A fast route to a European low cost, reusable, manned launch vehicle.

GW Johnson · 2023-06-08 15:48:22

Up to a certain size, most all the large solid makers can cast a non-segmented booster. Beyond a certain size (like the Shuttle SRB's), there's not many companies that can make a mix large enough to do an unsegmented design. Maybe 1 or 2 in the entire world. Thiokol was not one of them, which is why the SRB's were segmented, introducing multiple possible failure modes, one of which finally killed a crew.

The GEM 63XL motors are listed as 1.61 m (63.2 inch) diameter and 22.0 m (865 inches) long. They are to be used on the Vulcan Centaur, being a stretched version of the GEM 63's used on Atlas-5. Loaded motor weight is 53,400 kg (117,700 lbm), and propellant weight is 48,000 kg (105,800 lbm). Thrust is listed as 2026 KN (455,400 lb) with an 84 sec burn time. Vacuum Isp is listed as 280 sec, so I presume the thrust is vacuum, but it does not say.

The thing is a carbon fiber - epoxy composite case that is wet-wound around a slowly-spinning mandrel. The fiber lay directions are multiple, because one must "play God" and resist stresses that are different in all 3 directions at every element of the case. That used to be more expensive than plain steel cases, but the technology has matured, and it is now cheaper for those experienced at doing it.

They are actually pretty good motors! Each is made in a single cast.

GW

PS - a crude estimate says the sea level thrust is about Wp*Isp/tb = 353,000 lb (1570 KN). That means the exit area is pretty close to 6966 sq.in (4.49 sq.m). That would be 71.06 inch dia (1.80 m dia). Yet the picture does not show a nozzle exit that large. So there is some thrust profile with time. Max is not average.

Last edited by GW Johnson (2023-06-08 15:56:10)

RGClark · 2023-06-09 22:37:43

Got this response when I asked ChatGPT what is the difference between “Old Space” and “New Space”:

______________________________________________

The terms "Old Space" and "New Space" are not widely recognized or universally defined. However, they may refer to different eras or approaches within the space industry.

"Old Space" could be used to describe the traditional space industry that emerged during the Cold War era, dominated by government space agencies like NASA and Roscosmos. This period focused on large-scale, government-funded space exploration and scientific missions.

On the other hand, "New Space" typically refers to the more recent emergence of private companies and startups in the space industry. These companies, such as SpaceX, Blue Origin, and Virgin Galactic, aim to commercialize space activities, reduce costs, and develop reusable space systems. They have introduced innovative approaches to space exploration, satellite launches, and space tourism.

While "Old Space" often involves national space programs with substantial government involvement, "New Space" emphasizes private sector initiatives and entrepreneurial efforts to expand the boundaries of space exploration and exploit its commercial potential.

It's important to note that these terms are not universally defined, and their meaning can vary depending on the context or perspective of the discussion.
______________________________________________

A fairly reasonable explanation. I would only quibble about the not universally recognized part. That’s probably true about the general public but anyone in the industry and anyone familiar with the industry would know what they mean.

The key problem though with these big government-paid projects is the Old Space companies serving them believe the costs also need to be big - and they’ll do whatever it takes to keep them so.

Bob Clark

Last edited by RGClark (2023-06-10 03:38:48)

RGClark · 2023-06-10 04:01:42

tahanson43206 wrote:

For RGClark re #19
Thank you for giving my proposal some thought!
Please consider NOT jettisoning the SRB's. It costs money to jettison the SRB's. It costs more money to make new casings than to refill the old ones.
…

It would be difficult to retain the empty SRB casings. The reason is they are so heavy. Remember for a rocket you want to minimize the dry weight of a stage, that’s why you drop off the prior stages. The dry mass of the Ariane 5/6 core is in the range 12 to 14 tons. If you kept the dry mass of the SRB’s that’s about another ~14 tons for each SRB, tripling the dry mass that had to be accelerated by the core stage engines. The payload capability would be greatly reduced.

It might work if the Ariane 6 SRB’s were small as is usually the case; commonly in other cases they’re at 1/10 the core stage mass. But if the SRB’s were that small the Ariane 6 couldn’t make its desired payload.

Robert Clark

tahanson43206 · 2023-06-10 06:09:17

For RGClark re #28

Thank you for taking up my question.

The goal here is 100% reusability. The Europeans have shown they are unable to achieve that due to the consensus nature of their organization.

As an American, you are not limited in the range of thought you can explore, so it is possible for you to imagine using the liquid core stage as a reusability device.

At lift off, the total mass of the stage would be unchanged. The total thrust of the stage would be unchanged.

What I am proposing is that instead of exhausting the propellant from the core stage, a portion should be reserved to return the vehicle to launch point.

The total mass launched to orbit will be reduced, but at this point we do not know the amount of reduction. What we ** do ** know with certainty is that the cost of reworking the stage components back at the launch site will be less than the cost of building new ones.

I do not know (obviously) what your capabilities may be, but there are others alive on Earth today who might be willing to help if you get stuck.

It would be a great service to the Europeans if someone could show them how to save large sums by adding reusability to the first stage of their existing equipment.

Thank you again for taking up the question!

