Author Topic: CEV/EELV Discussion – Warning not for the Politically Correct  (Read 55727 times)

Offline Smatcha

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Since some managers at NASA have threatened to cancel current and future contracts for all who dare question the great and powerful Oz about details concerning process, methods, back-up, documentation, conclusions & decisions maybe we can discuss reasonable EELV alternatives to the SRB CLV here?

Only history will tell if this is an academic exercise or maybe something more important.  At a minimum it would be a good place to vent some frustration in a positive direction if only to outline “what might have been” for the sake of history.

I would strongly recommend everyone read Enchanted Rendezvous if they already haven’t.  It’s a good primer on how experts can be wrong and how good ideas rejected out of hand by the majority can win in the end due to the persistence of a minority.

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19960014824_1996007704.pdf

“The reasonable man seeks to conform to the world, the unreasonable man seeks to conform the world to him, therefore all progress is the result of the unreasonable man”

Kayla your responses, particularly to Doug are well written.  It was clear you know more than he does about your organizations launch products capabilities.  Why these facts didn’t make it to the final report is clear.

I haven’t seen anyone in here yet to present the Boeing Delta side. Hopefully some brave soul well emerge.  I’ll do my best until they do arrive.

Now on to the discussion, to begin with let’s shelve cost for now and focus on five primary questions related to payload performance.


First Question:
What are the off the shelf payload numbers to ISS for various models of EELV?
(see attached table for my first shot at this)

Second Question:
What is the performance reduction required to increase the structural and operational margins of the vehicle?

Third Question:
What is the performance reduction (assuming only an escape tower and no CEV/SM abort thrust capability) to eliminate the danger zones from the baseline EELV ascent ellipse?

Fourth Question:
What would be the required DeltaV of the CEV/SM in order to safely abort to an orbit of 185km at any point along the current danger zone of the ascent ellipse?  For nominal ISS missions the CEV/SM fuel could be used for final orbit insertion saving some of the weight penalty.

Fifth Question:
What would be a good compromise between depressing the EELV ascent ellipse and relying on some emergency correction from the CEV/SM engine?  i.e. either to shallow out the re-entry curve or abort to a temporary orbit.

In the end after interacting questions 1-5 what would be a reasonable (97% confidence) payload performance number for all launch vehicles to ISS.
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Offline vt_hokie

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I actually find the SDLV vs EELV argument to be somewhat amusing, as I think both approaches are wrong.  I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

Offline MATTBLAK

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vt_hokie - 23/5/2006 2:54 PM

I actually find the SDLV vs EELV argument to be somewhat amusing, as I think both approaches are wrong. I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

........Says hokie; flogging the same dead horse from one space blogsite to the next!! hokie; I don't wanna name drop here.... Okay I will. John Young told me in May 1996, when I interviewed him at JSC, before the LockMart X-33 design had been chosen --

(From literal, word-for-word transcription from audio tape):

JEFF: Do you favour any of the current shuttle replacements that are being considered at this time?

>>JOHN YOUNG: Well, its pretty hard to replace the space shuttle. The shuttle, when we finish upgrading the ascent capability, it'll be able to put up almost thirty-five tons into low earth orbit. You've got to develop an equivalent vehicle. A rocket scientist said that you could do a single stage to orbit with one percent of the mass fraction. Right now, the shuttle does about one-and-a-half percent mass fraction to orbit. To put the same weight to orbit as shuttle does, with a single stage.... You just cant do it. You can deliver people to orbit (with single stage), you can deliver payloads to orbit, but you wont be able to deliver both of them there simultaneously and do the kind of work that the shuttle does.<<

I read an interview with the man a couple months later where he said almost word for word the same thing, but added; "We might have to add a cluster of solids to that thing (Venturestar).



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Offline Tap-Sa

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vt_hokie - 23/5/2006  7:54 AM
  I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

scramjet Venture Superstar with 125 tonnes payload? Mmm-kay ... *ringing White House* Mr President, the VSE budget just tripled, will you cover it? Mr President? Hello?


back to topic, about question #4. I have a vague memory of chart showing the black zones are quite early in the flight, the CEV probably still has the LAS attached at that point so it pops just the CM and SM is of no use. Those who know better, please confirm or deny.

I have a subquestion regarding the black zone issue. Are those zones with current EELV trajectories really unsurvivable, or has NASA set some arbitrary g-limit for abort conditions (and if so what is it) ? The CM has to survive at least 10g if LAS is activated. Suborbital Mercuries pulled 11g during nominal mission.

Offline vt_hokie

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Actually, I'd favor the CaLV for heavy lift and a small reusable SSTO or TSTO space plane for crew transport, which would be able to fly more cheaply, more frequently and with less preparation than the CEV/"SRB Stick" crew transport.

Offline Jim

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Tap-Sa - 23/5/2006  3:01 AM

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vt_hokie - 23/5/2006  7:54 AM
  I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

scramjet Venture Superstar with 125 tonnes payload? Mmm-kay ... *ringing White House* Mr President, the VSE budget just tripled, will you cover it? Mr President? Hello?


back to topic, about question #4. I have a vague memory of chart showing the black zones are quite early in the flight, the CEV probably still has the LAS attached at that point so it pops just the CM and SM is of no use. Those who know better, please confirm or deny.

I have a subquestion regarding the black zone issue. Are those zones with current EELV trajectories really unsurvivable, or has NASA set some arbitrary g-limit for abort conditions (and if so what is it) ? The CM has to survive at least 10g if LAS is activated. Suborbital Mercuries pulled 11g during nominal mission.

Don't remember the specific numbers but it was high g's sustained over a matter of minutes.

Offline Jim

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Wrt to D-IV, the only option Boeing was working on was the heavy.  Also for ISS orbits, the CEV would be delivered into a lower altitude (apprx 150 nmi) and it would do the orbit raising.

Offline Jim

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vt_hokie - 22/5/2006  11:54 PM

I actually find the SDLV vs EELV argument to be somewhat amusing, as I think both approaches are wrong.  I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

X-34 wasn't going to orbit.  X-33 was never going to orbit either.

Offline Smatcha

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Jim - 23/5/2006  4:23 AM

Quote
Tap-Sa - 23/5/2006  3:01 AM

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vt_hokie - 23/5/2006  7:54 AM
  I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

scramjet Venture Superstar with 125 tonnes payload? Mmm-kay ... *ringing White House* Mr President, the VSE budget just tripled, will you cover it? Mr President? Hello?


back to topic, about question #4. I have a vague memory of chart showing the black zones are quite early in the flight, the CEV probably still has the LAS attached at that point so it pops just the CM and SM is of no use. Those who know better, please confirm or deny.

I have a subquestion regarding the black zone issue. Are those zones with current EELV trajectories really unsurvivable, or has NASA set some arbitrary g-limit for abort conditions (and if so what is it) ? The CM has to survive at least 10g if LAS is activated. Suborbital Mercuries pulled 11g during nominal mission.

Don't remember the specific numbers but it was high g's sustained over a matter of minutes.

Jim, attached is an industry day (Oct 31, 2005) chart given by Steve Cook on the ascent profile for the SRB/CLV.

Do you know if the orbit of 30x160 is in nmi?  No units are shown kind of important.

Theoretically if we match this we should be close to NASA blessed ascent ellipse.

Also notice that there is a delay of 80 seconds between the ignition of the 2nd stage and LAS jettison.  My guess is they might have an abort re-entry problem after SRB separation as well due to the ballistic disposal requirement of the SRB.

I would also assume that after the LAS is ejected the back-up for the crew abort is the SM thrust.  I have often wondered why we through away the LAS thrust.  It would seem that if we fire one solid at a time (as opposed to all at once for abort scenarios) we could then leverage some of that energy, leaving only the last one to clear the LAS from the vehicle, now even lighter from depleting most of its fuel.

Also if it is nmi than the apogee is 296 km.  I would also assume that circularization of the orbit is done by the SM.  This would be poor use of resources since that would mean utilizing a lower ISP fuel and require us to carry around to the Moon and back a larger tank than absolutely necessary.

Every kg to LEO about $10,000
Every kg to LLO about $20,000
Every kg to Lunar surface $40,000

A kg here a kg there before you know it you are talking real money.

I assume this high apogee as compared to Apollo’s 185km is due to provisions for orbital decay of the lunar cargo and TLI package yet another waste of resources, though reasonable priced at only $10,000 kg.  Fuel boil off or its mitigation (yet more required wt in LEO) will also be a big issue along with crew launch pressure or a billion dollars winds up in the Ocean.

What’s the longest duration we have held LH2 in a tank in space?

As far as the RLV stuff goes could you please start your own thread.
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Offline Jim

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SMetch - 23/5/2006  9:55 AM

[
Do you know if the orbit of 30x160 is in nmi?  No units are shown kind of important.

Theoretically if we match this we should be close to NASA blessed ascent ellipse.

Also notice that there is a delay of 80 seconds between the ignition of the 2nd stage and LAS jettison.  My guess is they might have an abort re-entry problem after SRB separation as well due to the ballistic disposal requirement of the SRB.

I would also assume that after the LAS is ejected the back-up for the crew abort is the SM thrust.  I have often wondered why we through away the LAS thrust.  It would seem that if we fire one solid at a time (as opposed to all at once for abort scenarios) we could then leverage some of that energy, leaving only the last one to clear the LAS from the vehicle, now even lighter from depleting most of its fuel.

Also if it is nmi than the apogee is 296 km.  I would also assume that circularization of the orbit is done by the SM.  This would be poor use of resources since that would mean utilizing a lower ISP fuel and require us to carry around to the Moon and back a larger tank than absolutely necessary.


nmi

It is not the orbit that determines whether there are black zones, it is how the LV flies to the injection point (this was illustrated in another thread).

The LAS is not just used for the first stage, there can be failures in the upperstage that require aborts.  The LASA  is used until the SM can provide sufficent separation in an abort.  One of the factors in determining this is how much propellant is left in the upperstage.  Abort entry problems are not a function of the LAS use.

The use of the LAS for propulsion is not a good idea (also see previous thread). Here are the reasons:
1.  There are not multiple abort motors in the system, just one motor with multiple nozzles.
2.  The CM/SM attach system is not made to take this type of load
3.  Loads on the crew.  Ok for abort but not for an everyday occurance

circularization does not take that much energy (see the shuttle) and this greatly aids in upperstage disposal.  Something similar is done for Mars trajectories.  Spacecraft are not targeted directly at Mars but offset.  This is to prevent a dead spacecraft and/or upperstage from hitting Mars.  The spacecraft must carry the fuel to cancel the offset.  Programs wanted to target the spacecraft directly to Mars and have the upperstage deflect after separation, but since this was another restart and blah, blah, blah, it wasn't allowed.

Offline Smatcha

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Jim - 23/5/2006  10:19 AM

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SMetch - 23/5/2006  9:55 AM

[
Do you know if the orbit of 30x160 is in nmi?  No units are shown kind of important.

Theoretically if we match this we should be close to NASA blessed ascent ellipse.

Also notice that there is a delay of 80 seconds between the ignition of the 2nd stage and LAS jettison.  My guess is they might have an abort re-entry problem after SRB separation as well due to the ballistic disposal requirement of the SRB.

I would also assume that after the LAS is ejected the back-up for the crew abort is the SM thrust.  I have often wondered why we through away the LAS thrust.  It would seem that if we fire one solid at a time (as opposed to all at once for abort scenarios) we could then leverage some of that energy, leaving only the last one to clear the LAS from the vehicle, now even lighter from depleting most of its fuel.

Also if it is nmi than the apogee is 296 km.  I would also assume that circularization of the orbit is done by the SM.  This would be poor use of resources since that would mean utilizing a lower ISP fuel and require us to carry around to the Moon and back a larger tank than absolutely necessary.


nmi

It is not the orbit that determines whether there are black zones, it is how the LV flies to the injection point (this was illustrated in another thread).

