The Ergosphere
Tuesday, August 02, 2005
 

It's about time

NASA is finally going to quit messing around with the Model T of space vehicles and go back to what works; the Shuttle's replacement will be two vehicles, one heavy-lift cargo hauler and one much smaller people-carrier.  Both will use Shuttle SRB's, and the heavy-lifter will use SSME's as well.

NY Times coverage.

(Hat tip:  Slashdot.) 
Comments:
Wonderful. Now we'll never get back to the Moon.
 
Never?  This re-establishes a Saturn-class lifter, which is a prerequisite for manned missions beyond LEO (the Shuttle is quite incapable of such travel).  Abandoning Shuttle is the first step on the road back.
 
Two problems.
1. The simplest problems are that this Saturn-class booster is not a) man-rated or b) designed to go to the Moon. It's designed to put large cargoes in LEO. This gets rid of one of the main advantages of the Shuttle--that you could put something large in orbit (like, say, space station components) and do something with it, and gets us no closer to the Moon.
2. It diverts money, time, and people away from developing an actual lunar program. It took eight years and change from the last landing on the Moon to the first Shuttle launch. Most of the engineering was done after 1975 because people were working on Skylab and ASTP. But the real problem here is that it took years to develop Shuttle. It will take years to redesign it as two separate spacecraft. It will then take years to land on the Moon again. Realistically, what is this vehicle going to do? If we abandon the Shuttle after the ISS is complete, and then want to land on the Moon, what is a heavy-lift LEO launcher going to do?

This represents a political decision. Somebody in power doesn't want to land on the Moon, be it George Bush, NASA management, or some other random turkey.
 
The shuttle is gigantic waste of money and I am glad to see it go.
I thought it was a good idea when I was 10.
If you think about it, they are unnecessarily returning to earth a large cargo hold. I am guessing that more than 20 times extra heat shielding is required than for just human re-entry. The heat shield is the most likely thing to fail on re-entry and its 20 times bigger than necessary.
Isn't one shuttle burn up enough for you?
Besides, I read somewhere that the current design was politically motivated. Something about having a cargo hold big enough to bring military satelites(sp) back in secret. I don't know if they ever used that capability, but let the military take those kind of risks when necessary.
 
What we have to do first is use the Shuttle to build the rest of the ISS, then retire it and go to the Moon.

The Shuttle was a compromise NASA-USAF design which was abandoned by USAF, meaning NASA is stuck with many design features they didn't like and can't use. A small manned capsule should have been put on a simple panel-based payload bay using a reusable LOX/LH2 staged system (probably 1 1/2 stage like the Shuttle). This would have allowed large payloads to be taken into orbit, and with modifications to the payload bay, back down. If we had only known in 1972.

Why, again, does building the Shuttle (almost) right the second time get us closer to the Moon?

Two shuttle accidents are more than enough for me. Apollo 17 was enough for me. We never should have ended the Apollo program. We should have just kept upgrading and kept going. The way it happened, we wasted all the infrastructure and R&D money by letting the results rot. Most of the engineers who were involved in building Apollo died with their knowledge of the systems. Redoing it today, we'd have to start from scratch and in today's sociopolitical environment, it would take longer to get back to the Moon than it took to get there the first time.
 
Goodness, Stewart.  It's been ages since I was a true space geek and I can still pick your objections apart point by point:

"1. The simplest problems are that this Saturn-class booster is not a) man-rated or b) designed to go to the Moon. It's designed to put large cargoes in LEO."

Curious you should say that.  Every part of the stack save the new fairing, from hydrogen motors to SRB's, is man-rated.  The Saturn V booster and the second stage happened to do nothing more than put the third stage of the stack into orbit.  On top of this, the second-stage motor for the manned booster is a re-engineered J-2... the same motor used in the two hydrogen-powered Saturn stages.

In short, you could trivially man-rate the heavy booster; you could put the manned capsule and its escape system on top of a cargo fairing.  Or you could just ignore the issue and launch your crew on a second vehicle, hooking up in orbit.  It's not like we forgot how to do a rendezvous.

"2. It diverts money, time, and people away from developing an actual lunar program."

As opposed to the Shuttle program, which did even more of the same without providing a vehicle which could get people anywhere near the moon.

"But the real problem here is that it took years to develop Shuttle. It will take years to redesign it as two separate spacecraft."

Is that so?  What part?

1.  The SSME's are done.
2.  The SRB's are done.
3.  The ET would no longer require insulation to prevent icing, or fancy non-shedding foam.
4.  We no longer need wings, landing gear or most of the other Shuttle systems.
5.  Are you seriously arguing that we no longer know how to build an Apollo-class re-entry capsule?!

If you gave this job to an outfit like Von Braun's NASA or the Lockheed Skunk Works, you'd have the vehicle flying within 2 years.  They have about as much to do as the engineers charged with making a first stage to launch the Apollo-Soyuz mission.  The Saturn 1B was cobbled together from parts of other rockets (link,
link).

"It will then take years to land on the Moon again."

And how long would it take with a Shuttle which can only loft 20 tons (and cannot be uprated) and is due for retirement?

"If we abandon the Shuttle after the ISS is complete, and then want to land on the Moon, what is a heavy-lift LEO launcher going to do?"

I dunno... maybe put the lunar-injection stage, the lunar lander and the return vehicle in place for the TLI burn?

" What we have to do first is use the Shuttle to build the rest of the ISS, then retire it and go to the Moon."

It would make a lot more sense to complete the ISS by building a 100-ton module or set of modules to go up in that payload fairing, and launch the whole kit and kaboodle with one rocket.  (Did you look at the diameter of that thing?  It's bigger than the ET, which is more than twice the size of the payload bay.  You could cluster 4 ISS modules inside it and have space left over for another Hubble telescope.)  You wouldn't have to worry about anyone dying if the launch failed, either.

"We never should have ended the Apollo program."

It's ironic that you lament the decision to return to a proven model of getting to space as "a political decision", then turn around and denounce the political decision which gave us the Shuttle.
 
Every part of the stack is man-rated (almost), but the stack itself isn't. Systems integration is a huge challenge in aerospace. Furthermore, you can't stack a payload on top of the ET as it is now. It was not designed to bear loads that way. 'Re-engineered' is the key word here. The fact that the system requires massive changes would make it a lot further away from being man-rated than you think.