(th)

Mars_B4_Moon · 2023-06-16 03:31:12

Final launch of Europe's powerful Ariane 5 rocket delayed indefinitely

https://www.space.com/ariane-5-rocket-f … unch-delay

RGClark · 2023-06-19 14:49:06

Towards a revolutionary advance in spaceflight: an all-liquid Ariane 6.
http://exoscientist.blogspot.com/2023/0 … ce-in.html

Most rockets have payload fractions in the range of 3% to 4%. The Ariane 6 using 2 and 4 SRB’s, because of the large size of the SRB’s and because solids are so inefficient on both mass ratio and ISP, the two key components of the rocket equation, it will count among the worst rockets in history at a payload fraction of only 2%.

In contrast a two Vulcain Ariane 6 could have a payload fraction of 7% and a three Vulcain Ariane 6 could have a payload fraction of 7.5%. This is well-above what any other rocket has ever achieved in the history of space flight.

To put this advance in perspective, it would be like SpaceX using the very same Merlin engine and the very same propellant tanks, and the very same size Falcon 9, suddenly being able to change the Falcon 9 payload from 22 tons to 40 tons.

It will be a paradigm shift in what payloads rockets should be able to deliver to orbit.

Bob Clark

Last edited by RGClark (2023-06-20 08:25:05)

GW Johnson · 2023-06-19 19:47:15

I would remind everyone that in the first 3-5 minutes of a launch to Earth orbit, the prime issue is raw thrust, not Isp.

It is only later as the air thins to nothing and the trajectory has bent nearly horizontal, that Isp is the prime factor.

Early on, you are pushing very heavy weights very nearly straight up, and fighting lots of aerodynamic drag, too. Neither is true in the second phase when Isp is paramount.

There's much more to it than just MR and Isp from the rocket equation. Look also at ignition thrust /weight ratio. If that is not at least 1.3, your payload fraction goes down, because you are simply not strong enough to lift the weight and fight the drag AND still accelerate!

1.5 is far better. Numbers higher than 3 have been seen, but ONLY with solids, and I mean ALL-solids! The ICBM's and similar. Also intercept SAM's. Lighten the payload on any ICBM and it becomes a satellite launcher. That's how the space launch thing got started in the first place.

The frontal thrust density is just higher with the solids in any practical design. Plus, there's no time-consuming checkout and verification expenses that nobody here on these forums wants to recognize as a cost. It's a wooden round. You hit the "go" button. That's all.

The liquids have the Isp, no doubt about that. But there is no fundamental reason (other than a bad design) why SRB's on a liquid cannot get you loads of takeoff thrust, right when you need it, and for a lower price than anyone here wants to admit.

GW

kbd512 · 2023-06-20 04:24:28

Is there an issue with understanding how total impulse affects burn-out velocity for a launch vehicle with a given total mass?

6 of the GEM-63XL solid boosters provide the same total impulse and ~23.5% more thrust than both P-120 solids due to greater nozzle-to-casing diameter / area, for about $10M less than the price of a single P-120 booster, at the expense of a higher inert mass fraction (enclosing an almost identical propellant mass in a greater number of casings). Both use filament-wound CFRP casings, HTPB propellant, are "single-mix / single-pour" from what I understand, and can be had with or without TVC. The Vega-C and Ariane-6 launch vehicles share the use of P-120C boosters to reduce costs, but it takes 33 days to wind a single P-120C casing.

RGClark is fixated on a reprise of Delta-IV, which never had anything like the payload performance he's touting. He doesn't seem to want to accept that the booster stage's job is to clear the atmosphere and get the vehicle pointed downrange so that the upper stage can pour on the speed without the aerodynamic drag and heating penalties associated with accelerating through the atmosphere. Remove too much thrust and you spend significantly more time in the lower atmosphere, where energy is sapped by drag. RS-68A has been available for awhile, and it has as much thrust as 2.75 Vulcain 2.1 engines, but would add 2.84t of inert mass over 3 Vulcain 2.1 engines. Adding 2.84t of booster mass over the top of 3 Vulcain 2.1 engines would not dramatically affect upper stage performance. If it would, then adding 2.6t of inert booster mass in the form of 2 additional Vulcain 2.1 engines, ignoring the weight added for additional plumbing and thrust structure, would also have the same deleterious effect on overall vehicle performance as RGClark's proposed booster stage without solids, to a slightly lesser degree. Vulcain 2.1 has a much better thrust-to-weight ratio than RS-68A, no doubt about that, but that is the only advantage it confers.

No "greater magic" is going to happen by attaching 1 or 2 additional Vulcain 2.1 engines to Ariane-6's core stage. You have a lot less propellant to consume, sans solid boosters, and therefore less payload performance, plain and simple. Thrust and inert mass fraction goes up, but total impulse does not, so unless the booster / upper stage / payload gets lighter, then the launch vehicle doesn't have the same payload performance.

Pure LH2 is legacy tech from the Delta-IV Heavy launch vehicle program. No such dramatic payload performance improvement was observed back then, either. Delta-IV Heavy triple Common Booster Core stage roughly doubled total payload, but not payload mass fraction (3.14% for triple CBC vs 3.43% for 1 CBC plus 4 GEM-60s). CBC is powered by 1 LH-2fueled RS-68A, GEM-60 is HTPB solid. In point of fact, by running the numbers I can see that payload mass fraction went the wrong way without the GEM-60 solids, and CBCs cost so much that ULA retired the vehicle.

