses this family as their HSF launcher would you ever need the “tall” core with two engine MPS.  You plan for that potential upgrade, but defer the actual development of that MPS until such a time there was a customer for it.

The upper stage can have two lengths as well.  A shorter for the smaller variants, and a stretched version for larger variants.  They are 5m methalox too, using all the same tooling and common bulkhead as the boosters.  Like SpaceX’s Merlin’s, the upper stage is powered by the single chamber version of this engine, optimized for vacuum running with a nozzle extension, etc.  It is a 500klbs engine, too powerful for the smaller LV configurations.  So you design it to be able to throttle down to around 20%.  The RD-0124 can throttle down to 30%, so I think 20% is doable.  That gives you a 100klbs engine, similar to the Merlin 1-vacuum.  Maybe you could have a smaller turbopump so the engine would be “permanently throttled down” when used on the smaller variants.  But the engine is just a variant of the booster engine.
When used for larger variants, then use the full 500klbs version.  If there was ever a need for more 2nd stage power than that, then you have a vacuum version of the 2-chamber variant, for 1Mlbs.  That should be all you never need on the upper stage.

So I envision it looking something like this:

10mt to LEO with the short core/1-chamber engine.

25-30 mt to LEO with the medium core/2-chamber engine. (Atlas Phase 2 basically, or Zenit)

35-ish mt to LEO with a long core/4-chamber engine.  The Dynetics booster basically.  Only necessary if contracted for NASA, or some quadzillionaire that wants to go to Mars or something. Otherwise, variants of the short and medium cores will take care of all the commercial and USAF payloads.

20-70mt to LEO, the “medium” core with flexible options of “short” or “medium” outboard boosters.  It can even launch with just one outboard booster.  No reason why not, and the engine gimbals should handle the off-line center of mass just fine.

For NASA, the tri-core “tall” version would have 12Mlbs of thrust, and should get well over 100mt to LEO.  Probably more like 150mt.

No variant has more than two boosters.  Just various core and booster lengths and engine chamber configurations.  This keeps the pad nice and simple, it only needs three ports, and would look like the Delta IV pad. With similar Mobile Service Structure.  Just make the MSS tall enough, and the flame trench strong enough, to handle up to the largest config, right from the beginning.

The single core short and medium version will cover 10-20mt and will cover that vast majority of USAF/DoD and commercial payloads, and those have the best mass fraction.  Add one outboard booster if you need a little more.  And two outboard boosters if you need more yet.  Make them the medium CCB’s for even more power.  Even a big NASA booster could have a “Tall” core with two “medium” boosters, etc.  They will be bottom lift.  MPS to MPS connected, so the upper attach point doesn’t have to carry the load like SLS.  So different booster core lengths could be mixed and matched for core and boosters.

Lastly for the engines,  GG vs. staged combustion?  Don’t really know there, but, if Aeojet thinks they can be competitive with AJ26, AJ26-500, and AJ26-1E6, and the Russians can make RD-180 pretty cost effectively (yea, their labor costs are low, but still) , I think you could do this engine in staged combustion, and just makes lots of them to drive the costs down.  ...With GG, the upper stage needs to work harder because of the lower efficiency.  So I think you can do this as a staged combustion methalox common platform engine.

For now, I will call this LV family the “Lobo-tomoizer” and the engine the “Hog”…as in Harley Hog.  Fast and Loud.

I now open it up to the slings and arrows of everyone on this thread who’ll probably point out reasons it won’t work that I don’t see.  (of course, that’s the fun of it all, eh?)

[Hyperion5]

Since Hyperion pointed me to this thread and ask for my normally incredibly verbose opinion, I'll share a thought here:

First, I don’t know a ton about the pros/cons of methalox vs. kerolox.  I do know that they seem to have pretty similar performance, and are relatively close in density/volume.
But by switching from kerolox to methalox, you’ve traded up one of the most easily handled, transported, and stored fuel’s in RP-1 for a mild cryogenic that’s similar to LOX, with those associated difficulties.

But it seems methalox upper stags can get better isp than kerolox.  Raptor is supposed to be about 380s, while RD-0124 (one of the most efficient kerolox upper stages I think) is about 359s.  Hyperion has pointed out some metallurgical advantages of methalox, and seems to like it, so we’ll pull on that thread some.

From our discussion about Titan and Saturn 1 over on the Apollo not being cancelled thread, it seems that core lengths are very easy, but core diameter changes are much harder.  So I think we should pick one common core diameter that can do it all, and then have different core lengths for different options and applications.