Let me make my views clear on the Shuttle. We need it to build ISS. We needed it to launch and maintain large payloads, like Hubble, for example. We don't need it to go to the Moon and in fact can't go to the Moon while we have it. What we really don't need is to build the Shuttle semi-right the second time. Then we're stuck with a system we shouldn't have built, can't use, and prevents us from doing anything while we have it. We need to complete the ISS and then go to the Moon using a totally different vehicle designed for that purpose. If I was not clear before I hope I am now.

The components are done, to a certain extent. The vehicle isn't. Cobbling something together that was designed to work in a specific configuration is harder than you think. If this was actually done, we would soon find that we needed to redesign many of the "done" components.

Yes, I am saying we no longer know how to build Apollo. Most of the plans were destroyed. Most of the engineers who knew the systems are dead. The technology is very different. We would have to start off from scratch.

The Saturn 1B was designed in the 1960s to test the Saturn V's engines and to work as a LEO launcher for Apollo. It was most certainly not designed between 1972 and 1975. The V-2 was a "cobbled-together" project--it was necessary because [brilliant insight] rockets don't scale up [/brilliant insight]. It took somewhere around six years. There is no time advantage to combining existing components; at most there is a cost advantage, and the only really effective COTS area is avionics.

We cannot land on the Moon with a Shuttle. Let's build a system that works for the purposes we'll be using it for.

A launcher made from Shuttle components will not work for translunar missions. There's too much propellant in too few stages, for one. A heavy-lift launcher cannot be easily converted (if at all) into a translunar booster. The missions are so fundamentally different that you have to design the boosters differently.

It would make a lot more sense to save $15 billion, use what we have for the next six years, and then build a lunar vehicle. Building something else to finish construction would only delay ISS construction by six to eight years, as well as delaying a Moon landing by an equal or greater amount of time.

The decision to go ahead with a shuttle replacement when we're supposed to be pursuing a lunar landing program would be a political decision that would be indicative of a lack of commitment to a lunar program. The Shuttle shouldn't have been designed the way it was. Now, however, is not the time to correct it when we have a completely different set of mission requirements. Of course we'll learn from it, but we shouldn't pretend it's 1972 and build another Shuttle.
 
The illustration in the times for the lifting vehicle was miss-labeled go to the SafeSimpleSoon.com web-site.

Personally, I am thrilled that people connected to a government program have figured out a way to solve problems by recycling and reusing existing components rather than just throwing wads of money at it.
 
OT, but interesting:
Shortages Stifle a Boom Time for the Solar Industry
 
What exactly has human spaceflight ever done for us peons anyway?
 
Funnily enough I suggested this configuration on the sci.space.tech forum years ago. I thought and still think that some things do not require a reuable shuttle. It is really only humans that require it.
 
The other massive problem is one I have highlighted before. The SSME if that is what is going to be used is right at the bleeding edge of materials and manufacturing. It has only acheived a measure of reliablity because of massive amounts of testing on a known platform.

It is not a trivial task to mount SSME's on a completely new platform. To my mind it would be akin to developing it all over again with all the costs and delays of the intial development.

If we had done this 25 years ago and kept the Saturn 5 as a HLLV and developed a DynaSoar size crew return vehicle perhaps NASA would have a budget now. Really we have completely wasted 25 years and trillions of dollars on a flawed grand experiment and now with the failure of this platform we are basically back before Apollo again.
 
What has human spaceflight done for us?
-MRIs
-Smoke detectors
-The modern (i.e., not teletype-swiches-lights) computer
-Much research in medicine that was done on space stations and shuttles
-Accurate weather forecasting
Just off the top of my head.
 
After properly reading the post and comments again I totally agree with Stewart Peterson.

Using SSMEs is the first critical mistake. This is like putting a 600hp V10 Formula 1 engine in a Mack Truck and expecting it to be reliable. The SSME are the highest performing rocket engines ever designed and stretch material and processes to the edge. You cannot simple bolt them onto a completely different platform and expect them to work. A far far better route would be the lower performance RD-180 used on the Atlas V. Additionally it uses LOX/RP7 so the lower stages can be more compact and easier to build. The Saturn V used LOX/RP7 for the very good reason that LH2/LOX's high specific impulse advantages are lost on lower stages. Trying to re-use shuttle components will doom this path to failure and not save any money or develpoment time. Really it would be better re-engineering Energia as this launch platform has the engines at the bottom of the tank and is designed for it. If you want a cut price low risk path to space fund the resurection of Energia.

The SRBs are man-rated but only over the most strident objections of just about everyone. Again Apollo did not use solids for the very good reason that once you light the blue touch powder you are riding that booster till the fuel runs out. The Shuttle is the first and hopefully the last to use them.

The capsule is just a total throwback. While it is cheap and 'easy' it has no cross range or divert capability. It has to be recovered. A winged re-entry vehicle has the advantage of landing like a plane. If you are only carrying passengers then the extra weight of wings etc is well justified. Looking to the future do you think mum and dad and the kids will want to go to space in a capsule? No way - the will want a aircraft like experience. Spaceship One has shown the way. In 20 years time how will NASA seem using dead end capsules from the 60s when Richard Branson can lift passengers to space hotels in an winged aircraft? NASA should contract the astronaut lifting part to Virgin Galactic.

Lastly the main reason that there are no 100t launchers at the moment is that there are no 100t payloads that need to be launched. All commercial launchers are in the 5 to 20 t to LEO range for a very good reason that the most complex expensive satellite that we can build is about this weight. Sure we could launch 5 of them in one go however if it crashed then this would be catastophic. This 100t launcher would have to be developed with no commercial use in mind totally from government money with no real use after the government is finished with it. Surely it would be better to design the modules for the lunar transfer vehicles with the Delta 4 Heavy and that Atlas V in mind. They can lift nearly 20 t to LEO. 4 launches of a reliable tested commercial rocket would probably be no more that the cost of the new one plus all the development costs and inevitable delays and overruns.

NASA and the whole American space effort is floundering and paying the price of putting all its eggs in the Shuttle basket and a complete lack of vision. Nobody expected the Shuttle to be so expensive and be such a dead-end that it would sink the whole manned space progam like it is danger of doing. Trying to re-use Shuttle components will only doom this effort and be the final nail in the coffin of American manned space flight.
 
Mr. Gloor, I find you saying more and more absurd things with each post.

"It is not a trivial task to mount SSME's on a completely new platform."

I'm sure that it takes reasonable care, but no more than that; if it did, standard articles like test stands would have been useless in its development.

