Author Topic: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)  (Read 66634 times)

Offline Chris Bergin

As always, keep it on topic.

DIRECT discussion thread: http://forum.nasaspaceflight.com/forums/thread-view.asp?tid=5016&start=1 - The DIRECT site appears to be down at the moment, I'm trying to find out why.

*Places PAO hat on* And now over to Dr Stanley for opening statements. *Takes PAO hat off*

Offline Doug Stanley

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RE: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #1 on: 01/13/2007 04:43 PM »
Hello friends...Thank you for all of your kind words after the last Q&A I did in late-November!  Chris has asked me back to do another Q&A and put me in the firing line on a very non-controversial topic -- The "Direct Launcher" proposal described at www.directlauncher.com and discussed extensively in the forums on this site.  

During my last Q&A, Ross Tierney, who has been the main spokesperson for the concept, personally asked me to look over it, analyze it, and let him know my findings.  I told him I would do so during the Christmas break.  I then shared the many issues I found directly with him to give him a chance to respond and make sure there were no misunderstandings.  In particular, I found one major problem that we both agree makes the concept infeasible (this will be discussed below in more detail) -- the performance assumptions for the RS-68 engine with a regenerative nozzle.  

Before getting into the technical issues, I would like to take a moment to thank Ross for his tremendous interest in this area and his sincere concern to try to make our exploration efforts as capable and efficient as possible.  I know his motives are pure, and he has been extremely open, helpful, and kind to me in all of our off-line discussions!

Offline Doug Stanley

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RE: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #2 on: 01/13/2007 09:37 PM »
As I noted above, there are a number of major problems with the "Direct" concept and claimed advantages, both technical and cost-related, but I will begin with the largest technical one, the RS-68 REGEN engine assumptions.  The Direct vehicle concept was analyzed by NASA and reviewed by me using the same models and assumptions used by ESAS and the current Ares V program to ensure apples-to-apples consistency.  When this was done the vehicle came up way short of the claims for payload capability.  It was over 16 tonnes short to LEO and 19 tonnes short to TLI -- per launch.  This means that the approach is not even close to being able to launch the lunar mission in two launches.  

One of the main reasons for this is the use of RS-68 Regen vacuum specific impulse assumptions of 435 sec rather than less than 420 sec as verified with NASA and Pratt&Whitney/Rocketdyne.  The thrust/weight was also found to be higher than possible.  Detailed performance analysis was run at the maximum power setting of the RS-68 and using a regen nozzle to gain additional performance.  This was run by the RS-68 contractor and verified by NASA's own internal analysis.  After consultations with the VP in charge at Pratt&Whitney and the Ares V Program Manager, they agreed to let me publicly release these performance numbers so they would not be from "anonymous sources".  They are willing to stand behind these numbers (posted below)!  Neither they nor I have been able to find any possible source of confusion that could have led to anyone quoting the higher numbers, which Ross said he got second hand from a source that is anonymous to me.

The second less major reason for the performance shortfall is in the assumptions about the amount of propellant that the core can actually hold.  Because of geometric and structural considerations, a core vehicle of this diameter cannot hold more than 1.6Mlbs of propellant due to tank clearance issues with the structure required to react SRB loads and thrust structure packaging considerations. A very detailed analysis of this was done during ESAS when we did the LV 24/25 concepts.  For a given diameter core, going to 5-segment solids has the added advantage of allowing you to make the hydrogen tank longer to hold more propellant.  

There are a number of more minor discrepancies, but Ross and I agreed that the RS-68 issue alone is enough to make the concept infeasible, so I won't bother discussing them here.

Offline Doug Stanley

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RE: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #3 on: 01/13/2007 10:10 PM »
I don't want to be pulled into a discussion of costs, because the numbers are SBU, but the claims of significant cost savings are just not correct.  I have every cost number we generated at the most detailed level.  I don't recognize at all the claimed numbers for CLV or CaLV DDT&E costs.  The DDT&E costs for the ESAS CLV is a factor of over 3 less than $16.8B.  Even if you add up all of the DDT&E plus the three test flights and all KSC ops costs and other "keep alive" costs, they don't add up to a number like that.  
 