There are customers for payloads in the less than 10mt range.  But I think that F9, Antares, Delta IV, Atlas V, etc all have a similar capacity of 10mt.  I think that’s a good place to have your basic building block.  If the LV is cheap, then it doesn’t really matter if it’s a bit oversized.  I think this is why SpaceX ended up bypassing Falcon 1 and Falcon 5, and went to Falcon 9.   There is still Delta II, but I think if you are looking for a “magic bulluet” LV, I’d probably start with that happy medium of about 10mt.

On the upper end, how big do we need to go?  Well, I think if you had a cheap, 30-40mt launcher, you are good for any payload but a NASA HSF payload for the foreseeable future.  And having a wide payload diameter capability is probably almost more important than pure mass.  But, if you could get this LV system up over 100mt, then I think you’d be in good shape for some wild hypothetical of NASA cancelling SLS and coming to us for our dirt-cheap common commercial launcher.  (Although DCCCL is a bit unwieldy as a name.)

I think a 5m diameter core is a good diameter to have a 10mt to 100mt+ range.  Maybe 5.5m.  Build them someplace where they can be loaded onto a transport ship and shipped the Cape for launching.  The Atlantic or Gulf coasts are the most likely.    They can be shipped to VAFB through the Panama Canal when necessary for polar launches.

So I say focus on that.  I think one larger core is cheaper than 3 smaller cores, and fewer cores have better mass fractions that big clusters of cores.  So there can be a balance.  I don’t think the core itself costs much more or less because of size (well, within reason).  I don’t think a 3m core, a 4m core, or a 5m core will really cost much more or less to build once the tooling is in place.  Unless you plan to use existing tooling, you will make the tooling new for your LV regardless of diameter.

Then design the engine like the RD-170/180/190.  1-chamber, 2-chamber, or 4-chamber, depending on the need, with common turbo pumps.  That makes for a modular and flexible engine system too.
Make the MPS the same for all, just a matter of if you are installing 1, 2, or 4 chambers of the engine.  (Except the long core will need a different MPS, but I’ll get to that in a moment.)

Make the single chamber the equal to the RD-190 series (or AJ26-500), about 500 klbs.  A two chamber will be basically an AJ-1-E6 or RD-180 at 1Mlbs.  The 4-chamber version will be basically a 2Mlbs RD-171.

Your basic 5m, “short” core is going to be the rough equivalent of Atlas V, Delta IV, Antares, and Falcon 9.  About 10mt-ish to LEO.  Like I said, that seems to be a pretty useful workhorse size.  It’s probably be about ½ the height of an Atlas Phase 2 core about 15m tall?  I think Atlas Phase 2 is about 30mt tall, and Delta IV is about 40m tall…all in 15mt.  ULA has a short Atlas phase 2 concept with a single RD-180.  The “medium core” could be basically an Atlas Phase 2, but with a 4-chamber single engine rather than two 2-chamber RD-180’s.  The short and medium cores will have the same MPS, so it’s just a matter of installing 2 or 4 thrust chambers into the engine mount. 

The short core would probably be about 1/3 the height of Delta IV, it’s mathalox, and it’d have a common bulkhead.  The “medium” core would be about 2/3 the height of the Delta IV, and the “tall” core would be about the same height as Delta IV or a little taller.

However, the “Tall” core will need two of the quad-chamber engines.  It will need a new MPS with two engines mounts.  It would look very similar to the Dynetics booster, but with two quad-chamber engines rather than two single chamber engines like the F-1’s.   However, a tri-core medium core will be about the equal to Atlas Phase 2-Heavy, or around 70mt.  Only if NASA chooses this family as their HSF launcher would you ever need the “tall” core with two engine MPS.  You plan for that potential upgrade, but defer the actual development of that MPS until such a time there was a customer for it.

The upper stage can have two lengths as well.  A shorter for the smaller variants, and a stretched version for larger variants.  They are 5m methalox too, using all the same tooling and common bulkhead as the boosters.  Like SpaceX’s Merlin’s, the upper stage is powered by the single chamber version of this engine, optimized for vacuum running with a nozzle extension, etc.  It is a 500klbs engine, too powerful for the smaller LV configurations.  So you design it to be able to throttle down to around 20%.  The RD-0124 can throttle down to 30%, so I think 20% is doable.  That gives you a 100klbs engine, similar to the Merlin 1-vacuum.  Maybe you could have a smaller turbopump so the engine would be “permanently throttled down” when used on the smaller variants.  But the engine is just a variant of the booster engine.
When used for larger variants, then use the full 500klbs version.  If there was ever a need for more 2nd stage power than that, then you have a vacuum version of the 2-chamber variant, for 1Mlbs.  That should be all you never need on the upper stage.