"The SSME are the highest performing rocket engines ever designed and stretch material and processes to the edge. You cannot simple bolt them onto a completely different platform and expect them to work.... To my mind it would be akin to developing it all over again..."

You're saying that:
1.  Replacing the Orbiter from the thrust-frame forward with a wingless cargo pod would be akin to developing the engines all over again.
2.  Moving the engines from a position beside the fuel tank to directly beneath it would be akin to developing the engines all over again.

To make such a bold yet highly questionable assertion you must have much relevant education or experience, like being one of the SSME designers or test engineers.  Would you care to list it?

"A far far better route would be the lower performance RD-180 used on the Atlas V. Additionally it uses LOX/RP7 so the lower stages can be more compact and easier to build. The Saturn V used LOX/RP7 for the very good reason that LH2/LOX's high specific impulse advantages are lost on lower stages."

The Atlas is typically used in conjunction with the Centaur upper stage.  The Centaur uses the RL-10 engine, which burns LOX/LH2.

RD-180:RL-10::SRB:SSME.  Not only is the RD-180 an inappropriate replacement for the SSME, using them on a Shuttle-derived HLV would mean clustering them, with the greater probabilities of failure.

"The SRBs are man-rated but only over the most strident objections of just about everyone."

So, launching the people-carrier with just one of them instead of two (half the probability of failure), and well behind the passengers instead of alongside them, is a BAD thing?

The design does come from the builder of the SRB.  If I were managing this and there were no pork requirements, I'd ask someone to look at pressure-fed "boilerplate" rockets burning cheap room-temperature liquids.  But I'm not, and there are.

"The capsule is just a total throwback. While it is cheap and 'easy' it has no cross range or divert capability."

The Apollo capsule had considerable lift, and hundreds of miles of cross-range capability.  The proposed capsule is a very flattened cone like Apollo and would have similar aerodynamics.  The Shuttle's cross-range capability was demanded by the military for a mission profile that has never been flown and never will be.

What divert capability do you need when you're trying to hit e.g. Nebraska?

"A winged re-entry vehicle has the advantage of landing like a plane."

In other words, if you can't get properly lined up on a nice, long, paved runway, your vehicle is toast and possibly cargo and crew as well.  A capsule needs an appropriate surface roughly as big as its underside; to a Soyuz, steppe is thousands of square miles of landing pad.

So it's not glamorous.  I rather like flying myself, especially those landings when I can just grease it on, but space is a different mileu and calls for different methods.

"Lastly the main reason that there are no 100t launchers at the moment is that there are no 100t payloads that need to be launched."

Not since Apollo.  In other words, you're ruling out both the Moon and Mars a priori.

Before commenting again, I'd appreciate it if you'd (a) RTFA, and (b) think about what you're saying.
 
No I am saying that developing the system that the SSMEs are a part of even with the small changes that you mention COULD take as much time, effort and money to debug than the intitial development of the SSMEs. A test stand is a rigid fixed mount with piping put anywhere it runs. A space vehicle is a compressed volume where fuel lines have to snake around components despite the best efforts of the designers to make them straight. ANY such changes to layout can result in surges and problems that cannot be foreseen. The SSME being such a high performance engine is more sensitive to such changes and will be much harder to re-integrate with the new vehicle. The unmanned cargo shuttle will be less of a problem however moving the engines to the bottom of the ET is a massive change. There are totally different vibration modes and fuel runs with this configuration that will cause problems that will take time and money to solve. My knowledge of the SSME is from a detailed article written by one of the team that designed and tested the SSMEs. Of course I cannot find the article despite tearing my garage apart looking for it - if I ever find it I will email you the title and author. However you will have to trust my flaky memory until then. Basically he says he crosses his fingers every single flight. This is why I do not regard the SSME as the best candidate for the new booster.

Now the RD 180 currently used in the Atlas V
Propellants: Lox/Kerosene Thrust(vac): 423,050 kgf. Thrust(vac): 4,152.00 kN. Isp: 338 sec. Isp (sea level): 311 sec. Area Ratio: 36.87

The SSME
Application: . Propellants: Lox/LH2 Thrust(vac): 232,301 kgf. Thrust(vac): 2,278.00 kN. Isp: 453 sec. Isp (sea level): 363 sec. Burn time: 480 sec. Area Ratio: 77.5.

or even the RS-68 as used in the Delta 4
Propellants: Lox/LH2 Thrust(vac): 337,807 kgf. Thrust(vac): 3,312.00 kN. Isp: 420 sec. Isp (sea level): 365 sec. Mass Engine: 6,597 kg. Area Ratio: 21.5.

Now you will note that the RD-180 is almost twice as powerful as the SSME and the RS-68 2/3s more powerful. Also as the sea level Isp of the the RD-180 is quite close (311 V 363) the advantages of hydrogen propellant is lost when the SSME is used as first stage engine. This was forced on the shuttle as the same vehicle had to be used for the total ascent. As they could not use an altitude compensating nozzle like an aerospike or dual bell they were forced to use a vacuum optimised engine (expansion ratio 77.5) at sea level. This is one reason the SRBs have to be so big. The SSMEs just do not perform well at sea level. Both the the RD-170 and RS-68 are designed with a lower expansion ratio which means that they give more thrust at sea level. In fact the RD-180 or RS-68 is a far better choice than the SSME for a first stage of a booster.

Possibly the most damning argument against the SSME is that the 2 most recent Amercan boosters designed after the shuttle NEITHER used the SSME as the main engine. If what you and the project authors say is true that it is cheaper to use existing hardware why did neither of the design teams for the Delta 4 or Atlas V use the SSME? Both are of a size that would use an engine of the thrust class of the SSME yet neither did. The Delta 4 used a totally new engine the RS-68 and the Atlas V uses the RD-170. The Delta 4 design team preferred to design a brand new engine from scratch rather then use the SSME. Obviously the SSME's poor sea level performance and high integration costs overshadowed its advantages of light weight and compactness. At the bottom of a HLLV compactness does not count for much. It is astounding to me that this team can now suggest that the SSME is a good choice for a HLLV when 2 other teams rejected it.