The cost of developing a heavy lift vehicle big enough to do the ESAS lunar mission in two launches (which the one in the proposal is not capable of doing) will be a few billion more than the baseline ESAS CLV.  You need a vehicle at least as big at the LV 24/25 vehicle from the ESAS report.  BTW, despite the claims in the proposal, we did look extensively at this issue.  The 24/25 vehicle was not quite large enough (but very close) to launch the final ESAS lunar mission even with the suborbital burning of the EDS.  There was also a bit of a mismatch between the needed launch mass for each vehicle (EDS on one and LSAM/CEV on the other -- as there also is with the Direct concept). Even doing the "split" LOR mission with two smaller EDS was marginal and this was unattractive for other reasons...Hence, we needed a larger vehicle (with 5 segment solids) to do the two-launch solution with sufficient performance margin.  

We could back off in requirements and make it fit on the 24/25 system with 4-segment solids.  This was still a possible option at the end of ESAS, since we selected a 4-segment CLV. One reason we selected a 4-segment CLV was to preserve the maximum flexibility in future CaLV decisions.  Although the final report presents the 1.5 launch solution as the preferred option, we preserved the option of eventually going to the two-launch option (with 5 or 4 segment SRBs) which also looked quite attractive. Our intent was that further study would confirm the best answer.  When ESMD later changed to a 5-segment solid on the CLV (which I don't necessarily agree with), the four segment option was taken off the table...although the 5-segment booster CaLV two-launch option remains...but NASA prefers the 1.5 launch option because it 1) has a better LOC and LOM, 2) can land significantly more cargo on the lunar surface in a single launch, 3) provides a larger vehicle for Mars missions, and 4) allows you the flexibility of launching humans to LEO or cis-lunar space without having to carry cargo also with a more costly vehicle.  

As the Direct proposal points out correctly (and as we did in the ESAS report), the two-launch option has somewhat lower annual operating costs and life cycle costs.  If Ares 1 is developed and then decommissioned, the 2-launch option has a LCC of a few billion less that the 1.5 launch option.  
 
This gets to the main unique issue with what was being proposed by Direct...why not forego the development of a CLV altogether and save some money.?  Although this is a moot point now, we did look at this, and the life cycle cost savings (purposely vague...) was less than $5 billion (not $17B).  The problem is that, despite any of your claims to the contrary, heavy-lift vehicle capable of a two-launch lunar solution cost a few (purposely vague...) $Billion more through first human flight due to the higher DDT&E cost relative to the simple ESA CLV and the higher costs of the 3 flight tests.  Despite assertions to the contrary the Direct core vehicle or the ESAS 24/25 core vehicles are almost completely new hardware developments with little STS ET heritage other than materials and diameter because of the different load paths.  The development time was also found to be over 2 years longer (using detailed apples to apples schedules).  The schedule driver now is actually the available budget...not technical considerations.  NASA wanted to close the gap as quickly as possible and wanted a system to go to the ISS with high safety (the ESAS CLV is significantly higher LOC than a HLLV).

Offline Doug Stanley

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RE: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #4 on: 01/13/2007 11:21 PM »
I will also take this opportunity to address on the record some of the alleged "issues" with the ARES 1 vehicle from "anonymous sources" that have been discussed in this forum and certain NASA-Related-Personal-Axe-to-Grind-Single-Source-is-Good-Enough-Blog sites.

An entire section of the Direct proposal is devoted to alleged "Flaws With the Ares Launch Vehicle Family". Very little that was written in the section concerning "problems" with the current Ares Program is correct. The premise that the current (or original ESAS) Ares 1 approach is "broken" and needs to be "fixed" by something like what is being proposing is simply not correct! I will attempt to address some of them in this section for the record in one place. All of the data in these responses come directly from the knowledgable NASA people in the responsible engineering or program office...