Someone can correct me if I'm wrong, but I can't recall an US engine being detuned like this to work if it was originally the booster engine.  It might save you a huge amount of expense, but might it not be cheaper just to go with the Spacex approach to one engine type on a rocket?  I do love the idea of a modular engine family like that of the RD-170/180/191 family from NPO Energomash. 

I do think the performance compromises (t/w ratio on the engine, possibly lower Isp due to throttling) might come to bite this plan of yours in the foot.  To me this sounds like a situation where we could really use the advice of a propulsion engineer to separate practical from unattainable. 

[Lobo]

Someone can correct me if I'm wrong, but I can't recall an US engine being detuned like this to work if it was originally the booster engine.  It might save you a huge amount of expense, but might it not be cheaper just to go with the Spacex approach to one engine type on a rocket?  I do love the idea of a modular engine family like that of the RD-170/180/191 family from NPO Energomash. 

I do think the performance compromises (t/w ratio on the engine, possibly lower Isp due to throttling) might come to bite this plan of yours in the foot.  To me this sounds like a situation where we could really use the advice of a propulsion engineer to separate practical from unattainable. 

I'm no rocket scientist, so I have no idea if an engine can be detuned like that.  Just throwing some ideas out there.  Maybe the larger combustion chamber would not work properly if there's too little propellant being pushed through it.  I only base it on the fact that most of the RD family can be throttled down to about 30%.  Other engines even moreso like the CECE.
if you can run 10% or 30% of propellent through a larger engine, then I don't see why you couldn't just put a smaller turbo pump on one and "detune" it down, so that the same basic engine came be used for commonality.
As far as a booster engine being used as an upper stage...Merlin 1 and NK33/43 are booster engines used as upper stage engines.

I'm guess it would work, but maybe it wouldn't work well, and would be better to go with a more specific built engine.  *shrug*

[Pipcard]

Is "fully reusable super-heavy two-stage single-core-only methalox rocket with common engines" now considered to be the ultimate commercial design?

[MATTBLAK]

1x 'family,' but 2x versions of a fully reusable, two-stage LOX/RP1 or LOX/Methane booster capable of placing 20 and up to 50-to-60 metric tons into high-inclination, low Earth orbits. Also able to deliver the largest Communications 'birds' to Geostationary orbits. The 50+plus ton rocket would also have the ability to use a large payload fairing for manned Exploration-class spacecraft and propulsion modules.

I guess I'm describing close analogues to Falcon 9 and a 'single stick' equivalent to Falcon Heavy. Actual turnaround and rational reflight rate to be advised by smarter folk than me. BUT: I could easily imagine an 'affordable', high flight-rate booster being used to assemble a manned Mars expedition pretty quickly, launching from two or three pads before a Mars launch window opened. This is not intended as an SLS slam - but such a booster would be an SLS-killer...

[su27k]

Is "fully reusable super-heavy two-stage single-core-only methalox rocket with common engines" now considered to be the ultimate commercial design?

Two leading commercial space companies are both pursuing designs like this, so I guess the answer is Yes.

[dror]

Is "fully reusable super-heavy two-stage single-core-only methalox rocket with common engines" now considered to be the ultimate commercial design?

Its the best design with current technology but I don't think it's ultimate.
If price goes near the price of fuel than ultimately you need to deviate from the rocket equation.
So I think ultimately somthing like a methalox hypersonic jet for first stage will allow the same payload to orbit with much less fuel and shorter turnaround time.

[AncientU]

Is "fully reusable super-heavy two-stage single-core-only methalox rocket with common engines" now considered to be the ultimate commercial design?

Without the 'super-heavy' -- yes IMO.
The market for launching unique design/orbit <10t sats will never go away, so highly efficient smaller launchers (probably as you describe configuration above) will have a place in the space economy for foreseeable future.

[Katana]

Is "fully reusable super-heavy two-stage single-core-only methalox rocket with common engines" now considered to be the ultimate commercial design?

Without the 'super-heavy' -- yes IMO.
The market for launching unique design/orbit <10t sats will never go away, so highly efficient smaller launchers (probably as you describe configuration above) will have a place in the space economy for foreseeable future.

But wouldn't they become obsolete or unneeded like Falcon 1 did? What about the launch industry shifting to a container ship-like business model? ("generic launch" as Space Ghost 1962 calls it, in which it is okay to launch a 1-tonne satellite on a full RLV for 100 tonnes)

Or:
Low cost expendable mini electric upperstage, one per one satellite when launching multiple satellites.

Allows relocating of different satellites to different orbits when using one large launcher.

Elrctric propulsion may have more develop potential than chemical.

[Pipcard]

(That post mentioning "generic launch" was deleted due to the keywords "unique design/orbit")