Author Topic: DIRECT v3.0 - Thread 1  (Read 1228211 times)

Offline robertross

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Re: DIRECT v3.0 - Thread 1
« Reply #1620 on: 06/25/2009 08:36 pm »

Even with the mitigation efforts, TO on Ares-I is expected to still be able to impart up to +/- 2.0g of vibrations on the Crew Module, although seat isolators are hoped to reduce that for the crew themselves.
Ross.


At those kind of alternating loads you have to start looking at Metal fatigue.. not just Ultimate and Tensile strength.  I would want a higher FS for material in that envirnoment.

Not to take this any further OT, but that brings to mind the comments made by the ULA rep at the panel hearing on D4H: 2x FS for untested systems, 1.4 FS for tested systems with (around) 1.2 FS as noted.

Ares I should have 2x FS throughout.

I'm sure the commission will weigh these points in their overview: tested vs untested systems (IE: rockets).

Offline kraisee

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Re: DIRECT v3.0 - Thread 1
« Reply #1621 on: 06/25/2009 09:14 pm »
I like your DIRECT approach, but there is one major problem. The manned Mars mission is still in the too-distant future (2032 or something).

Everything MUST be done in order to chop at least 10 years from the current DIRECT proposal. Please, look at this thread.

While I'd love to see it, I don't believe that mission is possible by 2022 because of funding shortages.

The only way to afford it, would be to close ISS, close the Lunar Program and close the NEO options as well, then dedicate everything for the next 15 years to the Mars effort alone.

I think it's a better option to pursue a stepping-stone approach which allows us to afford to go to all the destinations.

It might take a little longer to get to Mars, but I'd prefer a solution which allows Lunar, NEO's, Phobos and Mars -- all -- within 25 years than a program which concentrates on any "one trick pony" solution in 15 years.

Ross.
« Last Edit: 06/25/2009 09:43 pm by kraisee »
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Offline William Barton

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Re: DIRECT v3.0 - Thread 1
« Reply #1622 on: 06/25/2009 09:30 pm »
It's possible that a succession of non-lunar-landing missions, done at about one year intervals, might excite the public interest in (and willingness to pay for) a Mars landing, which would speed things up. Say, a lunar orbital mission (view or nearby Moon), a visit to Webb (view of Earth from much farther than the Moon), a visit to a "nearby" NEO (round trip of a few weeks), then a visit to a more "remote" NEO (round trip of a few months), finally a trip to Phobos. None of that would require the massive investment a Mars landing would require. And the view of Mars from Phobos must be staggering. You'd never manage a Mars landing by 2022, but you might manage Phobos. (Or some  such data, I haven't looked at the potential launch windows.)

Offline Ben the Space Brit

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Re: DIRECT v3.0 - Thread 1
« Reply #1623 on: 06/25/2009 10:33 pm »
Lets just say that they have been asking questions, we are preparing data for them and some of the team have made contact directly.   And the contacts have all been good so far.

That's good news.  They can't say: "Who are you people?" anymore.  They know the answer now because they've met you. :)
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Offline A_M_Swallow

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Re: DIRECT v3.0 - Thread 1
« Reply #1624 on: 06/25/2009 11:27 pm »
You'd never manage a Mars landing by 2022, but you might manage Phobos. (Or some  such data, I haven't looked at the potential launch windows.)

But what if VASIMR works well? In theory it can reduce the one-way trip time to 30 days.

One month to Mars, one month on Mars, and one month to Earth. That's only three months! Doesn't sound too risky anymore.

The 30 day trip assumes the energy is coming from a nuclear reactor and that the heavy concrete and lead shielding used on nuclear sea ships is not needed.  Three to four month trips may be possible using solar powered VASIMRs.
« Last Edit: 06/25/2009 11:44 pm by A_M_Swallow »

Offline A_M_Swallow

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Re: DIRECT v3.0 - Thread 1
« Reply #1625 on: 06/25/2009 11:43 pm »
{snip}I can understand EML1, as you are between the Earth and Moon.  For L2R, you'd need to be on the far side of the moon?  Wouldn't that burn extra fuel to get there than L1R?
Possibly.  The plan is to use the Moon's gravity as a brake, that saves a lot of fuel.

Quote

Why did Apollo take the CSM and LEM to LOI?
{snip}

Because that is what NASA planned when man still had not left LEO.

Over simplifying, Apollo found EML-1 and EML-2 and realised how useful they would be, next time.  This is the next time.