The LAS is not just used for the first stage, there can be failures in the upperstage that require aborts.  The LASA  is used until the SM can provide sufficent separation in an abort.  One of the factors in determining this is how much propellant is left in the upperstage.  Abort entry problems are not a function of the LAS use.

The use of the LAS for propulsion is not a good idea (also see previous thread). Here are the reasons:
1.  There are not multiple abort motors in the system, just one motor with multiple nozzles.
2.  The CM/SM attach system is not made to take this type of load
3.  Loads on the crew.  Ok for abort but not for an everyday occurance

circularization does not take that much energy (see the shuttle) and this greatly aids in upperstage disposal.  Something similar is done for Mars trajectories.  Spacecraft are not targeted directly at Mars but offset.  This is to prevent a dead spacecraft and/or upperstage from hitting Mars.  The spacecraft must carry the fuel to cancel the offset.  Programs wanted to target the spacecraft directly to Mars and have the upperstage deflect after separation, but since this was another restart and blah, blah, blah, it wasn't allowed.

Jim could you insert the links you referred to in this thread?
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Offline Jim

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Offline vt_hokie

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Jim - 23/5/2006  7:38 AM
X-34 wasn't going to orbit.  X-33 was never going to orbit either.

I know that, but if the programs hadn't been killed, we'd be beyond the flight test phase by now and could be applying the technology toward an orbital vehicle.

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SMetch - 23/5/2006  9:55 AM
As far as the RLV stuff goes could you please start your own thread.

By all means.

Offline Jim

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vt_hokie - 23/5/2006  6:24 PM

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Jim - 23/5/2006  7:38 AM
X-34 wasn't going to orbit.  X-33 was never going to orbit either.

I know that, but if the programs hadn't been killed, we'd be beyond the flight test phase by now and could be applying the technology toward an orbital vehicle.

Or they could have ended up like the X-wing aircraft (the combo helicopter/airplane), a dead end.

Offline yinzer

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Jim - 23/5/2006  5:10 PM

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vt_hokie - 23/5/2006  6:24 PM

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Jim - 23/5/2006  7:38 AM
X-34 wasn't going to orbit.  X-33 was never going to orbit either.

I know that, but if the programs hadn't been killed, we'd be beyond the flight test phase by now and could be applying the technology toward an orbital vehicle.

Or they could have ended up like the X-wing aircraft (the combo helicopter/airplane), a dead end.

Sure, that's a possibility.  But as it is, we have no knowledge as to the technological viability of reusable rocket boosters.  The only knowledge we gained was regarding the ability of MSFC to manage development programs that represent an incremental advance on the state of the art circa 1960 (i.e. not much), but we seem determined to ignore it.

The DC-X provided surprising and positive results to the questions about operability of reusable rocket-fueled vehicles, not to mention rocket-powered landing using throttleable RL10s (which is of direct relevance to the VSE).  It's too bad that the line of research was shut down to feed NASA's budgetary monster du jour.
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Offline mlorrey

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Jim - 23/5/2006  6:38 AM

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vt_hokie - 22/5/2006  11:54 PM

I actually find the SDLV vs EELV argument to be somewhat amusing, as I think both approaches are wrong.  I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

X-34 wasn't going to orbit.  X-33 was never going to orbit either.

X-33 was never intended for orbit, though the anti-spaceplane crowd always seems to portray it as if that was a goal it was unable to fulfill. The larger Venturestar was intended to reach orbit. X-33 was only intended to reach about Mach 14-18 on suborbital trajectories from Edwards to Utah, to test the technologies that would go into Venturestar.

Nor was X-33 or Venturestar intended as any sort of heavy lift launcher. It was intended to put men in orbit or medium payloads up to 10k lb. The plan was to leave heavy lift for EELV from the start. Of course, EELV was not invented by NASA, and given the "not invented here" syndrome at NASA, that was an unforgivable sin.
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Offline Jim

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Re: CEV/EELV Discussion Warning not for the Politically Correct
« Reply #16 on: 05/24/2006 05:56 pm »
There is no NIH wrt to the EELV's.  NASA has 10 on their manifest and maybe more to come

Offline yinzer

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NASA has no choice for satellite launches, the Commercial Space Act says they can only engage in development activities for requirements that can't be met by capabilities being used by national security and commercial launches.  It'd be hard to argue that EELVs can't launch the MRO or MSL or something like that; this is also probably a large reason why there's no NASA pictures of the Stick with a 3rd stage for GTO missions floating around.

But as soon as Griffin had the chance, he ditched the EELVs for the Stick.
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Offline Kayla

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Re: CEV/EELV Discussion Warning not for the Politically Correct
« Reply #18 on: 05/26/2006 12:45 am »
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Jim - 24/5/2006  12:43 PM

There is no NIH wrt to the EELV's.  NASA has 10 on their manifest and maybe more to come

Jim, I think you are talking about KSC and NASA's science and Earth observation missions. GRC actually was the key to Centaur’s original development and Shuttle Centaur back in the 80’s.

On the other NIH appears to be alive and healthy at MSFC.  Obviously not with everyone. A look at the ongoing consternation about the ability of the incredibly reliable RL10 to be used for human missions is just one view into this world.  I’m impressed that the RS68 is being seriously considered.

Offline Kayla

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SMetch - 22/5/2006  8:36 PM

First Question:
What are the off the shelf payload numbers to ISS for various models of EELV?
(see attached table for my first shot at this)

The following provides Atlas performance to generic orbits.  
http://www.lockheedmartin.com/data/assets/12455.pdf
Typically the ISS performance is about 5% below the due E. LEO performance (I personally do not have the official advertised ISS performance values in hand).

Since this link indicates Atlas V HLV dual engine Centaur (DEC) LEO performance at 29.4 mT, the ISS performance would be on the order of 27.9 mT.  This is also dependant on what altitude one defines as the desired ISS orbit.  But 27.9 mt is a good approximation for cargo delivery on an Atlas V HLV/DEC with a large, 87’ plf.

Quote

Second Question:
What is the performance reduction required to increase the structural and operational margins of the vehicle?
The interesting thing here is that when flying the CEV, one does not have the payload fairing above the Centaur and thus performance actually improves.  During the OSP days, the intent was to carry the OSP load with the payload fairing base module.  The PLF base module is the portion of the Atlas 5.4m PLF that surrounds the Centaur. Since than the Atlas program has further improved the concepts for carrying capsules.

With the low maximum dynamic pressure, ~400 lbf/ft2, relatively low acceleration and minimal bending moment NASA’s structural human rating requirements are actually satisfied without beefing up the rocket.  This is due to the fact that the Atlas V is designed around a family of rockets.  The Atlas 501 for example does not stress the rocket.  The Atlas 551 is the vehicle that with its high maximum dynamic pressure, ~900 lbf/ft2, defined the entire Atlas V family loading and structural design.  ESAS conveniently used the 1.25 structural factor of safety from the Atlas 551.  Applying this to the relatively lightly loaded HLV is truly either miss under standing the rocket or purposely misleading the study.

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Third Question:
What is the performance reduction (assuming only an escape tower and no CEV/SM abort thrust capability) to eliminate the danger zones from the baseline EELV ascent ellipse?

The Atlas V HLV/DEC provides sufficient upper stage thrust such that the nominal ascent profile only has to be slightly further depressed to close the black zones, causing almost no performance drop off.

My memory is that accounting for the LVHM and minor other human rating modifications, the Atlas V HLV/DEC ISS performance is ~28.9 mT.

Quote

Fifth Question:
What would be a good compromise between depressing the EELV ascent ellipse and relying on some emergency correction from the CEV/SM engine?  i.e. either to shallow out the re-entry curve or abort to a temporary orbit.

Atlas’s intent has always been, and continues to be to meet all of NASA’s requirements.  
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In the end after interacting questions 1-5 what would be a reasonable (97% confidence) payload performance number for all launch vehicles to ISS.

Atlas V HLV/DEC ISS performance is ~28.9 mT

The following link provides quite a bit of info regarding Atlas's plans:
http://www.lockheedmartin.com/wms/findPage.do?dsp=fec&ci=17607&rsbci=14917&fti=0&ti=0&sc=400

Offline wannamoonbase

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vt_hokie - 22/5/2006  10:54 PM

I actually find the SDLV vs EELV argument to be somewhat amusing, as I think both approaches are wrong.  I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

I would like the Starship Enterprise too but we don't have 46 Trillion to spend and any of current technologies obviously were not even close to be enough to making it possible.

If there were a silver bullet out there that would cut costs to 10% of the current value or revolutionize performance any one of the players world wide would have brought it to the table sometime ago.

We are stuck with ELV tail of fire technology for at least a few more decades.
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Offline Kayla

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wannamoonbase - 25/5/2006  9:57 PM

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vt_hokie - 22/5/2006  10:54 PM

I actually find the SDLV vs EELV argument to be somewhat amusing, as I think both approaches are wrong.  I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

I would like the Starship Enterprise too but we don't have 46 Trillion to spend and any of current technologies obviously were not even close to be enough to making it possible.

If there were a silver bullet out there that would cut costs to 10% of the current value or revolutionize performance any one of the players world wide would have brought it to the table sometime ago.

We are stuck with ELV tail of fire technology for at least a few more decades.

These comments get to the heart of why I have suggested the use of smaller launchers with cryo transfer instead of the CaLV.  The CaLV is so large that it will only fly occasionally, strictly for these NASA exploration missions.  It will be a monopoly that no one will dare compete against.  The monetary cost of entry is too huge, and the market opportunity is minimal.

On the other hand, if NASA were to open up the 300 mT/year launch requirement for two lunar missions to competitive launch, this has just created a huge market.  Added together the entire current US launch market consists of under 200 mT/year: 4 Shuttles (20 mT), 8 EELV’s (10 mT), few commercials (10 mT).  In one swoop of a pen, NASA could over double this launch market, and I’m not including the crewed launches.

If the exploration architecture is independent of rocket size, this new market opens the door wide for launch competition!!!  It also means that to compete one doesn’t have to develop a Saturn V class rocket to get maybe 2 launches a year.  Small Delta II, or EELV class rockets work just fine.  Who can provide the low cost, reliable on orbit delivery?  

With this impetus, I’d love to see the incredible launch technology advancement 20 years from now.  Going down the current path, one doesn’t need a crystal ball, the launch technology will not have changed.  The high prices will still haunt us.  Real space access will still be a dream.

Offline Smatcha

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Kayla - 25/5/2006  6:17 PM

The Atlas V HLV/DEC provides sufficient upper stage thrust such that the nominal ascent profile only has to be slightly further depressed to close the black zones, causing almost no performance drop off.

My memory is that accounting for the LVHM and minor other human rating modifications, the Atlas V HLV/DEC ISS performance is ~28.9 mT.


Thanks for the info

One of the ways we could save a lot of money in servicing the ISS would be to ship up a lighter CEV/SM (ISS Config) vehicle in the first place.  I’m pretty sure the Russians are not sending people up to ISS with anything heavy as the current SRB/CLV does.

One of the big problems with the SRB/CLV is the first stage only has one general power setting.  You have to add mass to hold it down.  One size never fits all.  As such they must send up a lunar configured CEV/SM to the ISS every time.

What are the ISS performance (man launch) numbers for all the entire Atlas Line?  I’m trying to match an ISS CEV/SM with another lunar architecture we are working on.

If we can get the ISS mission done with a smaller EELV utilizing the same CEV with an ISS and Lunar SM variant we could save a lot of money and time.

For the Boeing Delta Line, one core booster might be sufficient.  With less engines to go wrong this should help in the reliablilty department.