Your comment about the Centaur upper stage is exactly what I said. In most cases RP7/LOX is better for lower stages however LH2/LOX is better for upper stages where the high Isp of hydrogen really shines. That is why von Braun let the hydrogen boys use the J2 in the 3rd stage of the Saturn V but kept the 1st two stages kerosene. One of the things in the report is that they want to get rid of the foam danger. The most obvious way to do this is to stop using liquid hydrogen. This is an example of what seems to be the ideal fuel with a high Isp leading to unforeseen problems later on. LH2 is a good fuel however it leads to large tanks that are difficult to keep cool. In the end if the original designers of the shuttle had chosen kerosene there would have been no foam problems and no disaster. It just reinforces that unforeseen problems can bite you long after the design is done.

I agree that boiler plate liquids instead of the SRBs would be better. However if you use better performing sea level engines then the SRBs could be smaller and safer.

A smaller crew return shuttle could even have a go-around capability. Because you are not trying to lift freight the weight of jet engines could be justified. I just think that the capsule is a bad idea. If Burt Rutan with 20 million dollars can fly a wing to 100km why can't NASA with it billions of dollars do the same. Is the capsule going to splash down in the ocean? What if the parachutes fail? What if it sinks on landing? A capsule is not a sure thing. Just now the shuttle could land in Edwards rather than Florida. Support ships for a recovery take time to move. Even if the capsule can land cross range can the recovery team get there in time. The capsule could have quite limited time to spare. If a large hurricane shuts down the recovery area where is it going to go? At least in an extreme emergency a winged lander could conceivably land at a commercial airport of which there are thousands all over the world.

The reason that the Apollo system needed a 300t capability to LEO was the method of landing choice - LOR and the race mentality of the moon landings. They had neither the time or expertise to assemble a transfer vehicle in orbit from lighter components.

There is no engineering reason why the components of the moon lander have to weigh 100t or why they all have to be all lifted at the same time. The space station was not lifted all at once. We are not in a race with the Soviets and can take our time. The Moon and/or Mars orbital transfer vehicles can stay in orbit and be re-fuelled and re-used if we take our time and do things from a perspective of a long term future in space. This is what Apollo did not do. It was a magnificent achievement but it left no long term facilities or capability in space for future operations - it was a dead end. This is what von Braun feared and why he wanted the Earth Orbit rendezvous method chosen.

You also have to take into account that the 100t booster will have to be developed with no commercial use so it will have to be totally paid for by the government. Also it will probably have a low launch frequency the facilities that have to be built to launch it will not get used much. This is a lot of money sitting around doing nothing.

Using a smaller commercial launcher and being smarter about how we go about getting to the Moon and Mars will lower the cost and risk and ensure that we actually do it rather than spending billions of dollars and then it all disappearing into projects that never were feasible in the first place. You do not have to look any further than the X-33 to see where this project could go.

BTW I did RTFA and I do think about what I say.

You are not questioning a glaring assumption on the part of the people that are running this project. They blithely assume that re-using shuttle components will lower costs with no supporting evidence. It could well turn out that the shuttle components could save money however the chances are higher that the attempt to save money could lead to the exact opposite as unforeseen integration costs outweigh any savings from using existing hardware. The end result could be exactly the same as starting from scratch with proven lower performance hardware that is in commercial use.

Or why develop a totally new booster at all. From this site the Delta 4 Heavy can now lift 25 000 Kg to LEO and there are already plans to stretch this to 35t by adding SRBs. With a few more modification this could even be streched to as much as 70 000 Kg. Lets use this to build a facility in space where Mars and Moon spacecraft modules can be checked out, assembled and re-fuelled. This could sustain a permanent expansion into space and the solar system rather that what seems to be a pork and politically motivated stunt. The astronaut lifting could be done by a commercial outfit supported by NASA rather than spending billions to re-invent the wheel. The spacecraft modules could be lifted by existing commercial launchers. This would leave billions of dollars being available for exploration instead of financing boosters that we do not need.

Rather than creating a new booster put the money into evolving the Delta 4 which has a large growth potential right up to 100 000kg to LEO. Put more money into the Spaceship 1 derivatives for the crew launcher and use the money saved to actually explore the planets rather than supplying pork.
 
Sorry the link did not come out. I got all spacecraft information from http://www.astronautix.com/index.html
 
Stephen, perhaps you'd like to dispute the conclusions of a real rocket scientist.

Namely, Robert Zubrin.

He doesn't agree with anything you've said.
 
Mr Engineer Poet I would ask you to do the same that you asked me to do and RTFA. The article that you referred to only mentions the uprated shuttle in one or two paragraphs. Now I agree that if you really want to re-use Shuttle components and that is your overiding objective over all the engineering problems then this configuration, called the Shuttle-C would be the easiest. Because it only replaces forward of the engines with fairing then most of the design parameters would be the same. Now if you read the article properly he says that:

"The fastest route to creating a HLV at this point is by reconfiguring the hardware of the Space Shuttle stack, deleting the Orbiter and replacing it with a fairing and an upper stage. A variety of such Shuttle derived HLVs are possible, with LEO delivery capabilities ranging from 70 to 130 tonnes, with the more capable versions costing more to develop."

Notice that he says that the 70 to 130 t version would cost more to develop. I agree with this and this is what I have been saying - changing the configuration will cost money.

Now for an engineer that sort of agrees with me. Homer Hickam, whom you will remember was a Rocket Boy in the film October Skies, is quoted in this article. In it he says that:

"More critical of the shuttle program was Homer Hickam, a retired NASA engineer who in a commentary published in The Wall Street Journal said the shuttle "is still not a reliable vehicle and never will be.

"You simply don't place a fragile bird at the base of a big, quaking nightmare of rocket engines and a massive, debris-shedding fuel tank and get anything but an engineering debacle," Hickam said before recommending the shuttle fleet to the junk heap.

"When your design stinks, Engineering 101 says admit your mistakes and go back to the drawing board," said Hickam, asserting that most of the engineers he knows at NASA "have wanted to retire the shuttle for a very long time and build a reliable spaceship worthy of our country." "

So it is my rocket engineer against yours I guess. This article written after the Columbia disaster echos some of my own concerns. I am not sure who wrote it so it is not really an authority however it is an interesting read.

Links
Shuttle C http://www.astronautix.com/lvs/shuttlec.htm

Homer Hickam article
http://www.spacedaily.com/news/nasa-05w.html

Is The Shuttle Fatally Flawed
http://www.spacedaily.com/news/oped-03l.html
 
Mister Gloor, you're not paying attention.