DIRECT Assertion: “The original design, of 4-segment SRB with Space Shuttle Main Engine Upper Stage, would probably have lived up to expectations - if the SSME could have been air-started. It can not. NASA is left with a compromise which attempts to fulfill the same requirements, but which fails to.”

NASA Response: This is not true. NASA was confident in its plan to air-start the SSME and no showstoppers were identified at the time NASA elected to change the Ares I baseline. NASA switched to the 5 segment/J-2X approach to achieve greater commonality with the Ares V, reducing the number of developments required - resulting in significant development and recurring costs savings (billions). This included moving from 2 SRB’s (4 and 5 segment) to one (5 segment), 2 upperstage engine developments (alt start SSME and J-2X) to one (J-2X), and moving to a low cost, commercially developed core stage engine flying on the Delta IV today (RS-68) vs. an engine unique to NASA needs (SSME derivative).

DIRECT Assertion: “The “Stick” Crew LV's biggest selling point was its high safety figures. However, the difficulties the design is suffering from today are continually whittling those away, with each 'fix' causing ever larger penalties to the performance.

NASA Response: This is not true. NASA currently projects a loss of crew of 1 in 2,150 - a robust vehicle when compared to STS and with any other alternatives evaluated and consistent with ESAS projections.

DIRECT Assertion: “The new 5-segment SRB's and J-2X engines are both completely unproven.”

NASA Response: This is not true. A 5 segment ground test motor was fired in October, 2003. The J-2X is a derivative of the Saturn J-2 and J-2S engines, elements of which (turbopumps) were recently utilized on the X-33.

DIRECT Assertion: “Together their performance is so desperately low that other parts of the vehicle are having to be designed down to dangerously minimal weight, in order just to get the system to fly at all.” Performance of just 22mT -30x100nm 28.5deg is at best, mediocre, at worst, anæmic. This poor performance is causing detrimental domino effect throughout every phase of the development of the new vehicle.”

NASA Response: This is not true. The original ESAS baseline CLV delivered ~27mT (without performance margin) to LEO/28.5°. This was with a much lighter launch abort system and before wind tunnel data was available. Using much more detailed models, the current Ares I is projected to deliver ~26mT (without performance margin) to LEO/28.5° - equivalent to ESAS. Orion is being designed to weigh no more than 22mT (in ESAS, this was 23mT, but was a 5.5m diameter capsule). Ares I will be the largest heavy lift capability in the U.S. until Ares V is developed.

DIRECT Assertion: “A normal rocket is naturally stabilized throughout its flight by having the Center of Gravity (CofG) ahead of the Center of Pressure (CofP). Like a thrown dart, the rocket will naturally fly nose-first. But the Ares-I's CofG is behind the CofP - which causes the rocket to want to flip around in mid-air. Only with very precisely applied Thrust Vector Control, can the rocket be kept on track without applying very high stress loads to the structure. The first stage has a very slow Thrust Vectoring system, simply because it is a Solid Rocket Booster. This is causing concern during the first minute after launch, before speed builds and aerodynamics affect the ascent. It is the job of the SRB's Thrust Vectoring system to keep the very tall and ungainly rocket stable and pointing in the right direction as it lifts from the Pad. It is a problem often equated to balancing a pencil, on end, using your finger. The nozzle at the bottom of the SRB is proving to be a very slow ‘finger’ performing the balancing act. If the rocket becomes unbalanced, perhaps due to crosswinds, the nozzle may be too slow, and be forced to apply very high bending moment forces on the structure in order to try to re-stabilize.”

NASA Response: This is not true and shows a lack of understanding of large rocket design. Typically, large, orbital capable rockets have a C.G. aft of the C.P., hence you utilize a TVC system. NASA has conducted over 1,500 wind tunnel tests of the Ares I configuration, and conducted analyses on the flight control system design. While Ares is a long and slender vehicle, it is within the control dynamics experience base of previous programs, most notably the Saturn V. 6DOF simulation results indicate a ~2x margin on first stage thrust vector control (angle and rate) and an ~8x margin on the vehicle structural response to control frequency ratio.