Offline kraisee

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Re: DIRECT v3.0 - Thread 1
« Reply #1626 on: 06/26/2009 02:52 am »
EML-1 and EML-2 are primarily useful for architectures planning to use re-usable landers and ISRU -- neither of which is going to happen cheaply, nor soon.

If you aren't, they actually cost more in terms of delta-V than the EOR-LOR/Loiter plan which is currently baselined.


Another concern, is that if you fly the Orion separately from the Altair, you have no Apollo-13 style "lifeboat" capability at all and that results in a significant hit to your overall mission safety.


Given a fixed starting mass delivered to LEO by two vehicles, the bottom line is that the EOR-LOR/Loiter is THE most efficient means to get to the Lunar Surface and back with the current requirements for Global Access and Any Time Return capability.   It delivers the highest payload mass to the surface for every flight -- which is the real yardstick you need to measure things by.


EML-2 would make for a truly wonderful staging area for any mission heading out into the rest of the solar system though.   If you could assemble all of your Mars (and later Jovian) vehicles there and fuel from the Lunar surface that would make for a stunning capability.

But trying to establish that sort of architecture straight out of the gate on day 1 is like planning to build a complete national highways system in a single week.   You're biting off more than you can realistically chew in one mouthful, and all you're actually going to end up doing is choking on it -- and a Lunar Landing is already a *really* big bite all on its own without ever trying to complicate it in any way.

No, what really you need to do is a very simple business and management technique:

1) Identify where you are and what resources you have right now.

2) Identify where you wish to be and what resources you need to get there.

3) Identify means to break that "giant leap" into a series of smaller, easier, steps.

4) Begin the process of achieving each of those steps, in order, in an orderly manner.

5) When you achieve the last of those steps you will have reached your ultimate target.   Job Done.   What's next?


Ross.
« Last Edit: 06/26/2009 05:50 am by kraisee »
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Offline kraisee

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Re: DIRECT v3.0 - Thread 1
« Reply #1627 on: 06/26/2009 06:16 am »
It's possible that a succession of non-lunar-landing missions, done at about one year intervals, might excite the public interest in (and willingness to pay for) a Mars landing, which would speed things up. Say, a lunar orbital mission (view or nearby Moon), a visit to Webb (view of Earth from much farther than the Moon), a visit to a "nearby" NEO (round trip of a few weeks), then a visit to a more "remote" NEO (round trip of a few months), finally a trip to Phobos. None of that would require the massive investment a Mars landing would require. And the view of Mars from Phobos must be staggering. You'd never manage a Mars landing by 2022, but you might manage Phobos. (Or some  such data, I haven't looked at the potential launch windows.)

William,
Sorry I missed this post earlier.

I completely agree with you.   I believe that NASA ought to look at doing something "interesting" every year, and doing at least one "showcase" mission (preferably 2) every Presidential term.   And I don't limit that to HSF either -- there are a number of "showcase" robotic missions which I believe should be funded too.

What NASA has been missing for many, many years is "vitality".   We have been locked in the mode of "same old thing" for way too long.   Yes, each mission is a challenge.   Yes, each mission is worthwhile.   But the sad fact is that the public switched off a long time ago and only tunes in again once in a while when NASA does something new and different, or when things go horribly wrong.

If we had strong public interest, it would translate to strong political interest as well.   And that would ultimately mean that the budget wouldn't be such a problem.

But the *ONLY* way to accomplish that is to plan a program which can:

a) Afford to do pay for a wide range of new missions, both human and robotic

b) Ensure we have an affordable, yet powerful and extremely flexible and reliable, infrastructure beneath all the endeavours.


What CxP is currently doing isn't meeting any of those objectives.   It is busting the budget on the launch vehicles alone, it is duplicating existing capability when there really is no need, it is creating a huge capability which just isn't going to be an economical one, and it is preventing any money being provided to pay for any interesting missions -- even costing so much as to force the cancellation of many aspects of this program which it was supposed to support in the first place (Pressurized Lunar Rover & Lunar Base for example).


Something has to change.

Ross.
« Last Edit: 06/26/2009 06:17 am by kraisee »
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Offline Steven Pietrobon

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Re: DIRECT v3.0 - Thread 1
« Reply #1628 on: 06/26/2009 07:05 am »
The most critical factor in any 2-launch Lunar mission architecture is maximizing the amount of propellant lofted to LEO for the TLI.   Anything which might reduce that capacity, reduces Lunar performance by a factor of more than 3, so if you lose just 300kg of TLI propellant to LEO, the effect is actually that you lose about 1 full ton of payload performance actually being sent to the moon.