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Offline Smatcha

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Kayla - 26/5/2006  4:15 AM

Quote
wannamoonbase - 25/5/2006  9:57 PM

Quote
vt_hokie - 22/5/2006  10:54 PM

I actually find the SDLV vs EELV argument to be somewhat amusing, as I think both approaches are wrong.  I'd rather see us revive programs like X-33 and X-34 and start building some new hardware based on new designs.

I would like the Starship Enterprise too but we don't have 46 Trillion to spend and any of current technologies obviously were not even close to be enough to making it possible.

If there were a silver bullet out there that would cut costs to 10% of the current value or revolutionize performance any one of the players world wide would have brought it to the table sometime ago.

We are stuck with ELV tail of fire technology for at least a few more decades.

These comments get to the heart of why I have suggested the use of smaller launchers with cryo transfer instead of the CaLV.  The CaLV is so large that it will only fly occasionally, strictly for these NASA exploration missions.  It will be a monopoly that no one will dare compete against.  The monetary cost of entry is too huge, and the market opportunity is minimal.

On the other hand, if NASA were to open up the 300 mT/year launch requirement for two lunar missions to competitive launch, this has just created a huge market.  Added together the entire current US launch market consists of under 200 mT/year: 4 Shuttles (20 mT), 8 EELV’s (10 mT), few commercials (10 mT).  In one swoop of a pen, NASA could over double this launch market, and I’m not including the crewed launches.

If the exploration architecture is independent of rocket size, this new market opens the door wide for launch competition!!!  It also means that to compete one doesn’t have to develop a Saturn V class rocket to get maybe 2 launches a year.  Small Delta II, or EELV class rockets work just fine.  Who can provide the low cost, reliable on orbit delivery?  

With this impetus, I’d love to see the incredible launch technology advancement 20 years from now.  Going down the current path, one doesn’t need a crystal ball, the launch technology will not have changed.  The high prices will still haunt us.  Real space access will still be a dream.

While I agree with your overview I have a few questions.

First) Isn’t there a diminishing cost savings as volume increases?  While 2-4x the volume will help significantly in lowering cost 10x the volume won’t help as much.

Second) Doesn’t orbital assembly of spacecraft in automated/manual fashion add weight and cost?

Third) Zero boil off technology, orbit fuel transfer, plus preventing LH2 from leaking through the tank walls are non-trivial.

Fourth) Orbital decay is another method of burning money.

Fifth) Launch Window complexity interacting with a string of events that must happen.  Ie building a space station every time we want to go to Mars or put something heavier on the lunar surface than a Apollo Lunar Lander.

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Offline Jim

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SMetch - 26/5/2006  12:50 PM

Quote
Kayla - 25/5/2006  6:17 PM

The Atlas V HLV/DEC provides sufficient upper stage thrust such that the nominal ascent profile only has to be slightly further depressed to close the black zones, causing almost no performance drop off.

My memory is that accounting for the LVHM and minor other human rating modifications, the Atlas V HLV/DEC ISS performance is ~28.9 mT.


Thanks for the info

One of the ways we could save a lot of money in servicing the ISS would be to ship up a lighter CEV/SM (ISS Config) vehicle in the first place.  I’m pretty sure the Russians are not sending people up to ISS with anything heavy as the current SRB/CLV does.

One of the big problems with the SRB/CLV is the first stage only has one general power setting.  You have to add mass to hold it down.  One size never fits all.  As such they must send up a lunar configured CEV/SM to the ISS every time.

What are the ISS performance (man launch) numbers for all the entire Atlas Line?  I’m trying to match an ISS CEV/SM with another lunar architecture we are working on.

If we can get the ISS mission done with a smaller EELV utilizing the same CEV with an ISS and Lunar SM variant we could save a lot of money and time.

For the Boeing Delta Line, one core booster might be sufficient.  With less engines to go wrong this should help in the reliablilty department.


Why ISS orbit?  Injection would be at the same inclination, but not at the same altitude.  
 if you add SRM's with TVC for the Delta IV reliabity goes down, you are better off with a heavy

As for the SM, just offload fuel.  Having two different SM's would be more expensive

Offline Smatcha

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Jim - 26/5/2006  10:06 AM

Quote
SMetch - 26/5/2006  12:50 PM

Quote
Kayla - 25/5/2006  6:17 PM

The Atlas V HLV/DEC provides sufficient upper stage thrust such that the nominal ascent profile only has to be slightly further depressed to close the black zones, causing almost no performance drop off.

My memory is that accounting for the LVHM and minor other human rating modifications, the Atlas V HLV/DEC ISS performance is ~28.9 mT.


Thanks for the info

One of the ways we could save a lot of money in servicing the ISS would be to ship up a lighter CEV/SM (ISS Config) vehicle in the first place.  I’m pretty sure the Russians are not sending people up to ISS with anything heavy as the current SRB/CLV does.

One of the big problems with the SRB/CLV is the first stage only has one general power setting.  You have to add mass to hold it down.  One size never fits all.  As such they must send up a lunar configured CEV/SM to the ISS every time.

What are the ISS performance (man launch) numbers for all the entire Atlas Line?  I’m trying to match an ISS CEV/SM with another lunar architecture we are working on.

If we can get the ISS mission done with a smaller EELV utilizing the same CEV with an ISS and Lunar SM variant we could save a lot of money and time.

For the Boeing Delta Line, one core booster might be sufficient.  With less engines to go wrong this should help in the reliablilty department.


Why ISS orbit?  Injection would be at the same inclination, but not at the same altitude.  
 if you add SRM's with TVC for the Delta IV reliabity goes down, you are better off with a heavy

As for the SM, just offload fuel.  Having two different SM's would be more expensive

Not if the SM Thrust Package and Life Support Extension Package were separate stacked units

Since all mission’s will naturally have difference in duration’s vs manrated DV this would be a good place to split the two and arrive at any number of missions.  

"TVC"?
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Offline Jim

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thrust vector control

Offline Norm Hartnett

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So in amongst the massive amount of OT stuff in this thread we got some possibly good data on the Atlas packages. Anyone want to present the Delta packages?

Everything I read indicates that Griffin is not in love with the Stick and this thread http://forum.nasaspaceflight.com/forums/thread-view.asp?tid=2710&posts=5&start=1 indicates he isn't wed to ESAS's version of manrated either.
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Offline Jim

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Norm Hartnett - 29/5/2006  12:55 AM

So in amongst the massive amount of OT stuff in this thread we got some possibly good data on the Atlas packages. Anyone want to present the Delta packages?

There aren't any equivalent Delta IV charts.  Their launch infrastructure prevents them from widening the CBC.  Any upgrades are to the Heavy and involve either strapon SRM's, upgraded RS-68 and upperstage engines, or larger Upperstage tanks.   There is a chart elsewhere in the forum showing this.

Offline gladiator1332

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http://www.lockheedmartin.com/wms/findPage.do?dsp=fec&ci=17607&rsbci=14917&fti=0&ti=0&sc=400

That link has some interesting information on the Atlas V evolution plan. There is a table in there that shows the performance of each of the vehicles.

There are some vehicles on the table that are not capable of luanching the current CEV. The Phase 1A LV is simple, however there needs to be more upperstage engines to allow it to launch the CEV. So for now I would consider it unworkable.

The Phase 1B LV is capable of launching the CEV, however it needs 3 solids to do so.

The Phase 1C LV seems to be overkill to launch the CEV...nearly 40 tons, the CEV weighs about half of that.

The Phase 2A LV is the preferred choice I believe for the CLV. It does not require solids and it seems that Lockheed can easily upgrade to the 5.4 meter booster. It provides more than enough lift without being overkill.

The last two options are overkill for the CLV and seems to be a little light for the CaLV.

Offline mlorrey

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That phase 2a launcher looks a lot like my proposed alternative to the CLV: using an ET derived structure, SSMEs or RS-68s, single stage to orbit expendable. The 2e looks like my proposed clustering of these (though Lockheed doesn't go as far as my five and seven tank clusters.)
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For awhile I liked the ET derived structure for the CLV as well. I guess the main reason for its decline in my opinion is that we are now using a wider core for the CaLV. The Et derived CLV made more sense when it was using the same core as the CaLV.
It still is an interesting idea, and the fact that there are different cores does not kill it.  
I don't know if I can fully support an EELV derived design, as I am a fan of the SDLVs. If the stick is ever to go, I would like to see it replaced by something like the LV you describe above.

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Had some time to kill so I modeled your concept in Gmax. Instead of using the 8.4m ET derived core, I used the 10m CaLV core. The concept to the left is your single stage idea, and to the right is my old two stage idea. THe second stage is derived from the CLV, however it is only 20m in height, rather than 32m.

Offline mlorrey

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Here's my proposed LV constellation, which would put into LEO a regular supply of ETs for salvage or use in space station modules. The ETs would launch with framing, grille decking, hatches, and conduits installed in the tanks, with a fuel scavenge system in the intertank and deployable solar panels on the sides. These ET launchers would drop all but one main engine at 65-70% of fuel burned, to boost payload in orbit. The engines may or may not be recoverable and/or reusable.
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Offline Jim

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The wet design was rejected for Skylab but of qualifing materials for the cyrogenic environment.   I am assuming that the fuel is H2 otherwise purging that tank would be hard.   Is the outside of the vehicle also mmod proof?  What would they be salvage for?

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Jim - 5/6/2006  6:03 PM

The wet design was rejected for Skylab but of qualifing materials for the cyrogenic environment.   I am assuming that the fuel is H2 otherwise purging that tank would be hard.   Is the outside of the vehicle also mmod proof?  What would they be salvage for?

The basic concept is essentially a modified shuttle ET, with 5-6 expendable SSMEs, or 4 RS-68s. LH2 would be ventable or scavengable. MMOD would be simply keeping the foam, which will be encased in a plastic sheath by orbital personnel to prevent possible "popcorning", by either a wrap or sleeve method.

If the ET were redesigned to handle LOX and RP-1, it would be structurally a bit heavier, but could carry twice as much payload into orbit for the same tankage volume. RP-1 remnants should evaporate to space upon venting.

Here's a close up of the CEV layout. It uses a more Gemini-like capsule design (I always liked that one better) relying on SHARP TPS materials for the heat shield and a higher, more gradual reentry trajectory. Note the air lock with MMUs in the nose
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Re: CEV/EELV Discussion Warning not for the Politically Correct
« Reply #36 on: 06/06/2006 04:41 am »
In the old ET orbital use studies, it was found that the current foam was not  good MMOD protection and also deteroirated quickly.

Offline mlorrey

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Re: CEV/EELV Discussion Warning not for the Politically Correct
« Reply #37 on: 06/06/2006 04:57 am »
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Jim - 5/6/2006  11:28 PM

In the old ET orbital use studies, it was found that the current foam was not  good MMOD protection and also deteroirated quickly.

The foam, uncontained, apparently does popcorn off, particularly in areas which have not glazed over with a rind from atmospheric heating at max Q, which is one reason I've proposed sleeving it in. Do you have any references to studies of the foam for its micrometeorite resistance versus materials currently used on ISS?
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There are some critical points that are being overlooked here.  The exploration program is entended to be sustainable.  Unique NASA-only LV's will require a unique infrastructure and workforce (a la shuttle) and thus be hugely expensive.  All NASA's money will go into launching to LEO (like it is today).

Lockheed's solution to the sustainablility dilema is a modular system whose low end replicated the cost and performance of today's Atlas V 401.  The 401 is the simplest (and theoretically) cheapest rocket on the face of the planet, delivering 10mT to LEO with a 2 stage, 2 engine vehicle: 1 RD-180, 1 RL-10.  This is what Elon Musk is claiming for falcon 1 but at 10 times the performance and with the distinct advantage that it works.