DIRECT Assertion: “The two issues above can cause forces which, quite literally, try to bend the vehicle in half. The SRB is a very strong structure. The pressurized Upper Stage tanking is also a very strong structure. But the Interstage between them is a hollow cylinder, 18ft (5.5m) wide, and 40ft (12m) long, with walls only 1.25" (3cm) thick - and complicated further by a conical structure changing diameter from 13ft (3.9m) to 18ft (5.5m). The Interstage will be the “weak point” if the vehicle suffers instability issues during flight. It is the structure which would fail first if the rocket goes off-course and takes too much time to be forced back on course. The Ares-I test vehicles’ Interstages are being specifically over-built to combat this problem in a bid to dissuade disparaging comment from the space community, who is already well aware of this concern. But the final flight versions of Ares-I must be built down to the lowest possible weight limits in order to keep performance high enough - which means this will be the weakest structural point in the final design. The SRB first stage is currently 18,000lb overweight because the seals around all of the segments need additional, unplanned, strengthening. This is because the in-line design, with the stage and payload located above the booster instead of beside it, are experiencing different loads during flight from the SRB’s intended design - so require additional strengthening at these joints to compensate.”

NASA Response: This is not true. While the Shuttle RSRM was not originally designed to have a second stage ride atop it on the way to orbit, this is a very robust stage which carries the entire load of the Shuttle External Tank and offset load of the Shuttle Orbiter. In addition, it was sized to carry the offset load of 3 Space Shuttle Main Engines firing at ignition (“twang load”) which Ares will not have due to its single engine, in-line first stage configuration. The NASA team is using proven, validated engineering tools and loads models and conservative margin factors at this stage of the design. Analyses performed to-date indicates that the existing Shuttle RSRM cases, joints and aft skirt have sufficient design capability to support the Ares in-line configuration and are not “overweight” as characterized above. In addition, the upperstage and interstage are being designed for the loads expected on the ground and in-flight. The upcoming Ares I-1 flight test in 2009 will give NASA important data early in the development cycle.

DIRECT Assertion: “The roll-control system was not predicted to be as considerable an issue as it is proving to be. It requires an extra system which was unplanned originally, which impacts the weight of the vehicle, and increases the number of systems which can cause an expensive Loss of Mission or, worst of all, a Loss of Crew contingency.”

NASA Response: This is not true. Characterizing and controlling roll torque has been a high priority since Ares’ inception. NASA has utilized what it believes are worst case roll torque predictions and then designed the control system to handle 1.7 times that torque using RCS thrusters. Our goal now is to further refine the roll torque predictions through ground test firings of the motors with calibrated sensors, analyzing similar launch systems (Athena, for example) and the Ares I-1 flight test in 2009. We believe we have utilized very conservative predictions and then used a conservative design approach.

DIRECT Assertion: “The original “Stick” launcher utilized the Upper Stage to reach an initial elliptical orbit of 60x160nm, then that Upper Stage to then perform the Circularization burn to achieve the stable 160x160nm orbit. The Orion is now required to perform a 1000ft/s high-Delta-V burn to reach an initial orbit of just -30x100nm - that means the low-point is 30 nautical miles under the Earth’s surface.”

NASA Response: This is not true. In ESAS and until last Spring, the Ares I injected the Orion into a 30x160nmi transfer orbit and the Orion then circularized itself, to avoid the complexity of deorbiting the large upperstage. Working with Constellation and CEV project teams, the program elected to change to a -30x100nmi orbit to move the ocean impact of the CLV upperstage to the Indian Ocean from the South Pacific to stay away from populated islands. This also allowed the impact point for both ISS and lunar missions to be in the same general vicinity. Appropriate performance was transferred from Ares to Orion so that the spacecraft was not penalized. Performing multiple OMS types burns is commonplace on STS today and does not increase risk. Also, Orion does not have to do a burn to reach -30x100nmi - Ares places it in that orbit. Orion carries 1,000 ft/s to perform all orbital maneuvers, including transferring from -30X100 to 160 circ, and on to 220 for ISS, rendezvous, prox ops and docking, and deorbit.