The factor of three is incorrect. I reported this error in my Direct Rebuttal review. Here is what I wrote.

"p.71 Direct claim that for every kg of EDS stage mass increase, this results in 3 kg loss in payload mass through TLI. This is incorrect. You first lose 1 kg of payload mass due to EDS stage mass increase. You then lose 0.93 kg by not having the extra 1 kg of propellant available. Total payload loss is 1.93 kg, 35% less than what Direct claim.

(The rocket equation is given by dv = ve*ln(1+mp/mf) where dv = 1.01*3175 = 3206.8 m/s is TLI delta-V (1.01 factor is flight performance reserve increase), ve = 4393.4 m/s is J-2X exhaust speed, mp is propellant mass and mf is final mass. Rearranging the rocket equation have mf = mp/(exp(dv/ve)-1) = 0.93*mp. Therefore 1 kg loss of propellant results in 0.93 kg loss in payload.)"
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline Steven Pietrobon

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Re: DIRECT v3.0 - Thread 1
« Reply #1629 on: 06/26/2009 07:27 am »
It's theoretically possible.   It's a 7.3mT PLF so that's a significant weight penalty which will increase the propellant needed to do the burn.

My calculations show that the total propellant mass for separate burns by Orion and Altair is 2% greater than for for a single burn by Altair. This is because Altair uses its more efficient LOX/LH2 engine.

Injection Orbit: 56x241 km
Orbital insertion delta-V: dv = 55 m/s
Orion exhaust speed: veo = 3098.9 m/s (316 s ISP)
Altair exhaust speed: vea = 4167.8 m/s (425 s ISP)
Orion mass: mo = 20.2 t
Altair mass: ma = 45.0 t
PLF mass: mplf = 5.7 t (the LAS is 7.3 t see J246-41.4004.10050_CLV_090606.pdf)

Rocket equaltion: dv = ve*ln(1+mp/mf)
dv = delta-V
ve = exhaust speed
mp = propellant mass
mf = final mass

Rearrange: mp = mf*(exp(dv/ve)-1)

For Orion  mp = 20.2*(exp(55/3098.9)-1) = 362 kg
For Altair mp = 45.0*(exp(55/4167.8 )-1) = 598 kg
Total propellant = 363+598 = 960 kg

For Orion/Altair/PLF mp = (20.2+45.0+5.7)*(exp(55/4167.8 )-1) = 942 kg

Thus separate burns uses 18 kg or 2% more propellant than a single burn.
« Last Edit: 06/26/2009 07:34 am by Steven Pietrobon »
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Online MP99

Re: DIRECT v3.0 - Thread 1
« Reply #1630 on: 06/26/2009 07:31 am »
The most critical factor in any 2-launch Lunar mission architecture is maximizing the amount of propellant lofted to LEO for the TLI.   Anything which might reduce that capacity, reduces Lunar performance by a factor of more than 3, so if you lose just 300kg of TLI propellant to LEO, the effect is actually that you lose about 1 full ton of payload performance actually being sent to the moon.

The factor of three is incorrect. I reported this error in my Direct Rebuttal review. Here is what I wrote.

"p.71 Direct claim that for every kg of EDS stage mass increase, this results in 3 kg loss in payload mass through TLI. This is incorrect. You first lose 1 kg of payload mass due to EDS stage mass increase. You then lose 0.93 kg by not having the extra 1 kg of propellant available. Total payload loss is 1.93 kg, 35% less than what Direct claim.


But the burnout mass of the EDS through TLI is also increased by 1 kg, which consumes an additional 1.07 kg of fuel which is unavailable for injecting payload.

1.93 + 1.07 = 3.0 kg reduction.


I got this wrong recently in claiming that a 4mT increase in EDS mass would make DIRECT struggle to match CxP's 71mT TLI requirement, using 2:1 disadvantage.

That should have been 3mT (from 13mT to 16mT burnout), due to the 3:1 disadvantage.

cheers, Martin

Online MP99

Re: DIRECT v3.0 - Thread 1
« Reply #1631 on: 06/26/2009 07:33 am »
For Altair mp = 45.0*(exp(55/4167.8)-1) = 598 kg


I notice you edited this to add a space to remove the smiley.