Liquid plus solid is equally well proven on Titan, Delta-IV, Atlas-V, STS, SLS, Ariane-5, Ariane-6, Vega-C, and the list goes on. Each technology set does what it does. Solids provide greatly improved thrust performance at liftoff to allow otherwise anemic LH2-powered vehicles to get out of their own way long enough to provide overall lower liftoff mass, which is why NASA started using it to begin with (the assertion being that total vehicle weight drives cost, which is true to a point). We've fixated on that to the point that now the lowest possible weight greatly increases cost through the use of exotic materials and manufacturing solutions, so we've come full circle, right back to where we started with our cost vs weight conundrum.

LCH4 is the lowest cost / highest total performance booster propellant and engine technology, but is itself another compromise solution that trades the specific impulse advantage of LH2 for greater thrust. Kerosene is a very close second, and so close that LCH4 doesn't offer much of a material advantage. Apart from autogenous propellant tank pressurization to avoid the weight / expense / complexities of external pressurization systems, kerosene is the easiest liquid fuel to work with. Solids offer the highest frontal thrust density performance for the reasons GW already listed off (namely, exit nozzle surface area to frontal stage surface area ratio), but when they get large enough they run into costly manufacturing and transport issues.

I would like to see development of more compromise solutions, such as LOX / paraffin wax (as a potential PBAN / APCP / HTPB successor) and LOX / Propane (as a potential LCH4 successor) and in-space / upper stage propulsion with HAN-based propellants to replace the various Hydrazine-based propellants. We've already taken, LH2, LCH4, kerosene, and traditional solids about as far as we reasonably can. The only dramatic performance improvements to be had will require much more exotic engines, such as hybrid solids, SABRE (liquefying O2 from the atmosphere using LH2 as a heat sink- pretty much the only way you'll ever see a practical SSTO), supersonic flow turbo machinery cores (RAMGEN), ramjets (we've spent decades trying to develop these, and to the extent that they're usable at all they tend to get used in weapons more than spacecraft), and rotating pulse detonation wave engines (the only meaningful performance improvement over an ordinary LH2 rocket engine). When all of those technologies have been fully developed and deployed aboard operational vehicles, chemical propulsion technology has reached its zenith.

BNNT composite propellant tanks with SABRE technology combined with a RAMGEN turbofan engine core and pulse detonation wave rocket engines to take over outside the atmosphere is the level of airframe structure and engine tech required for a SSTO that comes close enough to the performance of any existing TSTO to eliminate the requirement to use TSTO to achieve a decent payload mass fraction. Even then, a TSTO will have a higher payload mass fraction and be more economical to operate if the booster is using the same materials and engine tech and is a VTVL design. Basically, SpaceX has the right idea. If input energy to synthesize propellant is not a significant limitation, then a big dumb booster is likely to remain the most economical to operate in the long run. Skylon D1 (16.4% inert mass fraction) would have a 4.6% payload mass fraction. All those technologies I mentioned combined into a single Skylon-like vehicle could just about provide a 10% payload mass fraction in a practical SSTO design that takes off and lands like a conventional airliner. A rocket can still do better. STS achieved a 6.49% payload mass fraction (note that it also used solid boosters). The TSTO Saturn-1B achieved a 4.76% payload mass fraction using late 1950s era technology, which is already above where the as-designed Skylon D1 would be, assuming none of the even more advanced engine and materials tech, if it ever becomes operational. Skylon already is another STS-like boondoggle. Yes, it would work if the engineering is on point, but no it would never reduce cost to orbit by one red cent. That hasn't stopped anyone from spending the money to find out the hard way. There's always someone who thinks we can make one minor tweak or change that will dramatically improve cost or capability. There are very rare cases of that happening, but most of the time cost only goes up and capability only goes down.

The only way forward from our most advanced combustion engineering technologies involves electrical / electromagnetic / plasma / nuclear / warp drive propulsion schemes primarily intended for dramatically more efficient in-space propulsion.

Fixating on a non-competitive project to transform Ariane-6 into something it was never designed to be is a waste of time and money. ESA / ArianeGroup is better off designing a new family of partially or fully reusable rockets to compete with Falcon-9 and Starship, else there is no international competitive market for launch vehicles. Heck, even China figured this out. SpaceX also thought "strapping on booster cores" to a Falcon-9 would be "quicker and easier" than designing a new rocket. In the end, they wound up with a completely redesigned core stage to accommodate the strap-on booster cores, at which point Falcon Heavy had very little in common with a Falcon-9 booster core except the engines. Everything else was a complete redesign. Why repeat a course of action that didn't work for Falcon-9 and likely won't work here, either?

tahanson43206 · 2023-06-20 06:19:02

For kbd512 re #33

Glad to see you feeling better after your recent serious sounding health scare....

SearchTerm:Solid vs Liquid analysis of space craft design - includes advantage of raw power to accelerate from ground level.

SearchTerm:Liquid vs Solid

(th)

GW Johnson · 2023-06-20 11:47:03

I just posted a "how-to" article for TSTO launch over at my "exrocketman" site. It is titled "TSTO Launch Fundamentals", and is dated today as 6-20-2023.

The article is fairly short and has only 2 figures in it. These show where all the numbers must come from, and what Newton's 3rd Law must look like from the free-body diagrams at 3 typical points along the ascent trajectory. The two burns have different Isp values and different weight statements. Choice of propellants may vary between stages. The payload fraction at launch optimizes with a search for "best" staging velocity.

