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http://www.astronautix.com/props/loxlh2.htm

Propellant Formulation: LOX/LH2.
Optimum Oxidiser to Fuel Ratio: 6.
Temperature of Combustion: 2,985  K.
Ratio of Specific Heats: 1.26.
Density: 0.28 g/cc.
Characteristic velocity c: 2,435 m/s.
Isp Shifting: 391.
Isp Frozen: 388.
Pp Isp Shifting: 109.
Mol: 10.00 M.
Isp: 391.00 sl.
Isp: 451.00 vac.

Oxidiser: LOX.
Oxidiser Density: 1.14 g/cc.
Oxidiser Freezing Point: -219.0 °C.
Oxidiser Boiling Point: -183.0 °C.

Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. Liquid oxygen, as normally supplied, is of 99.5 percent purity and is covered in the United States by Military Specification MIL-P-25508. High purity liquid oxygen has a light blue colour and is transparent. It has no characteristic odour. Liquid oxygen does not burn, but will support combustion vigorously. The liquid is stable; however, mixtures of fuel and liquid oxygen are shock-sensitive. Gaseous oxygen can form mixtures with fuel vapours that can be exploded by static electricity, electric spark, or flame. Liquid oxygen is obtained from air by fractional distillation. The 1959 United. States production of high-purity oxygen was estimated at nearly 2 million tonnes. The cost of liquid oxygen, at that time, ex-works, was $ 0.04 per kg. By the 1980's NASA was paying $ 0.08 per kg.

Fuel: LH2.
Fuel Density: 0.071 g/cc.
Fuel Freezing Point: -259.0 °C.
Fuel Boiling Point: -253.0 °C.

Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks. The United States mastered hydrogen technology for the highly classified Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950's. The technology was transferred to the Centaur rocket stage program, and by the mid-1960's the United States was flying the Centaur and Saturn upper stages using the fuel. It was adopted for the core of the space shuttle, and Centaur stages still fly today.

In Russia hydrogen fuelled upper stages were designed and developed by the mid-1970's, but the Russians never seem to have found the extra performance to be worth the extra cost. Europe and China developed liquid oxygen/liquid hydrogen engines for upper stages of the Ariane and Long March launch vehicles.

The equilibrium composition of liquid hydrogen is 99.79 per cent parahydrogen and 0.21 per cent orthohydrogen. The boiling point of this composition is -253°C. Liquid hydrogen is transparent and without a characteristic odour. Gaseous hydrogen is colourless. Hydrogen is not toxic but is an extremely flammable material. The flammable limits of gaseous hydrogen in air are 4.0 to 75 volume percent.

Hydrogen is produced from by-product hydrogen from petroleum refining and the partial oxidation of fuel oil. The gaseous hydrogen is purified to 99.999+ per cent, and then liquefied in the presence of paramagnetic metallic oxides. The metallic oxides catalyse the ortho-para transformation of freshly liquefied hydrogen. Freshly liquefied hydrogen which has not been catalysed consists of a 3:1 ortho-para mixture and cannot be stored for any length of time because of the exothermic heat of conversion. The delivered cost of liquid hydrogen in 1960 was approximately $ 2.60 per kg. Large-scale production was expected to reduce the cost to $ 1.00 per kg. In the 1980's NASA was actually paying $ 3.60 per kg.

Page contains a table of LOX/LH2 Rocket Motors

I assume that when they talk about the optimum propellant ratio, they are talking about volumes-6 units of LH2 to 1 unit of LOX. I believe this means that not all of the LH2 is burned as it leaves the engine.

I remember that the SII second stage of the old Saturn Vs was built with only a single wall separating the volumes of LH2 and LOX, in order to save weight. For whatever reason, the engineering pressure to reduce weight seemed to fall mostly on the design of the second stage.

I assume that when they talk about the optimum propellant ratio, they are talking about volumes-6 units of LH2 to 1 unit of LOX. I believe this means that not all of the LH2 is burned as it leaves the engine.