You can also just tick the "don't use smileys" tickbox.

cheers, Martin

Offline Steven Pietrobon

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Re: DIRECT v3.0 - Thread 1
« Reply #1632 on: 06/26/2009 07:40 am »
But the burnout mass of the EDS through TLI is also increased by 1 kg, which consumes an additional 1.07 kg of fuel which is unavailable for injecting payload.
1.93 + 1.07 = 3.0 kg reduction.

No, the increase in EDS mass has already been accounted for. You are counting this twice. EDS mass increases by 1 kg, TLI payload mass decreases by 1 kg. 1 kg less propellant means 0.93 kg less payload.

Using the rocket equation shows this. Have

mp = propellant mass
ms = EDS burnout mass
mc = TLI mass
dv = delta V = 1.01*3175 = 3206.8 m/s
ve = exhaust speed = 4393.4 m/s for J-2X

Rocket Equation: dv = ve*ln(1+mp/(ms+mc))

Rearranging mc = mp/(exp(dv/ve)-1) - ms = 0.93*mp - ms

Say ms increases by 1 kg to ms2 = ms+1. Then mp decreases by 1 kg to mp2 = mp-1. Then

mc2 = 0.93*mp2 - ms2
      = 0.93*(mp-1) - ms -1
      = 0.93*mp - ms - 0.93  -1
      = mc - 1.93

That is, TLI mass is decreased by 1.93 kg as expected.

You can also just tick the "don't use smileys" tickbox.

OK, thanks.
« Last Edit: 06/26/2009 08:07 am by Steven Pietrobon »
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline usn_skwerl

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Re: DIRECT v3.0 - Thread 1
« Reply #1633 on: 06/26/2009 08:47 am »
I didn't notice if my question was answered already. I didn't find it, so forgive me if it's there. Does the ET (core) need any mods internally (ribs and spars for example) to support the thrust structure and payload?

Thanks
Jeph
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Re: DIRECT v3.0 - Thread 1
« Reply #1634 on: 06/26/2009 09:08 am »
Noted that the June 22 issue of AW&ST's coverage of the Augustine Commission (on page 40) didn't have a single mention of Direct but did mention everything else, including Not Shuttle C.

 ???

« Last Edit: 06/26/2009 09:12 am by Envious »
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Offline Lab Lemming

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Re: DIRECT v3.0 - Thread 1
« Reply #1635 on: 06/26/2009 11:13 am »
Taking some snippets from http://www.spaceref.com/news/viewsr.html?pid=31601 titled "NASA ESMD Internal Memo from Jeff Hanley: 6/20 Cx Update - Moving Forward".

Quote
Much attention has been focused on the probability of loss of crew (pLOC) as a figure of merit in determining the crew launch aspect of the architecture, and we expressed that the ESAS pLOC numbers were all using the same methodology and that the value was in the comparative results and not in the absolute numbers. Very simply, Ares' clear advantage is in the comparative simplicity of its first stage (the shuttle SRM) and use of a single gas generator cycle upper stage engine. These two attributes alone provide substantial robustness over, for example, a more complex liquid pump fed first stage and a multiengine upper stage - simply put, they are more complex with more moving parts. What Ares affords us, in accordance with the findings of the CAIB, is a crew launch system that has the potential to achieve unmatched safety in human spaceflight history. And this is not just a Constellation 'claim' as some would suggest, but has been validated by independent experts in the field of physics based probabilistic risk assessment. There will be much more provided on this topic as well.

How many moving parts does the GEM 46 have?

Offline jeff.findley

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Re: DIRECT v3.0 - Thread 1
« Reply #1636 on: 06/26/2009 12:19 pm »
John & mars,
I really don't feel comfortable discussing that in public without seeking permission from the panel first.

Lets just say that they have been asking questions, we are preparing data for them and some of the team have made contact directly.   And the contacts have all been good so far.

We have decided to leave it entirely to the panel themselves to control the release of all such materials and discussions for themselves according to their own policies.


Thanks for this update.  I personally think that if the panel makes contact with enough members of the team, they'll be more willing to believe that Direct had to move underground to avoid persecution. 

The presentation to the panel about side-mounted SDV's started off with the presenter essentially saying there was no sane reason for Direct to go underground.  It was a direct attack on Direct.  I'm sure that presenter didn't feel threatened because there are so many downsides to a side-mount SDV that upper management wouldn't be threatened by that presentation.