DIRECT Assertion: “Together, this reduces the original “Stick” concepts Loss of Crew (LOC) figures below the stated 1 in 1918. The Ares-I’s fundamental design requires that the Upper Stage engine be ignited at altitude, only after the SRB First Stage has burned-out. There is no guarantee that any engine will start correctly, or safely, let alone at altitude. If there is a problem, the mission would become an abort, requiring the use of the escape system. NASA has yet to publish new, independent, ‘apples-to-apples’ comparison safety figures between the original CLV and the current evolution. The figures will obviously be lower today. Loss of Crew (LOC) safety figures of between 1 in 1500 to 1600 are rumored for the current risk factor as this paper was compiled - so the gap to DIRECT’s 1 in 1355 LOC risk is now very narrow indeed.

NASA Response: This is not true. The current Ares probabilistic risk assessment, which is much more comprehensive than the PRA used in ESAS is predicting a loss of loss of crew of 1 in 2,150 (mean). This is ~1.6x the DIRECT claim of 1 in 1,355 LOC (which is not supported with any analysis - the best ESAS vehicle in this “direct” class had an LOC of 1 in 1,170).


Offline Doug Stanley

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RE: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #5 on: 01/13/2007 11:22 PM »
I guess that is enough for now...At this point I will be happy to take your questions...

Sincerely,

Doug-I-Read-It-On-The-Internet-So-It-Must-Be-True-Stanley

Offline Chris Bergin

RE: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #6 on: 01/13/2007 11:57 PM »
Now live and open for questions....

Offline Smatcha

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #7 on: 01/14/2007 12:31 AM »
First Question:

Did you run a 4xRS-68Regen, 4 Seg SRB under a 8.4 meter core with an optimized upper stage?

Second Question:

Your “Maximum” Tank numbers for a 8.4m “is” the current SSTS tank capacity.  You could add about +60,000kg to tank by extending the LH2 tank down and LO2 tank up and still have room for the RS-68’s and Thrust structure.
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Offline JRThro

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #8 on: 01/14/2007 12:42 AM »
I'm not knowledgeable enough to ask an intelligent question, but I have to say that I appreciate Dr. Stanley's candor and his willingness to discuss DIRECT and the Ares program in this public forum.

Well, I will ask a question about what Dr. Stanley said in the following text bite, about the Orion's ability to circularize the orbit, etc.:
"Appropriate performance was transferred from Ares to Orion so that the spacecraft was not penalized. Performing multiple OMS types burns is commonplace on STS today and does not increase risk. Also, Orion does not have to do a burn to reach -30x100nmi - Ares places it in that orbit. Orion carries 1,000 ft/s to perform all orbital maneuvers, including transferring from -30X100 to 160 circ, and on to 220 for ISS, rendezvous, prox ops and docking, and deorbit."

Dr. Stanley, does this mean that the Service Module engine, or the amount of fuel carried onboard the Orion, have been modified in order to transfer the needed performance to Orion?  Can you expand somewhat on this point?

Thank you for your time.

-John Thro
-Houston, TX

Offline Scotty

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #9 on: 01/14/2007 12:57 AM »
About the existing RS68 engine.
Is or is not the following statement true?
The hydrogen now enters the combustion chamber from the fuel injectors about 20 degrees R colder than originally predicted.
This colder than expected hydrogen then negatively impacts the "C sub star" portion of the Isp equation.
Thus the lower than expected Isp for the existing RS68 engine.
A regenerative nozzel would add heat to the hydrogen, greatly improving the C sub star portion of the Isp equation.

I totally agree an Isp of 450, as many have stated as possible with an evolved RS-68, is flatly impossible.
But it sure looks like 430 to 435 is possible.


Offline Doug Stanley

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #10 on: 01/14/2007 01:41 AM »
Quote
Scotty - 13/1/2007  7:40 PM

About the existing RS68 engine.
Is or is not the following statement true?
The hydrogen now enters the combustion chamber from the fuel injectors about 20 degrees R colder than originally predicted.
This colder than expected hydrogen then negatively impacts the "C sub star" portion of the Isp equation.
Thus the lower than expected Isp for the existing RS68 engine.
A regenerative nozzel would add heat to the hydrogen, greatly improving the C sub star portion of the Isp equation.

