Author Topic: "DIRECT" Goes Live (Read 518894 times)

Jim · « **Reply #300 on:** 11/05/2006 04:33 pm »

the RS-68 will have less ISP, note the drop in payload capability.

JIS · « **Reply #301 on:** 11/05/2006 04:40 pm »

There is difference in orbits.

ESAS say lift capability of 86.5mT with net payload 73.5mT.
DIRECT says net payload of 70.9mT

What is the difference between lift capability and net payload?
Because ESAS considers lower orbit than DIRECT I think that net payloads 73.5mT and 70.9mT are NEARLY equivalent.

mong' · « **Reply #302 on:** 11/05/2006 09:31 pm »

what I really doubt in this plan is the RS68 upgrade.
I know an upgrade to regeneratively cooled nozzle was always kept in mind during developement of the engine but could it be done for $1Bn ? and what about the individual price of the engines after such an upgrade ?

Jim · « **Reply #303 on:** 11/05/2006 09:41 pm »

I doubt it is $1b, more likely less

Kaputnik · « **Reply #304 on:** 11/06/2006 12:39 pm »

Quote

JIS - 5/11/2006 5:23 PM
What is the difference between lift capability and net payload?

In the absence of any response to this question I would assume that 'lift capability' is the total mass placed into orbit including the upper stage of the LV, whilst 'net payload' is the useful paylaod excluding the rocket stage, fairing, residual propellants, etc.

kraisee · « **Reply #305 on:** 11/06/2006 04:47 pm »

Quote

mong' - 5/11/2006 5:14 PM

what I really doubt in this plan is the RS68 upgrade.
I know an upgrade to regeneratively cooled nozzle was always kept in mind during developement of the engine but could it be done for $1Bn ? and what about the individual price of the engines after such an upgrade ?

I don't have a final cost.

I have severely over-estimated the cost at $1Bn.

Understand that the work required to upgrade RS-68 to "Regen" spec is considerably less than the work to make the J-2X - and that contract has been issued to PWR already totalling just $50 million for complete DDT&E, 7 test units, 2 engines for certification and one full-spec flight engine. And remember that J-2X is going to have an all-new nozzle. This means it is going through it's own "Regen" development program as just one *part* of that contract.

Even though RS-68 is larger than J-2X, the Regen development will cost a lot less than the entire development program of the J-2X. I have still assumed a far higher number to cover ANY eventualities, including major element re-design if, somehow, it should be necessary.

If it does prove unnecessary (probably), then NASA would simply have a lot of spare cash left over

And that's a bad thing, how?

The RS-68's major components (pumps, combustion chamber, valves, injector etc) were designed from the start to be used with a Regenerative Nozzle upgrade at some point in the future. They have actually been tested for that performance already. They are howevere FLOWN considerably under the performance limits because they are using an ablative nozzle to keep costs down to just $12m per engine for Delta-IV.

PWR indicated, completely unofficially, that $5m per unit would be the additional cost for a Regen nozzle, assuming numbers around 10-20 per year.

Add in the $8m NASA has indicated will be the extra cost for the man-rated version and you get $25m per final RS-68R man-rated unit.

Ross.

kraisee · « **Reply #306 on:** 11/06/2006 05:08 pm »

Quote

Kaputnik - 6/11/2006 8:22 AM

In the absence of any response to this question I would assume that 'lift capability' is the total mass placed into orbit including the upper stage of the LV, whilst 'net payload' is the useful paylaod excluding the rocket stage, fairing, residual propellants, etc.

JIS and Kaputnik,
The difference is the mass of the other items brought up, but which have little to do with the payload itself. Things such as payload fairings, payload mounting structures attaching the payload to the final stage, any additional push-stages (if required for circularization for example) are all factored in to the total 'lift capability', but not included in the net payload numbers.

DIRECT's Lift Capacity is also considerably higher - a little over 83mT.