Maintaining the low end price point ensures that the system can be used economically for DoD missions, NASA science missions and commercial missions.  Its inherent low cost and high reliability makes the 401 ideal for the fledgling commercial tourism market, once people get tired of the 4 minute sub-leo experience.

Lockheed's phase 2 family (5.4m Lox/LH2 upper stage, 5.4m Lox/RP booster stage, RD-180 and RL-10 engines) provides vehicles that reach all the way back to the 401 class as well as extend up to over 70mT with the 3-body version.  Modularity provides rate for low cost and flexibility to cover the range of requirements.  The CLV member of the family is a 2 RD-180 booster coupled with a 4 RL-10 upper stage providing engine out capability on a single stick vehicle.  

All the phase 2 vehicles can be launched from the existing Atlas v pad.  Since the engines are existing and flying today and the other systems can be brought over from Atlas V, the phase 2 development effort consists of new 5.4 m tanks, not really much of a stretch.  Given the strong heritage, there would also be high confidence that it would work first time.

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Re: CEV/EELV Discussion Warning not for the Politically Correct
« Reply #39 on: 06/06/2006 05:18 am »
It deteriorates from exposure to the sun (that's why there is different shades of foam, the older foam is darker).  The ISS uses the bumper method, a shield standing off a few inches from the pressure vessel.  The SOFI has virtually no impact resistance at any level.  Long duration upperstage studies have to take this into account and add MMOD protection.

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mlorrey - 5/6/2006  10:55 PM

If the ET were redesigned to handle LOX and RP-1, it would be structurally a bit heavier, but could carry twice as much payload into orbit for the same tankage volume. RP-1 remnants should evaporate to space upon venting.


The bulk of the hydrocarbons may "evaporate" out, but there will always be a thin film left on the tank walls. Ever try to pump down a vacuum chamber after it was contaminated with oil? You can pump against it until you turn blue (Cryo Pump,Diffusion Pump,Turbo Molecular Pump,chose your weapon) and never get much beyond the milli torr range because of the hydrocarbons left behind. The RP-1 will have to be scrubbed out by hand.  I can lay may hands on a small vacuum chamber that had an oil handling accident (#@$! oil roughing pump) , It has been scrubbed, and I can still smell the oil in it.

The RP-1 tank will be 1/16th the size of the O2 tank. Why mess with the RP-1 tank at all.
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Offline simonbp

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kevin-rf - 6/6/2006  11:18 AM

Quote
mlorrey - 5/6/2006  10:55 PM

If the ET were redesigned to handle LOX and RP-1, it would be structurally a bit heavier, but could carry twice as much payload into orbit for the same tankage volume. RP-1 remnants should evaporate to space upon venting.


The bulk of the hydrocarbons may "evaporate" out, but there will always be a thin film left on the tank walls. Ever try to pump down a vacuum chamber after it was contaminated with oil? You can pump against it until you turn blue (Cryo Pump,Diffusion Pump,Turbo Molecular Pump,chose your weapon) and never get much beyond the milli torr range because of the hydrocarbons left behind. The RP-1 will have to be scrubbed out by hand.  I can lay may hands on a small vacuum chamber that had an oil handling accident (#@$! oil roughing pump) , It has been scrubbed, and I can still smell the oil in it.

The RP-1 tank will be 1/16th the size of the O2 tank. Why mess with the RP-1 tank at all.

Also, any foam dust that "popcorns" off will possibly redeposit itself on other parts of the station, like the solar wings, for example. Over time, this could be a major problem...

Simon ;)

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kevin-rf - 6/6/2006  11:18 AM

Quote
mlorrey - 5/6/2006  10:55 PM

If the ET were redesigned to handle LOX and RP-1, it would be structurally a bit heavier, but could carry twice as much payload into orbit for the same tankage volume. RP-1 remnants should evaporate to space upon venting.


The bulk of the hydrocarbons may "evaporate" out, but there will always be a thin film left on the tank walls. Ever try to pump down a vacuum chamber after it was contaminated with oil? You can pump against it until you turn blue (Cryo Pump,Diffusion Pump,Turbo Molecular Pump,chose your weapon) and never get much beyond the milli torr range because of the hydrocarbons left behind. The RP-1 will have to be scrubbed out by hand.  I can lay may hands on a small vacuum chamber that had an oil handling accident (#@$! oil roughing pump) , It has been scrubbed, and I can still smell the oil in it.

The RP-1 tank will be 1/16th the size of the O2 tank. Why mess with the RP-1 tank at all.

Well, I know one way to get rid of the RP-1: after venting, seal and fill with 100 millibars of pure O2. Light a match. The resulting combustion will not overpressure the tank.
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simonbp - 6/6/2006  11:35 AM

Quote
kevin-rf - 6/6/2006  11:18 AM

Quote
mlorrey - 5/6/2006  10:55 PM

If the ET were redesigned to handle LOX and RP-1, it would be structurally a bit heavier, but could carry twice as much payload into orbit for the same tankage volume. RP-1 remnants should evaporate to space upon venting.


The bulk of the hydrocarbons may "evaporate" out, but there will always be a thin film left on the tank walls. Ever try to pump down a vacuum chamber after it was contaminated with oil? You can pump against it until you turn blue (Cryo Pump,Diffusion Pump,Turbo Molecular Pump,chose your weapon) and never get much beyond the milli torr range because of the hydrocarbons left behind. The RP-1 will have to be scrubbed out by hand.  I can lay may hands on a small vacuum chamber that had an oil handling accident (#@$! oil roughing pump) , It has been scrubbed, and I can still smell the oil in it.

The RP-1 tank will be 1/16th the size of the O2 tank. Why mess with the RP-1 tank at all.

Also, any foam dust that "popcorns" off will possibly redeposit itself on other parts of the station, like the solar wings, for example. Over time, this could be a major problem...

Simon ;)

Any ET that uses RP-1 is going to need a lot less foam, ergo less popcorning, if any. However, the popcorn affect that is alleged is why I've proposed wrapping or sleeving the tank in orbit by orbital personnel.
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Offline kevin-rf

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mlorrey - 6/6/2006  12:22 PM

Quote
kevin-rf - 6/6/2006  11:18 AM

Quote
mlorrey - 5/6/2006  10:55 PM

If the ET were redesigned to handle LOX and RP-1, it would be structurally a bit heavier, but could carry twice as much payload into orbit for the same tankage volume. RP-1 remnants should evaporate to space upon venting.


The bulk of the hydrocarbons may "evaporate" out, but there will always be a thin film left on the tank walls. Ever try to pump down a vacuum chamber after it was contaminated with oil? You can pump against it until you turn blue (Cryo Pump,Diffusion Pump,Turbo Molecular Pump,chose your weapon) and never get much beyond the milli torr range because of the hydrocarbons left behind. The RP-1 will have to be scrubbed out by hand.  I can lay may hands on a small vacuum chamber that had an oil handling accident (#@$! oil roughing pump) , It has been scrubbed, and I can still smell the oil in it.

The RP-1 tank will be 1/16th the size of the O2 tank. Why mess with the RP-1 tank at all.

Well, I know one way to get rid of the RP-1: after venting, seal and fill with 100 millibars of pure O2. Light a match. The resulting combustion will not overpressure the tank.

Raised EyeBrow... So you have replaced uncombusted hydrocarbons with partially combusted hydrocarbons and soot that will be sticking on the walls. The heavier hydrocarbons that stayed behind (more tar like than RP-1 like) will not burn as easily as bulk RP-1. The lighter components will have boiled off.

The other problm is the tanks are an Al alloy. Most Al alloys are quite porous, making it dificult to clean out left over RP-1. We are not talking about Stainless...  

Brings up another question. How easy is it to burn the AlLi used in the ET? Anyone have the ingnition temp of AlLi handy?
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Offline Jim

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mlorrey - 6/6/2006  1:27 PM
QUOTE]

Any ET that uses RP-1 is going to need a lot less foam, ergo less popcorning, if any. However, the popcorn affect that is alleged is why I've proposed wrapping or sleeving the tank in orbit by orbital personnel.

RP-1 and LO2 tanks don't need insulation or foam

Offline J Britt RSA

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mlorrey - 5/6/2006  5:35 PM

Here's my proposed LV constellation, which would put into LEO a regular supply of ETs for salvage or use in space station modules. The ETs would launch with framing, grille decking, hatches, and conduits installed in the tanks, with a fuel scavenge system in the intertank and deployable solar panels on the sides. These ET launchers would drop all but one main engine at 65-70% of fuel burned, to boost payload in orbit. The engines may or may not be recoverable and/or reusable.

mlorrey - do you have a bigger version of the ext_tank_ssto_small.jpg ? As a hobby rocketeer, i'm interested in your design.

Offline mlorrey

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J Britt RSA - 6/6/2006  7:31 PM

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mlorrey - 5/6/2006  5:35 PM

Here's my proposed LV constellation, which would put into LEO a regular supply of ETs for salvage or use in space station modules. The ETs would launch with framing, grille decking, hatches, and conduits installed in the tanks, with a fuel scavenge system in the intertank and deployable solar panels on the sides. These ET launchers would drop all but one main engine at 65-70% of fuel burned, to boost payload in orbit. The engines may or may not be recoverable and/or reusable.

mlorrey - do you have a bigger version of the ext_tank_ssto_small.jpg ? As a hobby rocketeer, i'm interested in your design.

Yeah, but it's too big to upload. PM me your email and I'll send it to you.
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Offline publiusr

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Kayla - 26/5/2006  6:15 AM



These comments get to the heart of why I have suggested the use of smaller launchers with cryo transfer instead of the CaLV.  The CaLV is so large that it will only fly occasionally,

Same with Delta IV--which has no engine out. CaLV is a must for it reduces assembly and is a step into the future of real space. EELV assembly is daft--and ISS's delays are proof of that. This 20 ton at a time stuff is thre wrong way to go.

Offline bad_astra

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Raytheon and others had perfectly acceptable plans that wouldn't have required too many launches to build a lunar base. For whatever reasons ESAS chose to ignore all of them.
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Offline mlorrey

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bad_astra - 10/6/2006  12:32 AM

Raytheon and others had perfectly acceptable plans that wouldn't have required too many launches to build a lunar base. For whatever reasons ESAS chose to ignore all of them.

Mustn't talk about "reasons", you'll be accused of promoting "conspiracy theories".
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Offline gladiator1332

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publiusr - 9/6/2006  5:47 PM

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Kayla - 26/5/2006  6:15 AM



These comments get to the heart of why I have suggested the use of smaller launchers with cryo transfer instead of the CaLV.  The CaLV is so large that it will only fly occasionally,

Same with Delta IV--which has no engine out. CaLV is a must for it reduces assembly and is a step into the future of real space. EELV assembly is daft--and ISS's delays are proof of that. This 20 ton at a time stuff is thre wrong way to go.

Totally agree with you. We have barely flown the Delta IV so we do not know if there are problems hidden somewhere. We'll have a delay and a half finished spacecraft floating around up there wasting away.


Offline bad_astra

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FGB was given a waiver with the understanding it would be corrected later (insert eye roll). I don't see why the same thing could not be done in this case.
"Contact Light" -Buzz Aldrin

Offline Kayla

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gladiator1332 - 10/6/2006  8:08 PM

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publiusr - 9/6/2006  5:47 PM

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Kayla - 26/5/2006  6:15 AM



These comments get to the heart of why I have suggested the use of smaller launchers with cryo transfer instead of the CaLV.  The CaLV is so large that it will only fly occasionally,

Same with Delta IV--which has no engine out. CaLV is a must for it reduces assembly and is a step into the future of real space. EELV assembly is daft--and ISS's delays are proof of that. This 20 ton at a time stuff is thre wrong way to go.