I totally agree an Isp of 450, as many have stated as possible with an evolved RS-68, is flatly impossible.
But it sure looks like 430 to 435 is possible.


Yes...of course it is true that a regen nozzle heats the H2 by circulating around the nozzle and then when it is combusted it improves C star.  That is exactly what happens.  That is why the Isp with the Regen nozzle is a good bit higher (10 sec) than the baseline RS-68.  A full cycle balance and thermal flow analysis was done on the engine by the responsible experts, and the result is the posted performance...not anywhere near 435 sec.  The company and NASA have agreed to release these results to the public and stand behind them.  I don't know what more you want, neither do I understand the analytical basis for your statement that "it sure looks like 430 to 435 sec is possible"...

Offline Scotty

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #11 on: 01/14/2007 01:51 AM »
The basis for my statement is many long conversations with one of the fellows who did design work on the RS68's pumps and injectors.
Until I (and he) see and review the analysis data (that I feel will never happen), I have to go with what he has shown me in dry hard math.

Offline Generic Username

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #12 on: 01/14/2007 01:58 AM »
I know you said you diodn't want to get into costs, but I can't help myself.  :)  Recently there have been claims that the CLV first stage -1st stage, not whole vehicle - is going to cost $5billion to develop. How accurate, or inaccurate, is this number?
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Offline Paul Howard

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #13 on: 01/14/2007 01:58 AM »
I'd just like to thanks Dr Stanley for coming on this site again for a Q&A. This is the only site where this happens with NASA etc. people, and long may it continue.

And on that point,

Quote
Scotty - 13/1/2007  8:34 PM

The basis for my statement is many long conversations with one of the fellows who did design work on the RS68's pumps and injectors.


I'd also hope that when someone of Dr Stanley's standing comes here with actual figures, as posted, that we don't see too many of these "But I had a chat with an engineer that says different".

As Chris likes to say all the time, Documented Figures only please.

Offline Doug Stanley

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #14 on: 01/14/2007 02:01 AM »
Quote
SMetch - 13/1/2007  7:14 PM

First Question:

Did you run a 4xRS-68Regen, 4 Seg SRB under a 8.4 meter core with an optimized upper stage?

Second Question:

Your “Maximum” Tank numbers for a 8.4m “is” the current SSTS tank capacity.  You could add about +60,000kg to tank by extending the LH2 tank down and LO2 tank up and still have room for the RS-68’s and Thrust structure.

Thanks for the questions...

Yes, NASA recently ran a number of different combinations of RS-68's, upper stages, 4 and 5 segment solids, at both diameters.  For the 4 segment boosters, ET diameter, and even 4 RS-68s, there was not enough payload to do the 2-launch solution with 10 percent payload margin and still fit within the VAB.  You reach diminishing returns by adding first stage engines with 4-segment boosters.  With 4 engines, the propellant just burns out twice as fast and you have higher lift-off and first stage accelerations.  The higher g's at lift off, coupled with dynamic loads, add additional compressive loads on the intertanks and H2 tank with a full O2 load on top, which adds structural weight.  Yes, th epayload goes up, but not by enough.  If you go to 5-segment solids (which allows additional first stage propellant and larger diameters, you can get RS-68-based vehicles that can do the two-launch solution...and this could still be on the table...

Given the RS-68 engine length and required pad clearances, NASA designed and laid out an aft end with thrust structure/gimbal/lines and found that there really was no room left to drop the H2 tank by any useful degree...

Offline Doug Stanley

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #15 on: 01/14/2007 02:11 AM »
Quote
JRThro - 13/1/2007  7:25 PM

I'm not knowledgeable enough to ask an intelligent question, but I have to say that I appreciate Dr. Stanley's candor and his willingness to discuss DIRECT and the Ares program in this public forum.