DIRECT+EDS' Lift Capacity is over 110mT plus the 14.4mT EDS itself.

Ross.

kraisee · « **Reply #307 on:** 11/06/2006 07:15 pm »

The latest version of the DIRECT proposal has just been released.

v1.1.0 improves the readability of a lot of the images, and expands on the Development Phase section.

It is available on www.directlaucher.com.

Ross.

ryan mccabe · « **Reply #308 on:** 11/06/2006 07:38 pm »

Just an aesthetic note, but I like the improved banner on the directlauncher.com page. The first banner suffered from some crummy resolution. Looks much sharper.

A performance question: what is the status of the RL-60? How would performance and development/flight costs vary if a cluster of 3-4 RL-60 were used, rather than a single J2-X? The preliminary specifications for the RL-60 show much better ISP and mass efficiency than the first generation J-2. From an amature's standpoint, the RL-60 would also have some hope of application outside of the VSE.

Also, from the FAQ section of your webpage, you note the performance to LEO as the following:

DIRECT to 100x160nm 28.5deg:
70,916kg
or 156,343lb

DIRECT to 100x160nm 51.6deg:
70,907kg
or 156,323lb

Is it really just a 9 kg hit to reach a higher inclination ISS orbit? Seems very disproportionate to what the STS suffers on ISS flights

Regards

kraisee · « **Reply #309 on:** 11/06/2006 08:26 pm »

I've been looking at this again and the only reason I can come up with is the insertion point ithe initial orbit.

For Lunar CLV missions, 60x160nm 28.5deg orbit, the insertion point is at an altitude of 123nm (227km).

For ISS CLV missions, 60x220nm 51.6deg orbit, the insertion point is at only 92nm (171km).

That would seem to go a long way to explaining the difference.

Ares have both been further tweaked to force a lower insertion point and trade that for improved performance, but my tools don't allow me to do that manually to DIRECT, so I think there is a potential for additional DIRECT performance here...

Ross.

HIP2BSQRE · « **Reply #310 on:** 11/07/2006 05:58 am »

Good Evening,

In TeamVisions "Space Exploration Optimization (http://www.teamvisioninc.com/downloads/AIAA-2006-7517-146.pdf) they mention a vicheal called the Jupiter I. Is your vicheal divered off their vicheal? Do you see some ways that the "Direct" vicheal can benifit from some of the analysis that TeamVision have done? One of biggest differance is they postpone development of the JS2-X but would put approx 80mt into LEO (p.14). It is pretty funny, two organizations coming up with the same type of vicheal at approx. the same time.

kraisee · « **Reply #311 on:** 11/07/2006 08:51 am »

I haven't really looked into any of their proposals before.

I saw the photo of the massive Jupiter 3 uber-booster over on another thread and thought it was seriously wild. But honestly, I thought it was just for fun... I didn't know until just now that it was supposed to be serious. I also didn't know they had smaller booster concepts too, especially not one which is indeed quite similar to DIRECT.

Although neither of our groups were the first on this concept

The idea actually dates back to Challenger. It was a logical alternative (with 3 SSME's), along with Shuttle-C, for launching cargo missions, and was even considered for use with a re-started Apollo line of Crew capsules too. But there was simply no cash available to develop any new launcher at the time, so NASA just fixed Shuttle and carried on using that.

In 1991 it resurfaced as the National Launch System (NLS), proposed by NASA and the DoD with 4 less powerful, cheap main engines, as an alternative platform for launching Titan-IV class (and heavier) payloads.

And it also appeared in the ESAS Report with the SSME's again (LV 24/25).

It's not the slightest surprise to me that two separate groups have come up with this particular variant, using two RS-68's though. It's such a simple concept to create from what we fly today, that its really a no-brainer. The RS-68's simply remove the three very expensive SSME's ($90m each!), and replaces them with two, far cheaper, engines. That's what NASA did already with Ares-V anyhow, so it's a completely NASA-validated approach.