I did include a paragraph about what to do when adding solid SRB's. But that was out-of-scope for the basic article. It does split the first stage burn into two different effective first stage burns, each with a different Isp and weight statement, plus the second stage burn separate from all of that.

The notion of a single rocket equation calculation modeling how all this fits together is just UTTER NONSENSE! It is an iterative search, and there is NO general rule-of-thumb. The result is unique to each different design.

Maybe we can get Tom to post this thing in "dropbox" for us, since he says a lot of y'all have difficulty finding and looking at stuff posted on my "exrocketman" site.

GW

Last edited by GW Johnson (2023-06-20 11:51:43)

tahanson43206 · 2023-06-20 11:55:31

For GW Johnson re #35

Happy to oblige .... Just send the pdf and the images by email, and I'll set them up for viewing from the GW Johnson Postings topic.

(th)

RGClark · 2023-06-21 07:09:39

kbd512 wrote:

Is there an issue with understanding how total impulse affects burn-out velocity for a launch vehicle with a given total mass?
6 of the GEM-63XL solid boosters provide the same total impulse and ~23.5% more thrust than both P-120 solids due to greater nozzle-to-casing diameter / area, for about $10M less than the price of a single P-120 booster, at the expense of a higher inert mass fraction (enclosing an almost identical propellant mass in a greater number of casings). Both use filament-wound CFRP casings, HTPB propellant, are "single-mix / single-pour" from what I understand, and can be had with or without TVC. The Vega-C and Ariane-6 launch vehicles share the use of P-120C boosters to reduce costs, but it takes 33 days to wind a single P-120C casing.
RGClark is fixated on a reprise of Delta-IV, which never had anything like the payload performance he's touting. He doesn't seem to want to accept that the booster stage's job is to clear the atmosphere and get the vehicle pointed downrange so that the upper stage can pour on the speed without the aerodynamic drag and heating penalties associated for Falcon-9 and likely won't work here, either?

In regards to pricing of the solids, you have to consider the GEM 63XL is about 1/3rd the size of the Ariane 6 solids. The GEM 63XL has been estimated to cost from $5 to $7 million.

Here it’s estimated as $5 million:

Derek Newsome @DerekdotSpace
The AJ-60A typically cost around 7 million for Atlas V, with the replacement GEM-63 being slightly cheaper, but no set price has been given.

GEM-63XL is ~5 Million for Vulcan Centaur.
11:25 AM · Feb 22, 2023 · 2,019 Views 9 Likes 1 Bookmark
https://twitter.com/derekdotspace/statu … 57377?s=61

Here it’s estimated as closer to $7 million:

I believe we have enough information to determine the price of Vulcan.
…
Final Price Range
Vulcan 504: $98-$100 million
GEM-63XL: $6.67-$7 million
https://www.reddit.com/r/ula/comments/8 … utm_term=1

Based on the 3 times greater size than the GEM 63XL, we can estimate the SRB’s on the Ariane 6 as $15 to $21 million each. So $30 to $42 million for the two on the Ariane 62. This is in the range of the price I estimated in comparing how the price increased by adding two additional SRB’s in going from the Ariane 62 to the Ariane 64.

It is solely the price of the large SRB’s used on the Ariane 6 that accounts for why it is so expensive.

The idea of the importance of impulse as the most important factor is probably coming from amateur solid rocket makers, including those working with high power rocketry that mix their own professional solid propellants such as APCP. For them, they don’t care about how lightweight the solid rocket is. All the rockets are about the same mass ratio. They only care about how much propellant they can get into the casing.

But for orbital rockets the most important thing is not impulse, it is delta-v. And this depends on Isp and mass ratio. Because of mass ratio, essentially the ratio of propellant mass to empty mass, making the stage lightweight is of great importance.

The mass ratio for the Ariane 5 is quite extraordinary for a hydrolox stage at 16.3 to 1. This is in the range commonly seen by dense propellants. To use a colorful analogy, it’s like the ArianeSpace engineers found a way to make liquid hydrogen as dense as kerosene!

Obviously, this is not what happened. But they must have found a way to achieve extreme lightweighting of a hydrolox stage. To put this in perspective, the best mass ratio previously for a hydrolox stage was by the famous Centaur hydrolox upper stage at ~10 to 1, achieved back in the 1960’s. And the Delta IV hydrolox core is a quite ordinary 8.7 to 1. So the Ariane 5 core is about twice as good as the Delta IV core on this key mass ratio scale.

Because the Ariane 5 core has the high ISP of a hydrolox stage while achieving (somehow!) the high mass ratio of a dense propellant stage, it calculates out to have the highest delta-v of any rocket stage in the history of spaceflight.

Since delta-v is the single most important parameter for orbital rockets, you can legitimately say the Ariane 5 core is the greatest rocket stage ever produced in the history of spaceflight.