Offline winkhomewinkhome

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Re: DIRECT v3.0 - Thread 1
« Reply #1637 on: 06/26/2009 01:38 pm »
Noted that the June 22 issue of AW&ST's coverage of the Augustine Commission (on page 40) didn't have a single mention of Direct but did mention everything else, including Not Shuttle C.

 ???



Keep an eye out for a DIRECT only story.
Dale R. Winke

Offline adamsmith

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Re: DIRECT v3.0 - Thread 1
« Reply #1638 on: 06/26/2009 02:37 pm »
EML-1 and EML-2 are primarily useful for architectures planning to use re-usable landers and ISRU -- neither of which is going to happen cheaply, nor soon.

If you aren't, they actually cost more in terms of delta-V than the EOR-LOR/Loiter plan which is currently baselined.


Another concern, is that if you fly the Orion separately from the Altair, you have no Apollo-13 style "lifeboat" capability at all and that results in a significant hit to your overall mission safety.


Given a fixed starting mass delivered to LEO by two vehicles, the bottom line is that the EOR-LOR/Loiter is THE most efficient means to get to the Lunar Surface and back with the current requirements for Global Access and Any Time Return capability.   It delivers the highest payload mass to the surface for every flight -- which is the real yardstick you need to measure things by.


EML-2 would make for a truly wonderful staging area for any mission heading out into the rest of the solar system though.   If you could assemble all of your Mars (and later Jovian) vehicles there and fuel from the Lunar surface that would make for a stunning capability.

But trying to establish that sort of architecture straight out of the gate on day 1 is like planning to build a complete national highways system in a single week.   You're biting off more than you can realistically chew in one mouthful, and all you're actually going to end up doing is choking on it -- and a Lunar Landing is already a *really* big bite all on its own without ever trying to complicate it in any way.

No, what really you need to do is a very simple business and management technique:

1) Identify where you are and what resources you have right now.

2) Identify where you wish to be and what resources you need to get there.

3) Identify means to break that "giant leap" into a series of smaller, easier, steps.

4) Begin the process of achieving each of those steps, in order, in an orderly manner.

5) When you achieve the last of those steps you will have reached your ultimate target.   Job Done.   What's next?


Ross.

Ross,

First, let me say you and the gang are doing a terrific job.  And I am a true believer in the KISS principle.  But there is one flaw in your argument,  We've already done LOR. Is EOR-LOR that much harder? Where has  all the knowledge disappeared to?  It's time to explicitly put in place an architecture that will support the true long term goal of going beyond Cislunar space.  I can assure you that architecture using Jupiter launch vehicles would cost less that the current dead end NASA CxP, one spot on the moon, 1.5 launch architecture.  Indeed I would argue that go slow has many advantages and since I personally remember where I was on July 20, 1969, I feel I can argue that if I can be patient so can others.

Respectfully and sincerely yours,

Stan

Offline gregzsidisin

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Re: DIRECT v3.0 - Thread 1
« Reply #1639 on: 06/26/2009 02:57 pm »
Noted that the June 22 issue of AW&ST's coverage of the Augustine Commission (on page 40) didn't have a single mention of Direct but did mention everything else, including Not Shuttle C.

 ???



Just wanted to say that Shannon's / NASA's "Not Shuttle C" is actually similar to a proposal I made in 2003 for a "Shuttle B":

http://www.nsschapters.org/ny/nyc/Shuttle-Derived%20Vehicles%20Modified.pdf

http://www.spacedaily.com/news/oped-03zzs.html

The main similarity is in permanently attaching the engines to the ET. 

I did advocate using RS-68 engines, which would have significantly lowered performance.  It's interesting that now, after the SSME has been rejected as too expensive for so many previous concepts, people are saying that expending SSMEs isn't so bad, especially if you produce them in decent numbers.

I'm tickled that Shannon's Side-Mount has similarities to Shuttle B - I got to tell him so after the public hearing.

However, I'm also concerned that people will look at this and conclude that the reduced capabilities and growth potential will be worth a perceived political expedience (i.e., save jobs quickly).

A side-mount solution may have been best used alongside the current STS, to do heavier lifting alongside the human-carrying vehicle.  Although then the question might have become, why have that big Shuttle Orbiter anyway?
Greg Zsidisin

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"In essence, rocket science is about blowing a lot of hot gas out an orifice. There are more experts in this field than you might realize." -GZ, 2011

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