1.  The "more capable" variants are the ones with capacities much greater than 70 tons.  (A Shuttle is capable of landing at weights upwards of 99 tons, so a basic thrust structure and aeroshell weighing 20 tons should allow 79 tons of cargo.)  The SSME itself weighs 6990 pounds, and 9.5 tons should be plenty for a lightweight shell (what I can find about the Shuttle Orbiter thrust structure indicates that it weighs about 4500 pounds).

2.  Go back to your Hickam quote.  Read it again.  Notice that the Shuttle-derived vehicles, HLV and crew carrier both, get rid of:
    a.  The Orbiter with its fragile TPS.
    b.  People next to SRB's (the HLV carries no people, the crew capsule is on top of the people carrier).

"So it is my rocket engineer against yours I guess."

No, Hickam and Zubrin are in broad agreement; they both disagree with you.  The proposal is exactly the sort of "return to drawing board" that both of them have called for.
 
Mr Engineer Poet. Thank you for this interesting discussion. I am sure that you and the managers of this project are wrong however who am I? I could well be totally wrong and this new direction could be the best thing since sliced bread. Anyway as a final word from me I would just like you to consider a few points and see if they make sense. I really do not see your interpretation of Homer quote - I took it to junk the whole lot and start again. Not change and start again with the same components. Also the Shuttle-C can only lift 65 000Kg to LEO. It was extensively studied. Anything greater than this needs large configuration changes.

Firstly consider the X-33. This was a technology demonstrator designed to test some of the needed components for the Shuttle replacement the VentureStar. Chief amongst these components were flat, aerodynamic composite hydrogen tanks and aerospike engines. Now the XRS-2200 aerospike engine did work OK. It was needed to overcome the problems with the SSME sea level performance and provide maximum thrust from sea level to orbit, something no conventional bell nozzle rocket motor can do. The composite hydrogen tank was a failure. Despite assurances from learned people that a flat composite tank held flat by tension members formed from the substrate would be a piece of cake - it wasn't. This lead to 2 billion dollars being spent and the project cancelled leaving us in this position of an aging shuttle and no replacement.

Second consider the environment that you are dealing with - hypersonic speeds and high performance rocket motors. To get an idea of the difficulty of this environment you need go no further than the recent space walk to retrieve a small gap filler from the heat shield of the shuttle. To you or I, unused to such velocities, a small protuberance is no big deal. However to people that know this small thing at hypersonic speeds can lead to eddies that lead to local heating that lead to destruction. So the crew were commanded to remove this by people who know about such things. Also with the SSME the turbo pump is so highly loaded because of the high chamber pressure that small dents and eddies in the fuel lines can lead to fuel or oxidizer starvation which in milliseconds can lead to the destruction of the turbo pump and motor.

You are right that with due care this can be made to work. However due care is absolutely critical in this environment where are a protruding screw head can destroy the vehicle. This level of care does not come cheap. NOTHING in the Shuttle program so far, from conception to retirement, has been easy or cheap. EVERYTHING has been more expensive and taken longer that at first thought and mistakes have been disastrous. Please excuse for me for now being deeply suspicious when the same people then say "well we will just re-arrange a few things and put a few bits here and there and it will be easy and cheap" Nothing in the previous history of the shuttle program gives me any confidence that anything will change and this new direction will also take twice as long and cost twice as much.

Lastly consider that there is a new age dawning in manned space flight and 2 HLLVs are flying today, with much more suitable hardware, and can be found if we look hard enough. Since Spaceship One took 20 million dollars to go to 100 km everything has changed. We do not need huge government pork programs to send men and woman to orbit. It is very difficult to do and a lot of care needs to be taken however there are individuals and companies that can do this with little interference. Give 2 companies 200 million dollars, because it is at least 10 times more difficult to go to 300Km than 100km, and a promise of a contract in the future for the winner with a target of 4 astronauts safely in LEO and return in 3 years and let them go. These companies can then lift ANYONE to orbit for a fare as long as NASA is assured of a certain number of seats. Imagine what this would do for spaceflight.

The HLLVs already exist and are in production. They are the Atlas V and the Delta 4. Both programs, because they could not risk failure as this would have been financial ruin for the divisions of the companies building them, are very conservative designs with huge growth potential. There is no envelope pushing in these design and in fact they were criticized for being so conservative. However this means that with incremental and low risk changes they both can be developed over time to 70 to 100t to LEO launchers. Also as they are modular the launcher can be assembled to fit the payload. If there is only 50t capability needed you do not have to launch a half full 100t launcher. Finally because they are clean sheet designs they use far more suitable engines that perform well at sea-level and do not need huge SRBs to overcome this deficiency.

The main problem with the program of using shuttle hardware is that political needs are overriding good engineering. In trying to re-use shuttle components this is imposing artificial constraints on the design. Normally when this happens the result is disaster. Good engineering only occurs when the engineers are free to choose whatever components that fit within the budget are best suited to the job - not to be told what components they have to use and go from there. If an engineer had to choose a lower stage engine he/she would not choose an SSME as it is totally unsuited by design for this purpose. This is like using surplus F100 jet engines in the 787 commercial airliner.

My fear is that because this is not good engineering this program, like the X-33, will absorb billions of dollars while the engineers struggle with the problems they will encounter. This, again like the X-33, will lead to the eventual cancellation of the program with nothing achieved. I do not think that the American manned space program can survive 2 new directions being cancelled with no result and will founder, taking my dream of man on Mars with it.
 
I sort of agree with Stephen Gloor here. The SSME's should be avoided. And the design for the heavy lifter seems rather ambitious for the little we currently have planned. Also, I really don't trust Zubrin's judgement in this matter.

Having said that, I don't see a real problem with much of the designs. If they can find or make substantial launch volume for these rockets, then the costs will probably be reasonable.

However, it strikes me that Mr. Gloor is right in another area. We also have an opportunity to encourage private industry to enter in this range. All NASA has to do is promise enough business (IMHO).

At some point NASA will need to exit the launch industry. It's not their job and they'll be taking business away from US launch companies.
 
"The SSME's should be avoided."

Why?  They've never been associated with a mission failure; though a premature shutdown once did cause one mission to wind up in a lower orbit, it was a sensor malfunction and not an actual mechanical failure.  They're also the only hydrogen engines big enough to do the job, and they happen to be in production.

I suppose we could engineer something based on the aerospike engine of the VentureStar, which would probably have nice performance if it was expanded to cover the entire aft end of an ET with virtual nozzle.  But that would cost money to develop; we can buy SSME's off the line.

"And the design for the heavy lifter seems rather ambitious for the little we currently have planned."