Robert Clark

Last edited by RGClark (2023-06-21 14:07:34)

RGClark · 2023-06-21 07:30:28

GW Johnson wrote:

I would remind everyone that in the first 3-5 minutes of a launch to Earth orbit, the prime issue is raw thrust, not Isp.
It is only later as the air thins to nothing and the trajectory has bent nearly horizontal, that Isp is the prime factor.
Early on, you are pushing very heavy weights very nearly straight up, and fighting lots of aerodynamic drag, too. Neither is true in the second phase when Isp is paramount.
There's much more to it than just MR and Isp from the rocket equation. Look also at ignition thrust /weight ratio. If that is not at least 1.3, your payload fraction goes down, because you are simply not strong enough to lift the weight and fight the drag AND still accelerate!
1.5 is far better. Numbers higher than 3 have been seen, but ONLY with solids, and I mean ALL-solids! The ICBM's and similar. Also intercept SAM's. Lighten the payload on any ICBM and it becomes a satellite launcher. That's how the space launch thing got started in the first place.
The frontal thrust density is just higher with the solids in any practical design. Plus, there's no time-consuming checkout and verification expenses that nobody here on these forums wants to recognize as a cost. It's a wooden round. You hit the "go" button. That's all.
The liquids have the Isp, no doubt about that. But there is no fundamental reason (other than a bad design) why SRB's on a liquid cannot get you loads of takeoff thrust, right when you need it, and for a lower price than anyone here wants to admit.
GW

Because of the high thrust solids are good at liftoff, but only if they are small. And this is commonly the case like the solids used for the Atlas 5 or Delta IV, where the solids were only about 1/10th the size of the core stage. They become a poor choice when they are made large like happened with the Space Shuttle and SLS. Their large size, each comparable to the size of the entire core stage mass, contributed to both of those being financial disasters.

The large SRB’s on the Ariane 6 also are financial disasters. They are also each comparable in size to the entire core stage. They are the sole reason why the two versions of the Ariane 6 are so expensive, which will likely lead to ArianeSpace going bankrupt if they go ahead with fielding them at 2 to 3 times higher price than the going rate of the Falcon 9’s used price.

Bob Clark

Last edited by RGClark (2023-06-21 07:31:39)

Calliban · 2023-06-21 10:47:07

This sounds like a good argument for a very simplified, pressure-fed booster burning LOX/kerosene and using a non-regenerative water cooling system, with steam dumped over-board for extra thrust. We could mount the booster on wings and glide it back to a runway for landing and reuse. A hybrid using a cast rubber based fuel is a good contender as well.

kbd512 · 2023-06-21 15:37:04

RGClark,

But for orbital rockets the most important thing is not impulse, it is delta-v. And this depends on Isp and mass ratio. Because of mass ratio, essentially the ratio of propellant mass to empty mass, making the stage lightweight is of great importance.

The fundamentals of rocketry apply to all rockets at all times.

From NASA:

The total impulse (I) of a rocket is defined as the average thrust times the total time of firing.

Total impulse DOES determine your Delta-V or "burnout velocity", if all other factors are held constant, and here's why:

3 engines of the same type, producing the same amount of average thrust, are consuming propellant 3 times as fast as 1 engine of that same type. So long as the thrust-to-weight ratio of the complete vehicle is greater than 1, and preferably 1.5 times total vehicle weight or greater, that means 3 engines will accelerate a vehicle with the same form factor a bit faster, but they cannot substantially alter the booster burnout velocity. A minor alteration in burnout velocity will take place due to the rate of change of atmospheric drag since the rocket is ascending through the atmosphere faster, but not because more total energy is available to accelerate.

Again, assuming TWR > 1.5 or so, the figure of merit is the energy applied over a given period of time (predetermined by available propellant load and engine thrust), assuming the goal is to substantially increase burnout velocity. If you generate more thrust and therefore accelerate faster using 3 engines vs 1 engine (a more rapid acceleration rate), you also burn through an identical propellant load 3 times as fast, so the pair of additional engines can only provide the same thrust as 1 engine for 1/3rd as much time, unless you make the rocket substantially heavier by increasing its total propellant load, and therefore its total impulse capability.

Total impulse is what produces a given amount of force, operating on a vehicle with a given initial (all propellant, all structural weight, all payload weight) and final mass (remaining vehicle weight after all booster propellant has been expended and the booster(s) staged-off, leaving the upper stage to provide most of the velocity to orbit the payload). You do NOT get more total impulse by strapping two more engines to the back of your booster's propellant tank. Your rocket is now modestly heavier, and it should now have 2/3rds more thrust to accelerate faster, but the energy contained within the propellant hasn't changed one iota unless you increase the size and weight and drag of the propellant tanks.

Therefore, total impulse is also the total amount of force available for pushing the vehicle downrange, predetermined by the weight of the propellants and the thrust produced by the engine(s) consuming those propellants. Along with initial / final vehicle weight / thrust, total impulse ultimately determines what your burnout velocity will be, aka what your "Delta-V" or "change in velocity" will be.

You add engines to increase thrust, because if your thrust is less than 1.5 or so, then either you don't leave the ground at all if TWR < 1 or you don't accelerate fast enough to avoid burning through all of your propellant before exiting the atmosphere. That is what the pair of solid rockets are primarily there to do (increase liftoff thrust to a point where acceleration is quite rapid). Additionally, they also add a healthy amount of additional total impulse, thereby increasing burnout velocity or "Delta-V".

Calliban · 2023-06-21 15:48:34

Kbd512, the dV associated with gravity losses, is a linear function of the firing time.

V = A x t

V = (A - g) x t

Purely as example, if firing time with one engine is 300 seconds at 1.5g:

V = (15 - 10) × 300 = 1500m/s

If firing time is 100 seconds at 4.5g:

V = (45 - 10) x 100 = 3500m/s

One engine accelerating the craft at 1.5g will not give you the same burn-out velocity as three engines burning for one-third of the time. Gravity losses are minimised by accelerating at the highest rate possible. But of course bigger engines are heavier, so there is a trade off.