Well, I will ask a question about what Dr. Stanley said in the following text bite, about the Orion's ability to circularize the orbit, etc.:
"Appropriate performance was transferred from Ares to Orion so that the spacecraft was not penalized. Performing multiple OMS types burns is commonplace on STS today and does not increase risk. Also, Orion does not have to do a burn to reach -30x100nmi - Ares places it in that orbit. Orion carries 1,000 ft/s to perform all orbital maneuvers, including transferring from -30X100 to 160 circ, and on to 220 for ISS, rendezvous, prox ops and docking, and deorbit."

Dr. Stanley, does this mean that the Service Module engine, or the amount of fuel carried onboard the Orion, have been modified in order to transfer the needed performance to Orion?  Can you expand somewhat on this point?

Thank you for your time.

-John Thro
-Houston, TX

Yes, since ESAS, some of the performance requirements have been transferred from the ARES 1 to the Orion.  The service module of Orion must carry somewhat more propellant and the ARES 1 upper stage less...

Offline Doug Stanley

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #16 on: 01/14/2007 02:17 AM »
Quote
Scotty - 13/1/2007  8:34 PM

The basis for my statement is many long conversations with one of the fellows who did design work on the RS68's pumps and injectors.
Until I (and he) see and review the analysis data (that I feel will never happen), I have to go with what he has shown me in dry hard math.

If he works for Pratt&Whitney Rocketdyne he already has access.  What specific "hard dry math" did "he" show you.  I will try to get the same info.  Do you want C Star efficiencies, temperatures, pressures.  I don't know your aerospace education background.  Are you familiar with Red-Top or other propulsion models.  I may be able to provide you a generic model with the appropriate data.  JUst let me know...I will do my best to make it happen!

Offline Doug Stanley

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #17 on: 01/14/2007 02:19 AM »
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Generic Username - 13/1/2007  8:41 PM

I know you said you diodn't want to get into costs, but I can't help myself.  :)  Recently there have been claims that the CLV first stage -1st stage, not whole vehicle - is going to cost $5billion to develop. How accurate, or inaccurate, is this number?

If you are talking about just the development of the first stage...that is significantly inaccurate!! Where did you get that number??

Offline Generic Username

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Re: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #18 on: 01/14/2007 02:36 AM »
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Doug Stanley - 13/1/2007  8:02 PM

If you are talking about just the development of the first stage...

Yes.

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that is significantly inaccurate!! Where did you get that number??

I honestly don't know where it originated. But numbers of that ROM are now seemingly the "conventional wisdom" as to how much it is going to cost (not just here, but floating all over the net), thus "proving" that just staying with the 4-segment will automatically save that many billions at least. If you could set the record straight, at least an approximation as to 1st stage development cost, it'd be much appreciated.

Spacecraft and aircraft models, blueprints, documents:
http:www.up-ship.com

Offline marsavian

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RE: Q&A: ESAS Lead - Dr Doug Stanley (On DIRECT)
« Reply #19 on: 01/14/2007 02:42 AM »
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Doug Stanley - 13/1/2007  4:53 PM

 Despite assertions to the contrary the Direct core vehicle or the ESAS 24/25 core vehicles are almost completely new hardware developments with little STS ET heritage other than materials and diameter because of the different load paths.  The development time was also found to be over 2 years longer (using detailed apples to apples schedules).  The schedule driver now is actually the available budget...not technical considerations.  NASA wanted to close the gap as quickly as possible and wanted a system to go to the ISS with high safety (the ESAS CLV is significantly higher LOC than a HLLV).

Firstly once again thanks for taking time out on a weekend in your spare time to answer Ares questions. If nothing else DIRECT has got more people interested in VSE and ARES and the detailed very interesting information you have provided in rebuttal will only further that interest and I believe support in VSE in the long run. I have two clarification questions arising from the above quote. Are you saying that a DIRECT vehicle would   take 2 years longer to develop then the current Ares I or the original 4 segment RSRB/SSME Ares I ?  As to the schedule driver, even with an unlimited budget, how soon could an Ares I be built considering the 5 segment RSRB and J-2X need developing and as such would have serial technical critical paths ?

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