The only difference between the Jupiter-I concept and DIRECT is the Regen version of the RS-68 which DIRECT uses to boost performance. This is not an expensive upgarde, certainly only a tiny fraction as expensive as the Ares-I's $4Bn program just to make and qualify the 5-seg SRB!

Jupiter-II is kinda similar to the DIRECT Growth Options which I showed in my proposal, but Jupiter-II has a third stage, so would be more complex. Thus Jupiter-II would also cost more to make & operate than DIRECT+EDS.

I should perhaps, at this point, also mention that there actually is no current requirement for any program to go much above 100mT to LEO per launch - unless of course your CLV is a truly puny little performer like Ares-I is. In that single case, your Cargo lifter is forced to make up the difference if you're trying to keep launch numbers down.

As for Jupiter-III - that behemoth will *never* be commissioned. It's got way too many engines to go wrong, too many elements which must all work together reliably, too many separation events, too many different propellant types, too much mass to use any existing facilities, too much thrust to launch from *ANY* existing site, and a list of development nightmares spottable from orbit! The cost for developing that beast would be truly obscene (100 billion or more, bare minimum), and with so many things to go wrong - it would simply *never* be considered safe.

Of course, if you want to launch 500mT - you could just spend $980m for 7 DIRECT launches, and save yourself a whole bunch o' cash.

If you really must have a new 500mT to LEO booster, it would be a LOT cheaper and easier to just build a really big dumb booster in the Nova class. There are quite a few Nova concepts which would be far simpler to make, some of which offer far higher performance than the Jupiter-III.

But man, what an incredible looking rocket concept!!! It's like Saturn-V, Shuttle and Nova bits all got 'Frankensteined' together!!!

It will never be considered, because of cost and safety, but I sure would love to see it 'brough to life' on some TV show or Film. I can see it now, lifting a daring crew off to rescue the Earth in "Armagedon-II"!

Ross.

Zond · « **Reply #312 on:** 11/07/2006 07:29 pm »

Quote

kraisee - 6/11/2006 6:30 PM
I have severely over-estimated the cost at $1Bn.

Understand that the work required to upgrade RS-68 to "Regen" spec is considerably less than the work to make the J-2X - and that contract has been issued to PWR already totalling just $50 million for complete DDT&E, 7 test units, 2 engines for certification and one full-spec flight engine. And remember that J-2X is going to have an all-new nozzle. This means it is going through it's own "Regen" development program as just one *part* of that contract.

Where do you get your number for a total development cost of $50 million for the J-2X? According to this press release, NASA Announces J-2X Rocket Engine Development Contract Pratt & Whitney Rocketdyne get $50 million just to initiate development.
I expect that J2-X development will cost over $1 billion. And if the rumors are true and Ares-I is seriously underperforming they will try to optimise the J-2X to the limit which will make the development costs soar.

kraisee · « **Reply #313 on:** 11/07/2006 08:50 pm »

Yeah, that's true.

If the J-2X spec stays around 275k/450s I think it will stay below $1Bn. If they have to squeeze every last scrap of performance to compensate for all the other shortcomings in Ares-I, then yes, the cost may well exceed that figure.

But RS-68 doesn't have to do that. The powerplant was designed to be operated roughly 5% higher than it is being operated today on Delta-IV. 5% extra performance is what you get by upgrading to the Regen. It is not a coincidence - it is part of the intended design. The pumps, injectors, combustion chambers have ALL been qualified to those limits already. All PWR expect to need is the right nozzle, and a test/re-qualification program.

I was actually told "under $300m" by my PWR contacts. While they said that was for a man-rated unit, I still assumed doubling that for NASA's "man-rating" requirements, and rounded THAT up to give plenty of room for unexpected growth. $1Bn is a *gross* over-estimate IMHO. If the final figure turns out to be $800m, or even $400m, then NASA can afford to launch an extra probe or two to somewhere interesting. If it does grow to $1Bn, it still stays "on-budget", and everyone's happy because it still costs $3Bn less than developing the 5-seg SRB's...