Last edited by Calliban (2023-06-21 15:53:37)

kbd512 · 2023-06-21 17:08:06

Calliban,

The gravity losses are no more than 10% or so, according to GW. Any change by switching engines in a practical vehicle will be minor.

TWR of the as-built Ariane-6 A62 configuration (10.671MN / 1,088,139kgf of thrust and 530t of weight), is 2.05:1.

Leaving off the pair of P120C boosters from the Arian-6 A62 configuration removes 309.2t of weight and 9.3MN of thrust from Ariane-6.

Your vehicle weight drops from 530t to 220.8t, and thrust drops to 1.371MN / 139,803kgf (not enough to leave the ground). This is the vacuum thrust of Vulcain 2.1, so sea level thrust will be less, not more.

You add 2 more Vulcain 2.1 engines, which weigh 1,650kg each, bringing your new vehicle weight up to 224.1t and thrust up to 4.113MN / 419409.3kgf.

TWR at liftoff is now sufficient, at 1.871:1, so the rocket is definitely leaving the ground when powered by 3 Vulcain 2.1.

Houston, we have a problem!

You actually lost TWR by switching to 3 Vulcain 2.1 (1.87:1) over 1 Vulcain 2.1 and 2 P120C solid boosters (2.05:1).

You will get zero performance improvement by doing that. In point of fact, you will lose performance.

Now what? 4 Vulcain 2.1 engines? Now you're burning through propellant 4 times as fast, so you're thrusting for 1/4 the amount of burn time of the single Vulcain 2.1. Think that'll help? It won't. Either redesign the vehicle to use 4 Vulcain 2.1 engines, or there's no point.

Sorry, but this doesn't work.

RGClark · 2023-06-21 18:40:03

You can not ignore delta-v for an orbital rocket. Getting sufficient delta-v for orbit is the entire point of the rocket equation:

Robert Clark

kbd512 · 2023-06-21 19:25:49

RGClark,

Do you obtain Delta-V by shoving combusted propellant out the back of a rocket engine?

If so, do you get even more Delta-V by shoving even more combusted propellant out the back?

If you answered both the first and second questions in the affirmative, then do you see any potential problems with having a lot less propellant to work with?

GW Johnson · 2023-06-21 21:45:50

What Calliban tried to say about gravity losses is true, except that he needs to consider vector, not scalar, math. The component of gravity acceleration that is parallel to the path is the one you multiply by time for gravity loss. The perpendicular component is what bends your thrusted gravity turn trajectory, and that is not a loss.

What Kbd512 said about total impulse is correct, precisely because total impulse divided by burn time literally is thrust, and because total impulse divided by specific impulse literally is propellant mass.

What Bob Clark said about delta-vee (dV) is correct. That literally is the result of the rocket equation. There is a caveat: it must overcome the gravity and drag losses, as well as reach the orbital energy requirement.

What I have been trying to point out is that all these things (and more besides) are true, but they do NOT always have equal importance!

In a TSTO system with a properly-optimized staging velocity, the initial first stage burn incurs all the drag loss, and most if not all the gravity loss, because it flies mostly straight up for over half the burn. Under these circumstances, thrust above and beyond drag plus weight is of paramount importance. The drag and the gravity losses figure into the dV demanded of the stage. Isp affects that, being a rocket equation calculation. But it is that Newton's 3rd Law force sum = mass x acceleration that is of overwhelming importance.

During the second stage burns (yes, plural), the trajectory is exo-atmospheric, and almost but not quite locally-horizontal. There is no drag loss, very little (if any) gravity loss, and the delta vee is from stage velocity to what amounts to the surface circular orbit speed, in order to get not only the kinetic but also the potential energy needed for the desired low circular orbit. That's the basic 2nd stage burn dV.

It is unlikely that second stage burnout occurs just as you reach orbit altitude, so you coast up to that altitude, then conduct a modest circularization burn, on the order of 100 m/s = 0.1 km/s worth of dV. The second stage usually does that circularization, so if it does, you just add that circularization dV to your required dV (they come out of the same weight statement, so you can sum dV's for the rocket equation -- but if the weight statement changes, you CANNOT sum dV's for one rocket eqn calculation, you MUST do separate calculations for each burn from a separate weight statement). Thus for the 2nd stage, it's all about the rocket equation, and not really much about Newton's 3rd Law, except that acceleration and burn time are related.

Weight statement changes could be either changes in payload or changes to inert mass (like dropping a shroud), or both. The notion here is that inert + payload = burnout, and inert + payload + propellant = ignition.

Bob's arguments really apply far more to the second stage burn. Kbd512's arguments really apply far more to the 1st stage burn. But you have to do both stages, it's a TSTO situation. It's just complicated.

As for the costs of solids, they can be quite cheap. They can also be quite expensive, if you kluge-up the design with unneeded things like segmented cases, or so large a motor you cannot cast it plus burn rate samples from a single mix. (That is EXACTLY the expense problem with the Shuttle and SLS and Araiane-6 SRB's by the way.) Flexible-joint thrust vector control nozzles are expensive and add weight, but sometimes they are necessary.

Although it appears largely forgotten today that attitude control jets can often do the same steering function on a solid (and did so quite well on "Scout"). Such can do it with something way easier to design and build than a flexible-joint TVC nozzle rig. And quite probably lighter, too, although sometimes not. Depends on what kind of thruster rig you use. But "way easier to design and build" is almost invariably cheaper!