Totally agree with you. We have barely flown the Delta IV so we do not know if there are problems hidden somewhere. We'll have a delay and a half finished spacecraft floating around up there wasting away.


You are worried about the fact that the Delta IV has barely flown????  With 5 successful flights under Delta's belt, and 8 for Atlas V the EELV program is well on its way of demonstrating its reliability and capability with 2 independent rockets.  Yes, Delta has had some major delays, but that is typical of new rocket programs.  The DoD is paying to get through this rough period and Delta will be fully operational in the next couple of years, long before the CLV even demo flight in 2009.  Atlas V, having the advantage of being more derived than Delta IV, has been very successful in meeting its launch day promises, averaging about 1 day of delay for the past 6 launches.

With the SDLV CLV on the other hand, NASA will have to pay for the full development cost and then all of the early flight glitches as well and then NASA gets the benefit of paying for the full infrastructure cost of this NASA only rocket system.  CLV is currently the long pole for replacing the Shuttle, with an ILC of 2012 and first crew flight in 2014.  Any additional delays will further delay America's human space program.  A couple of years delay is likely based on past programs.  Are we willing to have a likely 6 year period with no American crew access to space?  Do we continue to launch shuttles during this period, draining $4 to $5B from exploration.  These issues are likely to jeopardize sending astronauts to the moon prior to 2020!!!

Offline Kayla

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Norm Hartnett - 28/5/2006  11:55 PM

So in amongst the massive amount of OT stuff in this thread we got some possibly good data on the Atlas packages. Anyone want to present the Delta packages?

Everything I read indicates that Griffin is not in love with the Stick and this thread http://forum.nasaspaceflight.com/forums/thread-view.asp?tid=2710&posts=5&start=1 indicates he isn't wed to ESAS's version of manrated either.


Griffin:
"Many, if not most, unmanned payloads are of very high value, both for the importance of their mission, as well as in simple economic terms. The relevant question may be posed quite simplistically: What, precisely, are the precautions that we would take to safeguard a human crew that we would deliberately omit when launching, say, a billion-dollar Mars Exploration Rover (MER) mission? The answer is, of course, “none”. While we appropriately value human life very highly, the investment we make in most unmanned missions is quite sufficient to capture our full attention."

As Griffin says, crew or expensive hardware, NASA will want to the most reliable launch vehicle to fly the expensive payloads.  The CaLV (1:124 failure rate) has a substantially lower reliability than the CLV (1:460 failure rate) according to ESAS: http://www.nasa.gov/pdf/140637main_ESAS_06.pdf, pages 382 & 384.  Although, I by and large disagree with most of ESAS's findings, the concept that a larger, more complex rocket has a lower reliability than a smaller less complex vehicle makes sense.  Large rockets will also have very low flight rates, further reducing the demonstrated reliability.

This is one of the many reasons that I support utilizing smaller rockets than the CaLV to under take exploration.

Folks on this site that are suggesting that smaller launches will require unworkable complex orbital assembly are confusing the ISS experience with what can be for exploration.  I've been suggesting launching the EDS & LSAM on the CLV (EELV derived is my preference), empty of propellant.  Then launch propellant to the EDS on any and all rockets available.  There are daily launch opportunities for the propellant launches to dock with the orbiting EDS.  Passive boil-off measures can reduce boil-off rates to ~0.01%/day, sufficient to support the period that the EDS is in orbit waiting to be completely filled and for the crew to arrive, see: http://www.lockheedmartin.com/data/assets/12382.pdf. Cryo transfer of propellants is not a wild, futuristic technology.  Using a settled environment, cryo transfer is very similar to what cryo stages do today: http://www.lockheedmartin.com/data/assets/12384.pdf.  The Centaur Test Bed concept is an inexpensive way to demonstrate full up cryo transfer in the next few years: http://www.lockheedmartin.com/data/assets/12534.pdf.

This approach uses the most reliable rockets to launch people and the expensive exploration elements, encourages private investment to advance the state of launch vehicles by creating a large market for launched propellant, and reduces NASA's up front investment allowing exploration sooner.  The entire launch architecture is extensible to more advanced missions such as Mars while also reducing launch costs for all other NASA, DoD and commercial endeavors.

Offline meiza

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Kayla, your links don't work for me?

Anyway, technically, docking isn't needed if you use a propellant depot that has a robot arm that can grab and berth the incoming tankers or the to-be-tanked craft.

Offline hyper_snyper

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Just delete the last period from the links.

Offline mlorrey

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Kayla - 11/6/2006  9:25 AM

Griffin:
"Many, if not most, unmanned payloads are of very high value, both for the importance of their mission, as well as in simple economic terms. The relevant question may be posed quite simplistically: What, precisely, are the precautions that we would take to safeguard a human crew that we would deliberately omit when launching, say, a billion-dollar Mars Exploration Rover (MER) mission? The answer is, of course, “none”. While we appropriately value human life very highly, the investment we make in most unmanned missions is quite sufficient to capture our full attention."

This is a good point. I'm not sure what current life insurance actuaries say a human life is worth, monetarily (and of course, the investment in training someone to be an astronaut increases that singificantly, but I recall at one time it was an average of $6.5 million per person), but with payloads themselves worth hundreds of millions, if not billions, of dollars, the idea that a rocket needs to be sufficiently more reliable to be "man rated" is ludicrous from a utilitarian standpoint: you lose as much economically by losing an expensive payload.

Financially, losing 7 astronauts at once is a loss of less than a hundred million dollars worth of payload.

Granted, this is very callously utilitarian, and many may feel that a human life is "priceless", the fact is that courts of law find the worth of a human life in jury awards for wrongful death every day.
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Offline Kayla

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mlorrey - 11/6/2006  4:59 PM

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Kayla - 11/6/2006  9:25 AM

Griffin:
"Many, if not most, unmanned payloads are of very high value, both for the importance of their mission, as well as in simple economic terms. The relevant question may be posed quite simplistically: What, precisely, are the precautions that we would take to safeguard a human crew that we would deliberately omit when launching, say, a billion-dollar Mars Exploration Rover (MER) mission? The answer is, of course, “none”. While we appropriately value human life very highly, the investment we make in most unmanned missions is quite sufficient to capture our full attention."

This is a good point. I'm not sure what current life insurance actuaries say a human life is worth, monetarily (and of course, the investment in training someone to be an astronaut increases that singificantly, but I recall at one time it was an average of $6.5 million per person), but with payloads themselves worth hundreds of millions, if not billions, of dollars, the idea that a rocket needs to be sufficiently more reliable to be "man rated" is ludicrous from a utilitarian standpoint: you lose as much economically by losing an expensive payload.

Financially, losing 7 astronauts at once is a loss of less than a hundred million dollars worth of payload.

Granted, this is very callously utilitarian, and many may feel that a human life is "priceless", the fact is that courts of law find the worth of a human life in jury awards for wrongful death every day.

As you say, callous, but than this thread is "not for the politically correct"!

The reality though is that if either the crewed mission (CLV) or launch of the other elements EDS & LSAM (CaLV) has a failure crewed exploration missions will likely be on hold for years.  With the current architecture, even a CLV launch delay resulting in too much boil-off from the EDS will result in mission failure!  You think launching is expensive, down periods hurt more.  During the already 2.5 years of shuttle recovery NASA has spent ~$12B on the shuttle, for a single launch.  Now that is expensive.

On orbit refueling could be used with the current exploration architecture to ensure that launch delays are not failures, just replenish the EDS and LSAM.  Preferably though, eliminate the CaLV and just fill the EDS and LSAM on orbit.  A propellant launch failure does not mean mission failure if one has backup launchers for the propellant.

Offline Kayla

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SMetch - 26/5/2006  11:50 AM

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Kayla - 25/5/2006  6:17 PM

The Atlas V HLV/DEC provides sufficient upper stage thrust such that the nominal ascent profile only has to be slightly further depressed to close the black zones, causing almost no performance drop off.

My memory is that accounting for the LVHM and minor other human rating modifications, the Atlas V HLV/DEC ISS performance is ~28.9 mT.


Thanks for the info

One of the ways we could save a lot of money in servicing the ISS would be to ship up a lighter CEV/SM (ISS Config) vehicle in the first place.  I’m pretty sure the Russians are not sending people up to ISS with anything heavy as the current SRB/CLV does.

One of the big problems with the SRB/CLV is the first stage only has one general power setting.  You have to add mass to hold it down.  One size never fits all.  As such they must send up a lunar configured CEV/SM to the ISS every time.

What are the ISS performance (man launch) numbers for all the entire Atlas Line?  I’m trying to match an ISS CEV/SM with another lunar architecture we are working on.

If we can get the ISS mission done with a smaller EELV utilizing the same CEV with an ISS and Lunar SM variant we could save a lot of money and time.

For the Boeing Delta Line, one core booster might be sufficient.  With less engines to go wrong this should help in the reliablilty department.


NASA dropped the high performance cryo LO2/LCH4 propulsion module for the CEV because of the need to get the CEV flying quickly.  I completely agree that if we want to use the CEV for ISS access as soon as possible following 2010 a storable propulsion module is what is needed.  However, assuming that this storable propulsion module will also work for lunar missions is significantly degrading or increasing the launch requirements for the lunar missions.

If CEV were to have a minimal storable propulsion module to support ISS missions the CEV weight would be a fraction of the 25 mT full up CEV.  This is especially true if the CEV were delivered to an ISS proximity orbit.  With the CEV (no propulsion module) weight under 10 mT: http://www.nasa.gov/pdf/140636main_ESAS_05.pdf page 223, how light could a very simple propulsion module be.  Challenge the CEV team a little and it might be able to fly on the existing Atlas V single stick vehicle:

LEO performance for the Atlas V dual engine Centaur is 11.3 mT with an estimated reliability of over 0.996.
Or for a little more performance, the Atlas Phase 1 (4 RL10’s) performance is 12.4 mT (or 13.3 mT with 6 RL10’s) with greater reliability than today’s Atlas V.

Designing this minimal ISS only propulsion module now, would free up time for NASA to develop a high performance cryo (preferably LH2 over LCH4) propulsion module for the lunar missions.

Offline Kayla

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mlorrey - 5/6/2006  5:35 PM

Here's my proposed LV constellation, which would put into LEO a regular supply of ETs for salvage or use in space station modules. The ETs would launch with framing, grille decking, hatches, and conduits installed in the tanks, with a fuel scavenge system in the intertank and deployable solar panels on the sides. These ET launchers would drop all but one main engine at 65-70% of fuel burned, to boost payload in orbit. The engines may or may not be recoverable and/or reusable.


Although I completely agree that a robust exploration program can benefit from large elements, I personally do not see the benefit of trying to use the launch vehicle tank as the on-orbit station module.  Think about the number of astronaut hours that will be required to outfit the tank into a station.  Think about the cost of approving all of the vehicle changes to enable the tank to be used on orbit.

The rockets for exploration, be it the CaLV or the various Atlas variants from Phase 2 on, are intended to accommodate an 8.4m PLF that is larger than an ET!  Replace the PLF with the large habitat module and you can have your giant station, already primarily outfitted prior to launch.  And this is without any of the cryo compatibility issues.  Yes this means a dedicated launch, but in the end, I believe that this will be far cheaper than trying to use the CaLV’s tank.

Offline mlorrey

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I do like the larger size of the Phase 2 tanks, as I've said previously, it looks like they are going in the direction I've advocated previously. However, using a 1.5 stage architecture puts lots of cubic in orbit, and cubic is what is needed: cubic is salable/rentable real estate. A real estate market brings real estate customers.