That's another way of saying we shouldn't plan anything bigger, especially not Moon missions, Mars missions or e.g. new space stations lofted in a single shot.

The little lifter with the crew capsule would be much cheaper than a Shuttle.

But much as I'd love to see NASA get out of the launch business, I don't see that happening as long as Washington is footing the bill; the only way they're going to write the checks is if they can get pork for the folks at home, and the winner of a competitive bid is all too likely to be in the wrong district.  We'll see NASA get out when the business of space hotels or something like that gets going.
 
Engineer-poet - the SSME should be avoided not because they are unreliable but because they are just not suitable engines by design for 1st stages. Put them in an upper stage and they will perform magnificently as they are the highest Isp engines ever designed. But their high-expansion ratio nozzle makes their sea-level efficiency and thrust low.

Better off the shelf engines already exist such as the RS-68 if you want to stick to LH2/LO2. However if you accept that the RS-68 is a better choice of engine for a first stage then you might as well use the stage designed for it and tested, the Delta 4 Common Core.

I do not think that the HLLV is ambitious I just think we already have the launcher - we just do not realise this.
 
Engineer Poet - I just read one thing in your post I missed before. You said

" They're also the only hydrogen engines big enough to do the job, and they happen to be in production."

This is absolutely wrong and I would really like you to acknowledge this. I posted the performance figures of the RS-68, which is in production and used in the Delta 4 however I will do so again.

The SSME
Propellants: Lox/LH2 Thrust(vac): 232,301 kgf. Thrust(vac): 2,278.00 kN. Isp: 453 sec. Isp (sea level): 363 sec. Burn time: 480 sec. Area Ratio: 77.5.

RS-68
Propellants: Lox/LH2 Thrust(vac): 337,807 kgf. Thrust(vac): 3,312.00 kN. Isp: 420 sec. Isp (sea level): 365 sec. Mass Engine: 6,597 kg. Area Ratio: 21.5.

As you can plainly see the RS-68's vacuum thrust is 3,312 Kn whereas the SSME vacuum thrust is 2,278.00 kN. Clearly the RS-68 has 1/3 more thrust than the SSME which contradicts your statement. Also you can see the sea-level performance difference here:

RS-68
Sea Level Thrust - 2,891 KN
SSME
Sea Level Thrust - 1,668 KN

The RS-68 is over twice as powerful at sea level than the SSME backing up what I said about the SSME not being a suitable lower stage engine.

sources for the thrust figures
SSME
http://www.spaceline.org/rocketsum/main-engines.html
http://www.astronautix.com/engines/ssme.htm

RS-68
http://www.boeing.com/defense-space/space/propul/RS68.html
http://www.astronautix.com/engines/rs68.htm
 
E-P, IMHO the SSME just isn't suitable as an engine for several reasons.

The SSME haven't overtly been linked to a mission failure. However, I suspect they are a leading cause of mission cost and delays for two key reasons. First, they are extremely complex. That means increased costs of building one and of inspecting and refurbishing one after each mission. Second, they leak. I suspect hydrogen leaks are probably the leading cause of delay for a shuttle mission.

Well, here's some supporting evidence. Doesn't quite align with my suspicions, but agrees in principle.

"The SSME is the single largest contributor to Shuttle delays & escalating costs. The DDT&E cost accounted for 25 per cent of the total Shuttle development cost, and various technical problems pushed back the STS-1 launch date by two years. Nothing about the SSME is ordinary; it costs $40 million per copy and Rocketdyne needs four years to produce each new engine."

[something is wrong with the italics html tag]
 
Hmmm, the same author indicates that the SSME's have improved substantially in reliability in the last ten or so years. However, there's another issue here.

The role of the engine is being shifted from reusable to disposable. This is another good reason to avoid the SSME's. They're designed to be reusable. That drives up the cost. Given the complexity of the system, redesigning the engines for the reduced reliability needed for an expendable, may mean effectively a new engine.
 
Mr. Gloor:  You're right about the RS-68.  It came along after I quit paying close attention to the field, so you've got me there.

However, your cite for the RS-68 shows a relatively low T/W (51.2) and vacuum Isp (420 sec) compared to the SSME (73.1 and 453 sec).  The lower Isp is going to cut the mass fraction to orbit, and the greater weight is going to take a chunk out of the payload.

The RS-68's thrust is about 1.5x that of the SSME, not 2x.  A pair of RS-68's is a rough replacement for a trio of SSME's; you'd have less delta-V offset by slightly lower gravity losses.

As for the SSME's DDT&E, that is a sunk cost.  Scrapping it is not going to get any of that money back.

"SSME should be avoided ... because they are just not suitable engines by design for 1st stages."

The SSME cannot be started or re-started in flight.  The mission for which it is designed, and the only one for which it is suitable, is the sustainer of a parallel-staged booster.

" I just think we already have the launcher [Delta IV common core] - we just do not realise this."

The Delta-IV Heavy has a maximum payload to orbit of a mere 10,843 kg, less than a third of the typical Shuttle-derived booster.  This is not a vehicle for Mars missions.  Even the heaviest Atlas variant is only capable of 20 metric tons to LEO, far below the 35 tons of the minimal Shuttle-C.

Mr. Hallowell:  The Shuttle launch scrub data indicates that the SSME is now responsible for perhaps 20% of the launch scrubs, not the 1/3 it was in 1993 (one would expect engines of any type to require much more time to fix than e.g. the weather).  The RS-68 might be more reliable, and probably could be made more so, but an HLV clustering 2 or 3 RS-68's would present greater risk than the same using 3 SSME's (which have already flown over 100 missions with no vehicle or mission losses due to the engines).

Building a vehicle for manned lunar and Mars missions could be done with Delta V parts, but it would take paired RS-68's in combination with at least two SRB's to lift a Shuttle-C equivalent payload to LEO.  It looks like no such configuration has ever been tested.  The risk appears high compared to a slightly reconfigured ET feeding 3 SSME's with the same two SRB's, and while I have no objections in principle to working the cost down I would worry about both technical (proving a never-tested cluster configuration) and political (shifting lots of money from established interests, placing total support in jeopardy) risks.
 
E-P, you wrote:

"The Shuttle launch scrub data indicates that the SSME is now responsible for perhaps 20% of the launch scrubs, not the 1/3 it was in 1993 (one would expect engines of any type to require much more time to fix than e.g. the weather). The RS-68 might be more reliable, and probably could be made more so, but an HLV clustering 2 or 3 RS-68's would present greater risk than the same using 3 SSME's (which have already flown over 100 missions with no vehicle or mission losses due to the engines)."