GW

Last edited by GW Johnson (2023-06-21 22:07:52)

kbd512 · 2023-06-22 09:28:18

GW,

Yes, my arguments were intended to apply specifically to the first stage of Ariane-6, because that is the topic of discussion in this thread. We've been arguing over removing the solids and adding more liquid engines to the Ariane-6 core stage, in order to reduce cost. I think that's a mistake, because as the simple math shows, you lose almost 10% of your thrust unless you add a 4th Vulcain 2.1 engine. Performance will suffer as a result. That doesn't really help us accelerate any faster, does it? In the weight calculation I added for the 2 additional Vulcain 2.1 engines, no additional weight for propellant feed lines or thrust structure were included. Maybe it'll add nothing in terms of weight because the weight associated with the propellant tank attachment points for the solid rocket boosters can be rolled back into the propellant feed lines and thrust structure for the additional Vulcain 2.1 engines.

We have a larger problem, though, don't we?

By removing the solids and adding 3 more Vulcain 2.1 engines, we just completely redesigned the core stage into something it was never designed to be, since 1 Vulcain 2.1 engine was what the rocket was designed to use.

We're better off starting from scratch and designing a new core stage, even if we stick with the Vulcain 2.1 engine, which lacks the thrust to be a terribly useful booster engine.

If Delta-IV is any indicator, removing the solids and adding more LH2 engines and propellants hurts payload mass fraction and cost. In terms of dollars spent, we received more "bang-for-the-buck" by strapping 4 GEM-60s to the Delta-IV Medium core. A Delta-IV Medium (4 GEM-60 solids) gave us a 13.5t class payload for around $164M a pop. A Delta-IV Heavy (no solids), gave us 1 23t class payload for around $400M a pop.

How did America and now China solve that problem?

It wasn't using solids or LH2. SpaceX started using LCH4. ULA / Blue Origin switched to LCH4 as well. RP-1 thrust from cryogenic propellants at the same working temperature, but without the manufacturing and transport issues associated with solids or the soot and operating temperature limitations of RP-1. Problem solved. Even at that, we're still using lots of solids because they still have a role to play. LH2 is the best propellant available for upper stages, where Isp is much more critical. It's pretty hard to argue with that. As a booster propellant, in all actually-built designs, LH2 seems to leaves a lot of thrust performance on the table when compared to solids or RP-1 or LCH4. That seems to be why solids are so frequently paired with LH2. NASA / ESA / ISRO / JAXA all pair solids with liquids. I think China and Russia are the only space-faring nations that almost entirely rely upon liquids (hypergolics or RP-1) for the bulk of their orbital launch vehicles not based upon repurposed ICBMs.

Here's a thought, though:

Instead of screwing around with another brand new rocket design, how about ESA / ArianeGroup creates a Mercury-like single astronaut space capsule capable of their very first crewed flight?

Mercury was a 1.3t capsule, or thereabouts. Modern materials should allow us to reduce that weight further. Vega-C can deliver 2.2t ot LEO.

I'd much rather see ESA join the crewed space flight community with their first crewed spacecraft, than try to catch up to where SpaceX is in the commercial market, which would mean more billions down the drain on vehicle concepts that were fine for 10 or 20 years ago, but no longer cost-competitive. They can still pursue this after clearing the next logical hurdle to tackle, which is crewed space flight.

America, Russia, China, India in 2025, and potentially JAXA, will have this capability in the near-future. I think it's past time for ESA to join the club. They need to demonstrate crewed space flight using their own launch vehicle, their own space capsule, and their own mission control. After they do that, then go after the commercial market. Heck, they already know how to design the service module of a very large capsule. Finish that "last mile" of capability to prove to everyone that you can.

I would like to see competitors to Starship, because that will become the future of space flight- enormous low-cost rockets pushing enormous payloads into space, followed closely by "large ship" concepts like the one proposed by RobertDyck.

GW Johnson · 2023-06-22 18:06:12

Kbd512:

Every point well-taken. I quite agree.

GW

RGClark · 2023-06-22 23:29:06

If Delta-IV is any indicator, removing the solids and adding more LH2 engines and propellants hurts payload mass fraction and cost. In terms of dollars spent, we received more "bang-for-the-buck" by strapping 4 GEM-60s to the Delta-IV Medium core. A Delta-IV Medium (4 GEM-60 solids) gave us a 13.5t class payload for around $164M a pop. A Delta-IV Heavy (no solids), gave us 1 23t class payload for around $400M a pop.

The size of SRB’s used also has to considered. It’s the size of the SRB’s on the Ariane 6 that make them so expensive, resulting in the Ariane 6 itself being so expensive. Solid side boosters are commonly 1/10th the size in mass of the core stage they’re attached to. This is true of the Delta IV and Atlas 5 SRB’s, and continues to be true with the new Vulcan Centaur with its GEM 63XL side boosters.

Notable exceptions were the Space Shuttle system and the Space Launch System with their huge SRB’s, with both systems being financial disasters.

The SRB’s on the Ariane 6 instead of being 1/10th size of the core are the same size as the core in mass.

Imagine what the cost of the Delta IV would be if each SRB used was made 10 times its current size.