I don't even think the astronauts bringing up tanks need to deal with fitting them out. After the first few are set up and equipped, they are permanently staffed with folks whose job is to pick up other people's tanks and remodel them into livable cubic. The CEVs and whatever other payloads come up with the tanks can go on their merry way to whatever destination they are needed at.

The remodders only have to wire up the tanks. Whatever customers want to fill the space they want to rent they can bring up themselves and move in.

I'm not stuck on the STS ET. It is an example, of what is far better than wasting money on solid rockets, or tanks, that are going to fall back into the ocean. Every kg that makes it to orbit is precious. Every kg that gets 98% of the way to orbit is also precious. It is a huge waste to not take it the rest of the way.
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Offline Smatcha

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Kayla - 11/6/2006  4:47 PM
LEO performance for the Atlas V dual engine Centaur is 11.3 mT with an estimated reliability of over 0.996.
Or for a little more performance, the Atlas Phase 1 (4 RL10’s) performance is 12.4 mT (or 13.3 mT with 6 RL10’s) with greater reliability than today’s Atlas V.

Designing this minimal ISS only propulsion module now, would free up time for NASA to develop a high performance cryo (preferably LH2 over LCH4) propulsion module for the lunar missions.

Agree.  Which came first the SRB CLV or the 1.5 launch approach?  It seems the whole design approach has been backwards from the beginning.

Step 1) We are going to use the SRB for manned launches end of discussion.
Step 2) What do we need to put on top of the SRB so it won’t exceed the g-limits for astronauts?  Answer between 22-25MT of payload.
Step 3) That’s more mass than we need for ISS mission.
Step 4) Quick let’s utilize a 1.5 approach to justify the SRB/CLV for manned launches otherwise known as too much for ISS missions and too little lunar missions.

The cycle is complete.

A more rational approach

Step 1) What do we "actually" need to perform the ISS mission?
Answer a Med Class EELV cheaper (Rec and Non Rec) and safer (one vs three engines) than the Heavy variants.

With two other important benefits;
No gap in manned access to the station.
More money sooner to spend on returning to the Moon method TBD.

“Know do we want to go to the Moon or don’t we?”

Or do we just want to ship up yet more mass to a destination that we don't want to go to anymore in the first place.
“Do we want to go to the moon or not?”
John C. Houbolt - November 15, 1961
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Offline Propforce

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Kayla - 11/6/2006  6:55 AM

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gladiator1332 - 10/6/2006  8:08 PM

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publiusr - 9/6/2006  5:47 PM

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Kayla - 26/5/2006  6:15 AM



These comments get to the heart of why I have suggested the use of smaller launchers with cryo transfer instead of the CaLV.  The CaLV is so large that it will only fly occasionally,

Same with Delta IV--which has no engine out. CaLV is a must for it reduces assembly and is a step into the future of real space. EELV assembly is daft--and ISS's delays are proof of that. This 20 ton at a time stuff is thre wrong way to go.

Totally agree with you. We have barely flown the Delta IV so we do not know if there are problems hidden somewhere. We'll have a delay and a half finished spacecraft floating around up there wasting away.


You are worried about the fact that the Delta IV has barely flown????  With 5 successful flights under Delta's belt, and 8 for Atlas V the EELV program is well on its way of demonstrating its reliability and capability with 2 independent rockets.  Yes, Delta has had some major delays, but that is typical of new rocket programs.  The DoD is paying to get through this rough period and Delta will be fully operational in the next couple of years, long before the CLV even demo flight in 2009.  Atlas V, having the advantage of being more derived than Delta IV, has been very successful in meeting its launch day promises, averaging about 1 day of delay for the past 6 launches.

With the SDLV CLV on the other hand, NASA will have to pay for the full development cost and then all of the early flight glitches as well and then NASA gets the benefit of paying for the full infrastructure cost of this NASA only rocket system.  CLV is currently the long pole for replacing the Shuttle, with an ILC of 2012 and first crew flight in 2014.  Any additional delays will further delay America's human space program.  A couple of years delay is likely based on past programs.  Are we willing to have a likely 6 year period with no American crew access to space?  Do we continue to launch shuttles during this period, draining $4 to $5B from exploration.  These issues are likely to jeopardize sending astronauts to the moon prior to 2020!!!


Very good point. I wonder why do people have more faith in an UN-DEVELOPED, UN-PROVEN, launch system than the existing proven versions.   It reminds me of back in the SLI days, where NASA reliability folks have MORE FAITH with "paper engines" than existing engines such as the SSME, because these "paper engines" have far more 9's of reliability "prediction" than the actually flown SSME.

When it comes to the CLV, at risk of keep beating a dead horse, there's absolutely no technical, nor economic, reasons why the CLV should be built at all.  NASA is already short on funds, why drain an additional $6B from the money you don't have to develop a vehicle that goes nowhere?  Spend $1B to "man-rate" (whatever that means) the EELV fleet.  Let both Lockheed and Boeing, or any other aspiring U.S. launch services companies,  compete for "human-launch" contract from NASA.  Competition works, and it brings out the BEST that any firms can offer, it works in the commercial sector, it worked on the X-prize, and it works in the military aerospace sector, why can't it work with the NASA civil sector?  

The CLV can never replace the Shuttle, not in its payload capacity nor in its mission capability.  So why bother to have an artificially imposed deadline of "must" having an alternate people-taxi come online by the time the Shuttle retires?  We can do the people-taxi thing with the Soyuz or either of the EELV fleet.  Trying to design & develop a brand new launch vehicle in a hurry is a sure receipe for disaster.  Instead, NASA should really focus on the CaLV development, as there's no vaiable alternative to its payload capability and there're plenty of technology risks (don't let the simple powerpoint charts fool you guys) in this new vehicle development.  It will be the biggest venture NASA takes since the Apollo era.  

But this is NASA, where it doesn't have to be efficient, rational, or even economical, where the concept of NPV & IRR is as foreign as the underground water on Mars, but it does have to provide JOBS to many direct civil service employees & contractors with various CONGRESSIONAL DISTRICTS.  So being either technically sound or financially viable is not a priority with regarding to the CEV/ CLV/ CaLV decision making process, but ensuring a continual support of various NASA centers & its people is a priority.  Nevertheless, one could not help but to hope that NASA will do the right things.  A launch architecture as complex as Space Exploration surely would take more than a 6-month ESAS team study, then to have many of its key decision reverted in a few short months.  If such decisions can not withstand the daylight scrutiny, how would it hold up under the sun after billions of dollars of investment made?


Offline zinfab

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A launch architecture as complex as Space Exploration surely would take more than a 6-month ESAS team study, then to have many of its key decision reverted in a few short months.
I accept that people disagree with NASA choosing for political reasons over cost-effective ones, but this particular point makes no sense.

You ping them for not being adaptive and agile, then they make changes/transitions and you call it "reverting". Almost every change they've made was a contigency/option component of the ESAS report or adaptive thinking. By attacking them for changing their minds for good and valid reasons, you only make it HARDER for them to do it again in the future.

What I think you mean to argue is that they'll make a "good" change for SDLV (SSME to RS-68), but will ignore the contractor changes for the "stick." I think everyone here has made valid points regarding safety and reliability. I simply think this particular example diminishes your argument, instead of strengthening it. NASA has actually been pretty good (for NASA) at being willing to consider options (if within a SDLV framework).

Finally, I'd be interested to know how many jobs NASA "saved" by maintaining the shuttle launch parts/manufactor. Would it be "cheaper" for NASA to lay off that many people? Would NASA get congressional funding at that point? I'm not thrilled about that rationale, either. In the big scheme of things, however, LM and/or Boeing would hire some lesser number of those (or other) workers, but not all. Everyone wants a cheaper NASA--just don't close MY NASA facility!

Offline Kayla

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What most people ignore here is that none of us are proposing reducing NASA’s budget.  NASA’s money in the end pays for jobs.  If we spend a lot on SDLV those jobs will be the current jobs tied up in shuttle with very little spent on the exploration end.  If we choose a cheaper alternative launch architecture, NASA will be able to afford more in-space and lunar/Mars work.  Once again this means jobs, not the shuttle jobs but new jobs.  This change scares people.  

What needs to be remembered is that NASA’s money will wind up enabling jobs regardless of the architecture.  With this in mind, we really should be seeking an architecture that creates results, actual astronauts on the lunar surface.  Crewed missions to Mars.  All while not sacrificing the unmanned science.  A NASA that is showing America’s exploration preeminence is much more likely to get Congresses support than a NASA that can barely support ISS.

Offline mlorrey

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Kayla - 15/6/2006  6:18 PM

What most people ignore here is that none of us are proposing reducing NASA’s budget.  NASA’s money in the end pays for jobs.  If we spend a lot on SDLV those jobs will be the current jobs tied up in shuttle with very little spent on the exploration end.  If we choose a cheaper alternative launch architecture, NASA will be able to afford more in-space and lunar/Mars work.  Once again this means jobs, not the shuttle jobs but new jobs.  This change scares people.  

What needs to be remembered is that NASA’s money will wind up enabling jobs regardless of the architecture.  With this in mind, we really should be seeking an architecture that creates results, actual astronauts on the lunar surface.  Crewed missions to Mars.  All while not sacrificing the unmanned science.  A NASA that is showing America’s exploration preeminence is much more likely to get Congresses support than a NASA that can barely support ISS.

Well, I for one wouldn't mind NASA having a smaller budget. If thats what it takes to get them out of the launch business, so be it. That being said, NASA is small fry in the greater behemoth of the federal budget, and generally pays for itself in technological innovation. In the end, I would be happy if they kept the same budget, and got a lot more Buck Rogers for the bucks they have through safer and more cost effective launch systems, and more focus on science and exploration.

The claim now is that ESAS will get us there, but I'll note that they said the same thing about Shuttle.
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Offline publiusr

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Kayla - 15/6/2006  6:18 PM

What most people ignore here is that none of us are proposing reducing NASA’s budget.  NASA’s money in the end pays for jobs.  If we spend a lot on SDLV those jobs will be the current jobs tied up in shuttle with very little spent on the exploration end.  If we choose a cheaper alternative launch architecture, NASA will be able to afford more in-space and lunar/Mars

Rockets came first--with little payloads--and will again. I think more of ESAS than Gump's t/Space...

Enough with the Ayn Rand worship.

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Kayla - 15/6/2006  7:18 PM

What most people ignore here is that none of us are proposing reducing NASA’s budget. NASA’s money in the end pays for jobs. If we spend a lot on SDLV those jobs will be the current jobs tied up in shuttle with very little spent on the exploration end. If we choose a cheaper alternative launch architecture, NASA will be able to afford more in-space and lunar/Mars work. Once again this means jobs, not the shuttle jobs but new jobs. This change scares people.

What needs to be remembered is that NASA’s money will wind up enabling jobs regardless of the architecture. With this in mind, we really should be seeking an architecture that creates results, actual astronauts on the lunar surface. Crewed missions to Mars. All while not sacrificing the unmanned science. A NASA that is showing America’s exploration preeminence is much more likely to get Congresses support than a NASA that can barely support ISS.
The only problem is that what you propose still lays off some people, even while hiring others. The bumps in the road for infrastructure (delayed/cancelled/deferred science programs) churns a lot of people through the employment pump. Districts don't like that. They like stable sources of income.

Government run bodies cost more than business. It's a trade-off. It always has been. At 1% of the US budget, I'm subjectively unimpressed with NASA's budget. The congress/white house/national sentiment seem to be following CAIB's advice-- for now.

We're crunching science to "rush" the CEV by 2 years. No matter how they white-wash it, that's what is happening. If NASA's "stakeholders" could be pleased with the original VSE schedule, science would not have suffered quite as early or as heavily.

I'd rather pay for both. I'm in the minority.