I think it would be better to accept the higher risk here. the RS-68's can be with experience made more reliable just as the SSME's were. Hmmm, in this news story there is a short discussion comparing the SSME with the RS-68:

"Compared to the SSME, development time for the RS-68 was cut in half, the number of parts was reduced by 80 percent, the hand-touched labor reduced by 92 percent and non-recurring costs were cut by a factor of five.

"We've been able to get a lot of hand-work out of the RS-68 and replaced it with numerically-controlled machines. So instead of having a thrust chamber built up of a lot of tubes, we've machined this thing out of a solid piece of metal, which increases reliability," Collins said."

When you toss in the claimed decline in costs and simplification of the design and conbstruction, this rocket appears to be a signficant improvement over the SSME. And worth sacrificing some reliability over.

E-P, you wrote:

"The risk appears high compared to a slightly reconfigured ET feeding 3 SSME's with the same two SRB's, and while I have no objections in principle to working the cost down I would worry about both technical (proving a never-tested cluster configuration) and political (shifting lots of money from established interests, placing total support in jeopardy) risks."

The cluster hasn't been tested either way though the SSME are I allow less risky in the near future. Second, we have already determined (see the history of the ISS and the numerous experimental programs to study replacements for the shuttle) that the established interests are harmful to effective NASA operations. IMHO, going with the better designs is more likely to extend NASA's useful lifespan (such as it is) than continuing to feed the established interests.
 
E-P

Comparing the thrust of the RS-68 and RS-24(SSME) at vacuum is meaningless for this discussion. This proposal uses SSME as first stage motors. Of course the SSME is going to outperform the RS-68 in a vacuum as this is what is was specifically designed for. The relevant part, which you have overlooked, is the SEA-LEVEL performance. Quoted below is the entries from the Encyclopedia Astronautica You will see that the SEA-LEVEL Isp of the RS-68 and SSME is 365 and 363 respectively. At SEA-LEVEL the RS-68 is superior rendering your argument void. Also slightly misquoted are the SEA-LEVEL thrust figures that I posted.

RS-68 Sea Level Thrust - 2,891 KN
SSME Sea Level Thrust - 1,668 KN

The difference is closer to 2891/1668 = 1.73 not 1.5 which is a quite significant difference. True it is not over 2 times as I mistakenly quoted, as I cannot divide 2 numbers sometimes, however it is not the 1.5 that you said either. Also because of this the T/W ratio is not as good either.

SSME sea level T/W = 1668KN/3177KG = .52 KN/Kg
RS-68 sea level T/W = 2891KN/6597 = .45 KN/Kg

Finally the proper figures for the Delta 4 Heavy are posted below. You can quite clearly see that the lift capability of the Delta 4 Heavy is 25,800 kg to a 185 km orbit (LEO) which is similar to the Shuttle that can lift 24,400 kg to a 204 km orbit. You quoted the payload to GTO which is quite different.

The Delta 4 has an upgrade path to over 70t to LEO - true that extensive modifications are needed however these can be evolutionary rather than total redesign. The design for the HLLV looks more like a complete re-design as the ET was never designed for the loads at the bottom of the tank. Even if they used RS-68s in place of the SSME then this would be a bit better. I think that this is all I will say on this as I am not an engineer and I have reached the limit of my incompetence. I just hope that the time and effort required to make this proposal work with Space Shuttle components will not be just as hard and expensive as staring from a clean sheet with more suitable hardware.

Manufacturer Name: RS-68. Designer: Rocketdyne. Developed in: 1998. Application: . Propellants: Lox/LH2 Thrust(vac): 337,807 kgf. Thrust(vac): 3,312.00 kN. Isp: 420 sec. Isp (sea level): 365 sec. Mass Engine: 6,597 kg. Chambers: 1. Chamber Pressure: 95.92 bar. Area Ratio: 21.5. Thrust to Weight Ratio: 51.2. Country: USA. Status: In Production.

Manufacturer Name: RS-24. Designer: Rocketdyne. Developed in: 1972. Application: . Propellants: Lox/LH2 Thrust(vac): 232,301 kgf. Thrust(vac): 2,278.00 kN. Isp: 453 sec. Isp (sea level): 363 sec. Burn time: 480 sec. Mass Engine: 3,177 kg. Diameter: 1.63 m. Length: 4.24 m. Chambers: 1. Chamber Pressure: 204.08 bar. Area Ratio: 77.5. Oxidizer to Fuel Ratio: 6. Thrust to Weight Ratio: 73.1197829645898. Country: USA. Status: In Production. First Flight: 1981. Last Flight: 1998. Flown: 279. References: 225 . Comments: Used in Shuttle Orbiter. Space Shuttle Main Engine. Staged combustion, pump-fed. Originaly specification was vacuum specific impulse of 455, but not achieved in the final design.

Delta 4 Heavy
Manufacturer: Douglas. Launches: 1. Success Rate: 100.00% pct. First Launch Date: 21 December 2004. Last Launch Date: 21 December 2004. LEO Payload: 25,800 kg. to: 185 km Orbit. at: 28.5 degrees. Payload: 10,843 kg. to a: Geosynchronous transfer, 27deg inclination trajectory. Liftoff Thrust: 884,000 kgf. Liftoff Thrust: 8,670.00 kN. Total Mass: 733,400 kg. Core Diameter: 5.00 m. Total Length: 70.70 m. Span: 15.00 m. Development Cost $: 500.00 million. in 2002 average dollars. Launch Price $: 254.00 million. in 2004 price dollars. Cost comments: The originally estimated launch price in 1999 was $170 million. Due to the collapse of the commercial launch market, this was revised by the USAF in November 2004 to $ 254 million.

Shuttle
Manufacturer: NASA. Launches: 116. Failures: 1. Success Rate: 99.14% pct. First Launch Date: 12 April 1981. Last Launch Date: 16 January 2003. Launch data is: continuing. LEO Payload: 24,400 kg. to: 204 km Orbit. at: 28.5 degrees. Payload: 12,500 kg. to a: space station orbit, 407 km, 51.6 deg inclination trajectory. Apogee: 600 km.