And it wil lnot cost anywhere near as much as J-2X. That engine is technically a much more complicated unit than RS-68, with a considerably wider range of requirements. It must be air-startable, must be re-startable and will be required to have an extensive on-orbit lifespan too.

No such complicated requirements apply to RS-68 - it is a ground lit, used-once, disposable engine, albeit a fairly big one - and the 'big cost' items are already designed to produce the power we want - so there is no program at all to improve performance beyond current design limits - like J-2X is having to go through. So NONE of those costs will apply.

Ross.

rumble · « **Reply #314 on:** 11/07/2006 10:16 pm »

Part of the RS-68 mods will be to eliminate the huge fireball at ignition (or so I hear). What causes that, and how can it be eliminated? The fireball blackened most of the insulation on the Delta IV Heavy flight! That seems a bit dangerous to me.

And besides...the SSME doesn't do it... ;-)

Jim · « **Reply #315 on:** 11/07/2006 11:37 pm »

Quote

rumble - 7/11/2006 5:59 PM

Part of the RS-68 mods will be to eliminate the huge fireball at ignition (or so I hear). What causes that, and how can it be eliminated? The fireball blackened most of the insulation on the Delta IV Heavy flight! That seems a bit dangerous to me.

And besides...the SSME doesn't do it... ;-)

It is a artifact of the fuel lead to condition the engine for ignition. It isn't dangerous and doesn't really need to be eliminated. The Delta IV design takes it into account.

rumble · « **Reply #316 on:** 11/08/2006 03:43 am »

Oh... I had always assumed that was part of what needed to happen to man-rate that engine.

How does the SSME avoid that? Would the down-draft of the sound suppression water tend to pull the H2 downward?

If it's really not an issue, then I guess these questions are just academic. It's just un-settling for ignition to scorch the paint & insulation 1/2 way or more up the rocket.

simonbp · « **Reply #317 on:** 11/08/2006 03:58 am »

Quote

Jim - 7/11/2006 6:20 PM
It isn't dangerous and doesn't really need to be eliminated. The Delta IV design takes it into account.

A massive, Hydrogen-powered fireball encasing your launch vehicle isn't dangerous?!?

It might be within engineering tolerances, but that doesn't make it inherently safe. I really doubt that it would not be a roadblock to man-rating....

Simon

kraisee · « **Reply #318 on:** 11/08/2006 11:11 am »

The "ball of flame" at RS-68 ignition is NOT a hinderance to man-rating. Recall that the Saturn-V had a similar event at ignition too, but it was a far larger vehicle, so while the ball of flame was a similar size, it *appeared* to be a much smaller event.

The TPS on the stage itself can easily handle this short-term effect. And in fact it will have to handle far worse heating later on the ascent as the plume starts to ride up the sides of the stage.

None-the-less, NASA has requested PWR to mitigate it in the man-rated variant. I understand this is primarily for PR concerns though, not technical ones. It simply doesn't look good on TV to the tax=paying public and the politico's who are unaware of what's really going on.

IIRC, the plan is to re-route the excess propellant back through the engine somehow, but I don't have any specifics to hand. Ultimately though, it's just a plumbing job

Ross.

rumble · « **Reply #319 on:** 11/08/2006 12:59 pm »

Well, could we take ADVANTAGE of the fireball, then? Could we leave either an un-scorchable or an easily blackened bit of lettering on the core stage? Maybe a clear paint with different thermal characteristics... So that after the fireball, it reveals "NASA" or "USA" or something... whatever.

The fireball would then be seen as an expected and designed-for event, reporters would be talking about it & looking forward to it, and it would then be perceived as the non-event it is.

Thermally-reactive paint may be cheaper than a re-plumbing job.