Offline Smatcha

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Kayla - 11/6/2006  4:47 PM

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SMetch - 26/5/2006  11:50 AM

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Kayla - 25/5/2006  6:17 PM

The Atlas V HLV/DEC provides sufficient upper stage thrust such that the nominal ascent profile only has to be slightly further depressed to close the black zones, causing almost no performance drop off.

My memory is that accounting for the LVHM and minor other human rating modifications, the Atlas V HLV/DEC ISS performance is ~28.9 mT.


Thanks for the info

One of the ways we could save a lot of money in servicing the ISS would be to ship up a lighter CEV/SM (ISS Config) vehicle in the first place.  I’m pretty sure the Russians are not sending people up to ISS with anything heavy as the current SRB/CLV does.

One of the big problems with the SRB/CLV is the first stage only has one general power setting.  You have to add mass to hold it down.  One size never fits all.  As such they must send up a lunar configured CEV/SM to the ISS every time.

What are the ISS performance (man launch) numbers for all the entire Atlas Line?  I’m trying to match an ISS CEV/SM with another lunar architecture we are working on.

If we can get the ISS mission done with a smaller EELV utilizing the same CEV with an ISS and Lunar SM variant we could save a lot of money and time.

For the Boeing Delta Line, one core booster might be sufficient.  With less engines to go wrong this should help in the reliablilty department.


NASA dropped the high performance cryo LO2/LCH4 propulsion module for the CEV because of the need to get the CEV flying quickly.  I completely agree that if we want to use the CEV for ISS access as soon as possible following 2010 a storable propulsion module is what is needed.  However, assuming that this storable propulsion module will also work for lunar missions is significantly degrading or increasing the launch requirements for the lunar missions.

If CEV were to have a minimal storable propulsion module to support ISS missions the CEV weight would be a fraction of the 25 mT full up CEV.  This is especially true if the CEV were delivered to an ISS proximity orbit.  With the CEV (no propulsion module) weight under 10 mT: http://www.nasa.gov/pdf/140636main_ESAS_05.pdf page 223, how light could a very simple propulsion module be.  Challenge the CEV team a little and it might be able to fly on the existing Atlas V single stick vehicle:

LEO performance for the Atlas V dual engine Centaur is 11.3 mT with an estimated reliability of over 0.996.
Or for a little more performance, the Atlas Phase 1 (4 RL10’s) performance is 12.4 mT (or 13.3 mT with 6 RL10’s) with greater reliability than today’s Atlas V.

Designing this minimal ISS only propulsion module now, would free up time for NASA to develop a high performance cryo (preferably LH2 over LCH4) propulsion module for the lunar missions.

Kayla, how do you handle aborts after the Launch Abort System has been jettisoned?  Under the current plan the SM Primary Engine performs that role.  The only other way would be to carry the LAS higher up into the orbit.  It seems to me if we had separate LAS motors we could fire at intervals on the way up we could add some deltaV, lighten the load and still have enough for lower deltaV abort requirements at the higher altitudes.

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Offline mlorrey

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If we are in a rush to replace the shuttle, then we should go with the option that requires as little new equipment as possible. On that vein, I (a major critic of Thiokols influence on the space program) propose the SRB-X, topped with an X-38 derived CEV.

WRT upper stages: why do we even need to develop new upper stages? We've got Centaur, among others.
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Offline Jim

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SMetch - 30/6/2006  1:28 PM
Kayla, how do you handle aborts after the Launch Abort System has been jettisoned?  Under the current plan the SM Primary Engine performs that role.  The only other way would be to carry the LAS higher up into the orbit.  It seems to me if we had separate LAS motors we could fire at intervals on the way up we could add some deltaV, lighten the load and still have enough for lower deltaV abort requirements at the higher altitudes.

The SM does it, just like the current CLV.  LAS is only for abort and simple as possible.  Adding all the extra circuitry, extra ignitors and modes defeats it's purpose.  Just one motor for escape.  The thrust for abort has to be enough for approx 15 g's.
Lighting the LAS motors, subjects the CM to exhaust products, and imposes loads on the CM/SM interface, which don't exist in a standard abort.  Breaking up the motor into smaller ones won't be too many

Jettisioning early probably gives more performance than can be supplied by the motor

There is a thread on this elsewhere

Offline Smatcha

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Jim - 30/6/2006  10:57 AM

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SMetch - 30/6/2006  1:28 PM
Kayla, how do you handle aborts after the Launch Abort System has been jettisoned?  Under the current plan the SM Primary Engine performs that role.  The only other way would be to carry the LAS higher up into the orbit.  It seems to me if we had separate LAS motors we could fire at intervals on the way up we could add some deltaV, lighten the load and still have enough for lower deltaV abort requirements at the higher altitudes.

The SM does it, just like the current CLV.  LAS is only for abort and simple as possible.  Adding all the extra circuitry, extra ignitors and modes defeats it's purpose.  Just one motor for escape.  The thrust for abort has to be enough for approx 15 g's.
Lighting the LAS motors, subjects the CM to exhaust products, and imposes loads on the CM/SM interface, which don't exist in a standard abort.  Breaking up the motor into smaller ones won't be too many

Jettisioning early probably gives more performance than can be supplied by the motor

There is a thread on this elsewhere

Jim, the purpose of this idea is to be able to lighten up and save money on the crewed ISS mission by not bringing along full service module just the RCS system.  Without the Lunar-SM we can get the Crew only CM to work on a single stick EELV.  The problem is that if we need to have a late abort after the upper stage is well into its burn we have to able to get away from the upper stage.  Traditionally this is accomplished via the SM primary engines after LAS jettison.  Though I have often wondered about this because it seems that an out of control nearly depleted upper stage would be able to out run even the SM primary engine since it thrust to weight ratio is designed around in space travel.  Anyway

If the capsule nose can handle 15g by firing all SRM’s at once it can handle 2 out 7 at a time.  Also it doesn’t get much simpler than putting separate igniters to each SRM.  I believe Apollo had separate motors in firing sequence as well in order to use offset thrust to move the Apollo capsule up and over from the Saturn V.  Again trying to get away from an out of control upper stage.  You can run but you can’t hide. It sure looks like it goes sideways a little in test videos.
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Offline Jim

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The late aborts not driven by "out of control" upperstages, but by reduced capabilitie like power failures in the upperstage, early depletion, engine shutdowns etc.    That's why the LES/LES is/was jettison, the abort modes are less severe.

The SM will need more than just RCS anyways.  Some Rendezvous adjustments will need higher thrust i.e. main engine.

 14 SRMs for an abort system?  Control of them and timing?  You just have complicated things greatly and for an emergency system

 The LES for Apollo had ONE abort motor with 4 nozzles and a very small pitch motor.

Offline Smatcha

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Jim - 1/7/2006  7:26 AM

The late aborts not driven by "out of control" upperstages, but by reduced capabilitie like power failures in the upperstage, early depletion, engine shutdowns etc.    That's why the LES/LES is/was jettison, the abort modes are less severe.

The SM will need more than just RCS anyways.  Some Rendezvous adjustments will need higher thrust i.e. main engine.

 14 SRMs for an abort system?  Control of them and timing?  You just have complicated things greatly and for an emergency system

 The LES for Apollo had ONE abort motor with 4 nozzles and a very small pitch motor.

Explain to me how an upperstage, which needs to be throttled back so as not over accelerating a combinded vehicle beyond 3g’s, does not over take a vehicle at its nose with a thrust to weight less than ½ g.

This has always been a problem.  Even during Apollo the crew would have been killed after the LAS jettison if the 2nd stage could not be shut down.  Given an explosive event in the 2nd stage they would be dead again with the SM engine not capable of getting outside the range of explosion in time.  Either way this has always been a dangerous condition which our solution will solve and at the same time not require carrying the complete weight of the LAS to orbit.  It’s amazing how this danger zone is somehow ignored when discussing steps that must be taken to “man-rate” a vehicle.  In fact this means that for half of the manrated ascent the crew will die, go figure.

The RCS will work fine with a direct injection profile.  You only need strong upper stages with lofted trajectories something we can’t use on manned mission anyway.  The Shuttle OMS are hardly strong engines when compared with the weight of the Shuttle and they work fine for orbit insertion and manvauering all the way to the ISS.

Its seven not fourteen SRM’s and building demolitions require the precise timing of 1,000 of explosives.  Besides the firing of the individual LAS motors on ascent doesn’t have to be that precise +/- 10 sec and the abort will be require firing all of the remaining ones at once just like it is now.  The purpose of seven is so a even thrust pattern can be maintained, if we can just redirect it through central nozzle like Apollo than the number of separate LAS motors could be reduced to 2-4.

Either way in our plan the crew will not die in the event of a 2nd stage failure where as now they do.  It’s amazing to me how everyone takes complex low margin cryogenic engines so lightly.  I hope the SSME’s never remind us of this fact.  It always seems were overly focused on solving the last problem.

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Offline Jim

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3 g's is not a requirement for the CLV (it is between 4-5 g's) and the J-2 does not throttle.  Part of the abort sequence IS shutting down all stages, which the range safety system does also.  Therefore, a stage can not be shutdown is not a credible failure.  There never had been a requirement to out run a stage (except for the current CLV SRM), never.  The requirement is to outrun a fireball/shockwave, which only happens in the lower atmosphere on lower stages.   At altitudes, the thinner atmosphere doesn't transmit the shock wave.  That is why the LAS can be jettisoned after a certain point.  There was never an issue with outrunning the shockwave after LAS jsettison.

There is only ONE motor for the LAS.  Simplicity is the rule.  No multiple motors and circuitry.  Building demolitions use ordnance lines not electric circuits.  and people's lives don't depend on them.  

Crew does not die with a J-2 failure.

OMS was .05 T/W and really didn't do the abort to orbit (the SSME still had to provide the bulk of the impulse) The RCS for the CEV is .005 to .02 much less.   Apollo's SME was 20K, close to 1

your plan will increase the per flight costs with an extra segment, extra SRB components and extra J-2.  Might as well do an heavy EELV.

Offline Smatcha

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Jim - 3/7/2006  12:12 PM

3 g's is not a requirement for the CLV (it is between 4-5 g's) and the J-2 does not throttle.  Part of the abort sequence IS shutting down all stages, which the range safety system does also.  Therefore, a stage can not be shutdown is not a credible failure.  There never had been a requirement to out run a stage (except for the current CLV SRM), never.  The requirement is to outrun a fireball/shockwave, which only happens in the lower atmosphere on lower stages.   At altitudes, the thinner atmosphere doesn't transmit the shock wave.  That is why the LAS can be jettisoned after a certain point.  There was never an issue with outrunning the shockwave after LAS jsettison.

There is only ONE motor for the LAS.  Simplicity is the rule.  No multiple motors and circuitry.  Building demolitions use ordnance lines not electric circuits.  and people's lives don't depend on them.  

Crew does not die with a J-2 failure.

OMS was .05 T/W and really didn't do the abort to orbit (the SSME still had to provide the bulk of the impulse) The RCS for the CEV is .005 to .02 much less.   Apollo's SME was 20K, close to 1

your plan will increase the per flight costs with an extra segment, extra SRB components and extra J-2.  Might as well do an heavy EELV.

So a 2nd stage uncontrolled failure, ie exploding J-2, is not possible?  While the shock wave at 177,000 ft may be less it will still have some power.  Also the exploding fragments will not be slowed down by the air resistance at higher altitudes.  In summary, we are attempting to cover a contingence that is currently considered a crew write-off event, i.e. exploding 2nd stage engines after LAS jettison.

After all why would one do a controlled 2nd stage shutdown?

Actual the OMS burn is what keeps the Shuttle from landing in the same location as the External tank.  Both have the same general orbit at separation, the only difference is the OMS burn at apogee to move the perigee from 60(External Tank) to 185km (Space Shuttle).  I don’t know what the orbit profile is for ISS. I would assume some 400km x60 because they still need to properly dispose of the external tank.