Bob Clark

Last edited by RGClark (2023-06-22 23:53:46)

RGClark · 2023-06-22 23:50:25

kbd512 wrote:

RGClark,
Do you obtain Delta-V by shoving combusted propellant out the back of a rocket engine?
If so, do you get even more Delta-V by shoving even more combusted propellant out the back?
If you answered both the first and second questions in the affirmative, then do you see any potential problems with having a lot less propellant to work with?

As I mentioned in post #43, the key equation for orbital rockets is the rocket equation. That equation makes clear the importance of having a low empty mass, i.e., dry mass, of a stage.

Suppose you have a tank. I don’t mean a propellant tank. I mean those big lumbering things with a turret on top that shoots explosive shells. You can fit, if someone wanted to, solid propellant in the empty volume inside the tank. If you want more propellant, and more impulse, just make the tank proportionally larger. That tank is not going to serve as a stage of an orbital rocket because it is too heavy for the propellant it can carry.

Robert Clark

Last edited by RGClark (2023-06-23 01:55:18)

kbd512 · 2023-06-23 02:32:03

RGClark,

You keep repeating this line about the rocket equation, as if it means something in this particular context, but it doesn't. The empty weight of the P120C motor casing is 11t, and it's staged-off as soon as it burns out. The rocket isn't lugging the empty motor casings into orbit. The pair of P120Cs provide almost all of the thrust at liftoff. The weight of 4 Vulcain 2.1 engines alone is 6.6t. The core stage probably weighs more than an empty P120C casing. Ariane-5's 5.4m diameter empty core stage weight is 12,600kg. The new Vulcain 2.1 engine is about 350kg heavier than Vulcain 2, so a reasonable guess would be that Ariane-6's core stage is even heavier since it's using a heavier engine and the rocket is also producing more thrust and therefore has greater loads to resist. Anyway, Ariane-6's core stage carries 140t of propellant and each P120C casing carries 143.6t of propellant. That means the core stage of the Ariane-6 weighs 11.3t, excluding the engine, so slightly more than each P120C motor casing. Adding 3 more Vulcain 2.1 engines won't make the core stage any lighter, either. Your plan is to cut 287.2t of propellant and then act as if that will have no effect on Delta-V or the rest of the rocket equation, because the motor casings are "heavy" (but somehow lighter than the core stage without any Vulcain engines attached). Good luck with that.

Total impulse is what produces your Delta-V (which is created by all that propellant you intend to get rid of). Regardless of how spectacular your engine's specific impulse is, insufficient fuel sill means you're not making it to orbit. The specific impulse difference between the LH2 core stage engine and the solids is fairly pedestrian, and the entire point behind using the solids is the thrust they provide and the total impulse they add to, by tripling the available propellant at liftoff, burning through the solid propellant very quickly, and then they're staged-off.

You don't have enough propellant in the core stage of Ariane-6 to generate payload performance equivalent to what it is with the pair of P120Cs. Delta-IV proves this. Payload mass fraction took a big hit when the solids were replaced with a pair of liquid boosters in the Delta-IV Heavy configuration. That means you'll have to make the core stage bigger and add at least 3 more Vulcain 2.1 engines to match or beat the liftoff thrust-to-weight ratio of the A62's configuration. That's equivalent to redesigning the entire rocket- engines / core stage / flight control software. Heck, you have to go back and model the thermal load all the additional engines impose on each other. They already redesigned Ariane-5 into Ariane-6 and spent 3.6B Euros to do it. SpaceX spent about 300M USD to develop Falcon-9 and 90M USD in the development of Falcon 1, so it was 12 times more costly to develop Ariane-6 when compared to Falcon-9. Deja-SLS all over again. Deja-Delta-IV all over again.

You want to redesign Ariane-6 again because you think doing that will make it cost competitive with Falcon-9. How much do you figure that will cost? Does ArianeGroup have to make money by selling launches at any point? Suppose it costs another 3 billion Euros. At $65M per launch, that means they need to fly 101.5 times to pay off the development costs alone, but rockets still cost money to build and fly, even after the development has been paid for. What will that add to the launch price tag? It always costs nothing to play with other peoples' money.

I don't think you thought this one through. You originally asserted that we could simply add another engine to the core stage. That didn't provide enough thrust to leave the ground, which is why I suggested adding 2 more engines. 3 Vulcain 2.1 would at least enable us to leave the ground, but then our thrust was still lower than the existing A62 configuration by nearly 10%, so the hoped-for acceleration increase wasn't achieved. Even after fudging the numbers a bit, I still showed no payload to orbit with 3 Vulcain 2.1 engines. Adding a 4th engine gives us a TWR of 2.48:1, which is better than 2.05:1, and at least I show payload to orbit with this configuration, but it's not going to make the vehicle more performant than the A62 configuration, because total impulse is a real thing and it really matters, and you really do need more propellant to add to it.

If we do what you're proposing (without redesigning the rocket to add more propellant), I get a payload to LEO estimate from Kourou of about 3.4t, and 6.7t at most (half as much as the A62 configuration), by monkeying with the specific impulse of the core stage to make it an average of 431s and by plugging in the vacuum thrust of Vulcain 2.1 (even though we won't achieve anything like that at sea level). A core stage with 4 Vulcain 2.1 represents about double to triple the payload performance of a Vega-C (2,300kg) if you get creative with your thrust and specific impulse figures. It's going to be useless for SSO / GEO / polar launches, and you can forget about launching something to another planet. We now have a very expensive rocket that has half of its existing payload performance and very few nice-to-have capabilities.

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