Delta 4 upgrade to 35t to LEO
Orbital launch vehicle. Family: Delta. Country: USA. Status: Study 2004.
Upgrade to Delta IV Heavy by adding 4 GEM-60 solid rocket boosters. 6.5 m diameter payload fairing. Introduction would require modifications to existing launch pads.
Manufacturer: Douglas. LEO Payload: 27,000 kg. to: 407 km Orbit. at: 28.5 degrees. Payload: 10,000 kg. to a: earth escape trajectory. Core Diameter: 5.00 m. Total Length: 71.00 m. Span: 15.00 m.

to

Delta 4 Heavy upgrade to 85t
Orbital launch vehicle. Family: Delta. Country: USA. Status: Study 2004.
Upgrade to Delta IV Heavy by clustering seven common booster modules, using a new RS-800K engine in the booster stages, an AUS-60 upper stage powered by 4 MB-60 or RL-60 27 tonne thrust Lox/LH2 engines, and aluminium-lithium lightweight alloy in place of the existing aluminium in all stages. Payload fairings over 6.5 m diameter could be accomodated. Introduction would require new launch pads and booster assembly infrastructure.
Manufacturer: Douglas. LEO Payload: 85,000 kg. to: 407 km Orbit. at: 28.5 degrees. Payload: 32,000 kg. to a: earth escape trajectory. Core Diameter: 5.00 m. Total Length: 67.00 m. Span: 15.00 m.


References
http://www.astronautix.com/lvs/shuttle.htm
http://www.astronautix.com/engines/rs68.htm
http://www.astronautix.com/lvs/delheavy.htm
http://www.astronautix.com/lvs/shuttle.htm
 
That's an impressive list of Delta-variant vehicles, but I have to note several things:

1.  You're flat wrong about the Shuttle cargo variants using SSME's as first-stage motors.  They use SRB's; the SSME's are burning from the pad to orbit, but the SRB's provide a majority of the total vehicle thrust for the first 2 minutes.

2.  None of the Delta IV configurations which have Shuttle-C class capability have ever been tested.

3.  All of the Delta vehicles require an upper stage.  Shuttle and its variants are 1.5 stages to orbit; all engines are burning as the vehicle leaves the pad, so there is no way for an ignition failure to result in failure to make orbit.  This is not true of the Deltas.

4.  While the basic Shuttle-C would have a capacity of 70 tons and upgrades could bring it to nearly twice that, the Delta IV Heavy can only orbit 28 tons and the top of the upgrade path (which requires 4 solid boosters) is only 85 tons.  This is insufficient for manned missions to the Moon or Mars.

5.  The Delta IV Heavy upgrade has twice the solid-motor risks of Shuttle-C.

There's one thing you don't appear to appreciate about specific impulse.  LEO orbital velocity is roughly 7700 m/sec, and typical solid boosters drop off somewhere around 1 km/sec; this means the liquid booster engines need to impart 6700 m/sec of delta-V.  (The vehicle is effectively in vacuum at that time, so its sea-level performance is irrelevant from separation of the solids to engine cutoff.)  If you have two engine options, of 453 seconds and 420 seconds Isp, the amount of fuel which gives you 6700 m/sec of delta-V in the first will only yield 6212 m/sec from the second.  That's half a kilometer per second you have to make up with more fuel, which could otherwise have been payload.  Extra engine mass comes straight out of payload.

Half a kilometer per second is a huge number.  IIRC, 1 m/sec of delta-V near LEO changes the altitude at the other side of your orbit by roughly a mile.  Losing 488 m/sec at MECO would put your perigee at almost 300 miles below sea level.

It might be possible to adapt the RS-68 to achieve SSME-class impulse using an aerospike nozzle; such an advance would eliminate the need for an upper stage to achieve orbit, and improve both the vehicle reliability and payload.  But this is a long-term project, and getting e.g. Mars Direct launched soon can only be done with something like Shuttle-C.
 
1. True enough however then the Isp situation is worse because the SRB Isp is only 240 or so.

2. True bit neither has the Shuttle-C

3. Very true and parallel staging can be more reliable - the Russians use it extensively

4. What upgrades? No-one knows how heavy a manned mars mission will be so it is premature to say this.

5. Solid motor risk is more related to size than number as there is more risk of manufacturing defects as the size increases. This is why the SRBs were a big departure.

Isp as such is a bit of a fools gold chase. Chasing 430 Isp cost the shuttle program 2 years and a lot of crossed fingers with the SSME's.

I am not sure that you have the 1 Km/sec bit correct as the SRBs burn for 2 minutes. This gives an average acceleration of v=.5 a * t^2
so 1000 m/s = .5 * a * (120 sec ^2)
a = 0.13 m/s^2
This seems a bit low as the main engines only continue for another 8 mins. I am sure the acceleration should be over 9.8 m/s^2 for a start.

If the average acceleration until T+120sec is say 20m/s^2 then the velocity at SRB seperation is 144 km/sec. Even then I think this is a bit low.
 
Got the last bit completely wrong - please disregard.

The 1km per second is about right.

The lower Isp of the RS-68 just means that a higher mass fraction is needed to achieve orbital velocity.

The rocket equation tells why

Vfinal = C * ln * (Initial Mass / Final Mass)

The higher Isp of the SSME means that is has a higher exhaust velocity (C) so therefore will give a higher final velocity for a given mass ratio. The problem that the shuttle designers had was that the shuttle itself is very heavy and does not have a good mass ratio so therefore it needed a very high Isp engine to get to to orbit.

An ummanned HLLV has no such restraints and can be made much lighter. Obviously the RS-68 works as the Delta 4 achieves orbit in combination with the high Isp RL-10.
 
The original Mars Direct program called for 45 tons landed on Mars.  The Hohmann transfer orbit data I've got calls for 2.411 km/sec minimum delta-V required to go from Earth's orbit to MTO; if you're starting from a circular parking orbit at 7700 m/sec the total delta-V is about 3.2 km/sec.  Applying this kick with an H2/O2 upper stage using RL-10's at 450 sec impulse requires a mass ratio of about 2.07; with the transfer stage tankage, plane-change requirements and whatnot, call it an even 100 tons to orbit.

You can put 100 tons into LEO with 3 SSME's, 2 SRB's and a standard ET.  You don't have to modify anything about the physical configuration of the engines or their fuel plumbing, though you could hard-couple them to the ET and save the weight and expense of the disconnects.  Doing another upgrade of the Delta IV heavy is a stretch upon a stretch; perhaps worth doing, but not the slam-dunk of the Shuttle-C.
 
Fair enough - I hope it all works
 
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