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Offline Jim

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OMS burns are just little burps, not the same abort to orbit.  It is the same thing they want to do with the CLV.  

There is no need to protect against an exploding J-2.  Those events happen at ignition, where the LAS is still there.  
You do a controlled shutdown if you lose TVC, or other associate systems.  Or if chamber pressure, temp, turbine speed or other parameters are trending to the bad.

The LAS is jettison at the appropriate time when analysis shows there is not threat from the shock wave and the CEV can survive.  Why else did they do it this way for Apollo.  It was to save weight, it was because it was no longer needed.

the LAS is mostly for first stage.  Once safe recovery altitude is reached

Offline Smatcha

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Jim - 3/7/2006  7:57 PM

OMS burns are just little burps, not the same abort to orbit.  It is the same thing they want to do with the CLV.  

There is no need to protect against an exploding J-2.  Those events happen at ignition, where the LAS is still there.  
You do a controlled shutdown if you lose TVC, or other associate systems.  Or if chamber pressure, temp, turbine speed or other parameters are trending to the bad.

The LAS is jettison at the appropriate time when analysis shows there is not threat from the shock wave and the CEV can survive.  Why else did they do it this way for Apollo.  It was to save weight, it was because it was no longer needed.

the LAS is mostly for first stage.  Once safe recovery altitude is reached

So given that an out of control upper stage could out run any in space designed spacecraft thrust unit the minimum design of an attached spacecraft thrust unit would for final orbit insertion and de-orbit maneuvers correct?

Stated another way after second stage ignition and LAS jettison, if the upper stage fails catastrophically or cannot be shut down the crew would be in trouble.

While our approach would save the crew in that contingency it has a price in terms of weight to orbit regardless of whether we partially fire the LAS motors on ascent or not.  If everyone is fine with the above failure scenario no need to penalize our approach.

Besides going to the Moon and coming back is much more dangerous overall than an upper stage failure.  A fact that should be considered in put important components of VSE on hold while we duplicate what we already have in EELV’s for the magical benifits of unproven new SRB base CLV.  While I'm sure the SRB will be fine its all the other new stuff that concerns me.
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SMetch - 6/7/2006  12:21 PM

So given that an out of control upper stage could out run any in space designed spacecraft thrust unit the minimum design of an attached spacecraft thrust unit would for final orbit insertion and de-orbit maneuvers correct?

Stated another way after second stage ignition and LAS jettison, if the upper stage fails catastrophically or cannot be shut down the crew would be in trouble.

While our approach would save the crew in that contingency it has a price in terms of weight to orbit regardless of whether we partially fire the LAS motors on ascent or not.  If everyone is fine with the above failure scenario no need to penalize our approach.


Not a credible failure.  Do you actually think Apollo ignored this?  Same with Gemini.  Shutting down engines is easy.  Ask Elon Musk. All engines use valves that need power  to stay open.  Remove power (multiple ways of doing this) and the engine shuts down.


Offline josh_simonson

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To put it another way, once the second stage is firing the rocket is more or less in space.  There is no SRB to pose an explosion hazard and the fuel is seperate so it can't explode either, there is no air outside so the fuel cannot burn if vented unless oxidizer is vented too.  Even if oxidizer and fuel mix, they aren't contained (even by atmospheric pressure) and so cannot explode, only go poof.  Max Q is long past, so massive structural failure is extremely unlikely.  

Highly visible launch failures always end in explosions because the range safety systems intentionally blow up the rocket (and sometimes SRBs do too) but we never hear about upper stages exploding in space.  They generally fail by shutting down or running out of fuel, stranding the payload in the wrong orbit, or fail to separate.

Offline Smatcha

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Jim - 6/7/2006  9:57 AM

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SMetch - 6/7/2006  12:21 PM

So given that an out of control upper stage could out run any in space designed spacecraft thrust unit the minimum design of an attached spacecraft thrust unit would for final orbit insertion and de-orbit maneuvers correct?

Stated another way after second stage ignition and LAS jettison, if the upper stage fails catastrophically or cannot be shut down the crew would be in trouble.

While our approach would save the crew in that contingency it has a price in terms of weight to orbit regardless of whether we partially fire the LAS motors on ascent or not.  If everyone is fine with the above failure scenario no need to penalize our approach.


Not a credible failure.  Do you actually think Apollo ignored this?  Same with Gemini.  Shutting down engines is easy.  Ask Elon Musk. All engines use valves that need power  to stay open.  Remove power (multiple ways of doing this) and the engine shuts down.


Actually on Apollo there was a memo that seriously questioned whether a safe abort was possible in the event of a main engine failure shortly after lift off and that was based on just the minimum time for the system to perform and escape the shock wave.  Even after a complete system failure in Apollo 12 they still didn’t pull the handle.  There is strong reluctance to chuck billion dollar vehicles at every little indicator light or bump.

In summary crew threatening second stage failures after LAS jettison are deemed to be to remote to worry about.  That’s fine by me.  I’m sure there will be more pressing issues to deal with between the Moon and back.
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Offline pierogoletto

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Kayla - 16/6/2006  1:18 AM

What most people ignore here is that none of us are proposing reducing NASA’s budget.  NASA’s money in the end pays for jobs.  If we spend a lot on SDLV those jobs will be the current jobs tied up in shuttle with very little spent on the exploration end.  If we choose a cheaper alternative launch architecture, NASA will be able to afford more in-space and lunar/Mars work.  Once again this means jobs, not the shuttle jobs but new jobs.  This change scares people.  

What needs to be remembered is that NASA’s money will wind up enabling jobs regardless of the architecture.  With this in mind, we really should be seeking an architecture that creates results, actual astronauts on the lunar surface.  Crewed missions to Mars.  All while not sacrificing the unmanned science.  A NASA that is showing America’s exploration preeminence is much more likely to get Congresses support than a NASA that can barely support ISS.

I think that, although CEV / CALV architecture might in my humble opinion need to evolve, this architecture will create results. Then we need more: a lunar base, then, many years from now, a Mars base; an ISS evolution (from both the European and the American standpoint).
Piero Giuseppe Goletto

Offline yinzer

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SMetch - 3/7/2006  9:34 AM
Explain to me how an upperstage, which needs to be throttled back so as not over accelerating a combinded vehicle beyond 3g’s, does not over take a vehicle at its nose with a thrust to weight less than ½ g.

This has always been a problem.  Even during Apollo the crew would have been killed after the LAS jettison if the 2nd stage could not be shut down.  Given an explosive event in the 2nd stage they would be dead again with the SM engine not capable of getting outside the range of explosion in time.  Either way this has always been a dangerous condition which our solution will solve and at the same time not require carrying the complete weight of the LAS to orbit.  It’s amazing how this danger zone is somehow ignored when discussing steps that must be taken to “man-rate” a vehicle.  In fact this means that for half of the manrated ascent the crew will die, go figure.

How is a second stage going to explode?  There is instantly available stored energy in the PV of the thrust chamber, PV of the ullage volume in the fuel tanks, and kinetic energy of the moving engine parts.  There's less instantly available chemical energy in the propellant flow, and even less available chemical energy in the stored propellant.

The amount of energy available in the first three sources is not enough to damage the payload (as can be seen by the second Delta III failure), and the latter two sources of energy can't be released fast enough to cause an explosion, at least not during upper stage flight.  Shutting down the engines is easy; a combination of spring-loaded normally-closed main propellant valves, shaped charges on the chamber, and pyrotechnically actuated vents on whatever the source for turbine drive gas is will do the trick.
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Offline josh_simonson

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Here's an interesting question:  Does the SM of the CEV have enough delta-V to save it if it were stranded on a lofted trajectory by an EELV failure?  If the SM fails, the crew will be lost anyway, might as well count on it to save the crew in an emergency - 1500m/s delta-V is quite a bit for use in a pinch.

Offline Jim

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josh_simonson - 13/7/2006  5:41 PM

Here's an interesting question:  Does the SM of the CEV have enough delta-V to save it if it were stranded on a lofted trajectory by an EELV failure?  If the SM fails, the crew will be lost anyway, might as well count on it to save the crew in an emergency - 1500m/s delta-V is quite a bit for use in a pinch.

No lofting.  That has be solved.

Offline yinzer

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The lofting can be solved - requires either reduced payload or additional upper-stage thrust.  Lockheed explicitly considered both, I'm not sure what Boeing did.  In any event, the current CLV/CEV combo relies on the SM motor during aborts to keep the capsule out of the North Atlantic during winter.

What could, and probably should (IMHO) be done, is to provide triple-redundant IMU/RCS so that they can control the entry and fly a lifting trajectory.  This would pay off especially well with the biconic/ellipsled entry vehicles and their higher L/D.
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Offline Jim

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yinzer - 13/7/2006  7:03 PM

The lofting can be solved - requires either reduced payload or additional upper-stage thrust.  Lockheed explicitly considered both, I'm not sure what Boeing did.  In any event, the current CLV/CEV combo relies on the SM motor during aborts to keep the capsule out of the North Atlantic during winter.

What could, and probably should (IMHO) be done, is to provide triple-redundant IMU/RCS so that they can control the entry and fly a lifting trajectory.  This would pay off especially well with the biconic/ellipsled entry vehicles and their higher L/D.

there will be a backup RCS/control system on the CEV

Offline yinzer

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Jim - 13/7/2006  5:15 PM

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yinzer - 13/7/2006  7:03 PM

The lofting can be solved - requires either reduced payload or additional upper-stage thrust.  Lockheed explicitly considered both, I'm not sure what Boeing did.  In any event, the current CLV/CEV combo relies on the SM motor during aborts to keep the capsule out of the North Atlantic during winter.

What could, and probably should (IMHO) be done, is to provide triple-redundant IMU/RCS so that they can control the entry and fly a lifting trajectory.  This would pay off especially well with the biconic/ellipsled entry vehicles and their higher L/D.

there will be a backup RCS/control system on the CEV

But not a full-flegded backup RCS/control system.  In the ESAS report, they used the requirement for passive entry to eliminate slender body re-entry vehicles, but then when it turned out that they couldn't guarantee safe re-entry of the capsule, they added the back-up RCS.
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Offline Jim

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with its own power supply and controls

Offline Smatcha

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yinzer - 13/7/2006  5:28 PM

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Jim - 13/7/2006  5:15 PM

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yinzer - 13/7/2006  7:03 PM

The lofting can be solved - requires either reduced payload or additional upper-stage thrust.  Lockheed explicitly considered both, I'm not sure what Boeing did.  In any event, the current CLV/CEV combo relies on the SM motor during aborts to keep the capsule out of the North Atlantic during winter.

What could, and probably should (IMHO) be done, is to provide triple-redundant IMU/RCS so that they can control the entry and fly a lifting trajectory.  This would pay off especially well with the biconic/ellipsled entry vehicles and their higher L/D.

there will be a backup RCS/control system on the CEV

But not a full-flegded backup RCS/control system.  In the ESAS report, they used the requirement for passive entry to eliminate slender body re-entry vehicles, but then when it turned out that they couldn't guarantee safe re-entry of the capsule, they added the back-up RCS.

Only if the capsule has shallow side walls, it becomes more stable the closer the CG is to heat shield and the steeper the side walls, just like the Russian capsule.

Either way there must be a DV from the SM that will significantly improve the abort scenarios while minimizing impact on the EELV's optimal lift capabilities.  Ideal the DV required for final orbit insertion could be used in an emergency to improve the abort trajectories entry interface location and angle of attack.

Do you have any knowledge along these lines?  It would seem the 500 m/s would be enough for just about any scenario.
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