Correct.  The density of LOX is 1.14 g/cc, while that of LH₂ is 0.07 g/cc.  The idea is to run the mixture slightly enriched in H₂ in order to reduce the effective molecular mass.  Isp is a function of the molecular mass, and so one can achieve a greater Isp with H₂ rather than H₂0.  So some portion of the H₂ absorbs the energy from the combustion (oxidation) reaction, but it is not combusted.

I remember that the SII second stage of the old Saturn Vs was built with only a single wall separating the volumes of LH2 and LOX, in order to save weight. For whatever reason, the engineering pressure to reduce weight seemed to fall mostly on the design of the second stage.

The first stage of Saturn used 5 F-1 engines.  The single-chamber F-1 used liquid oxygen (lox) and RP-1, a kerosene, and was the largest and most powerful single liquid-fuel rocket engine ever built.  ( http://www.boeing.com/history/bna/f1engine.htm )

The second and third stages had to be as small as possible for the payload (Apollo Command/Power Modules and Lunar Excursion Vehicle).  The LEM structure was also rather minimal.  The Apollo Command vehicle was somewhat heftier since it had to return through the atmosphere.

The history was complicated. The third stage of the Saturn V was the S-IVB, which was a pre-existing rocket stage originally built for the Saturn IB. Maybe they didn't want to tinker too much with a proven vehicle. Also for whatever reason, there didn't seem to be as much structural weight savings tried on the S-IC. In any case that phase of the boost experienced the highest acceleration during the ascent, 4Gs or so at SI-C burnout and maybe the engineers didn't want to take too many chances(?) Just guessing.

The S-II turned out to be far the most troublesome of the three rocket stages to engineer.

It's interesting that LOX, being so much heavier than LH2, is what limits the achievable specific impulse. In the old days they talked about the potential of mixing fluorine with the oxygen (perish the thought-HF is horrible stuff!) If LH2 could be burned with oxygen from the air during the initial boost phase-some kind of hybrid jet-rocket engine, a lot of propellant mass could be saved.

A very good (though rather expensive) book that covers both the political and engineering backgrounds of boosters:
TO REACH THE HIGH FRONTIER
A History of U.S. Launch Vehicles
Lannius & Jenkins, Editors
2002 University Press of Kentucky
ISBN 0-8131-2245-7

The history was complicated. The third stage of the Saturn V was the S-IVB, which was a pre-existing rocket stage originally built for the Saturn IB. Maybe they didn't want to tinker too much with a proven vehicle. Also for whatever reason, there didn't seem to be as much structural weight savings tried on the S-IC. In any case that phase of the boost experienced the highest acceleration during the ascent, 4Gs or so at SI-C burnout and maybe the engineers didn't want to take too many chances(?) Just guessing.

I'd say that's a reasonable guess.  Thanks btw to Spacecat for the book reference.

It's interesting that LOX, being so much heavier than LH2, is what limits the achievable specific impulse. In the old days they talked about the potential of mixing fluorine with the oxygen (perish the thought-HF is horrible stuff!) If LH2 could be burned with oxygen from the air during the initial boost phase-some kind of hybrid jet-rocket engine, a lot of propellant mass could be saved.

I remember a comparison of various fuel/oxidant combination and HF had the highest energy generation and therefore highest Isp.  However, having HF in the atmosphere is highly undesirable, as it nasty stuff and does nasty things to living tissue, not too mention the corrosion of metals.

Might have been OK for the 3rd stage, but then we'd have an ozone issue.

Photograph taken from the Apollo 8 spacecraft looking back at the Saturn V thir (S-IVB) stage from which the spacecraft had just separated following translunar injection. Attached to the S-IVB is the Lunar Module Test Article (LTA) which simulated the mass of a Lunar Module on the Apollo 8 lunar orbit mission. Sunlight reflected from small particles shows the "firefly" phenomenon which was reported during first earth orbital flight of Mercury program.  - Photo date 21 December 1968.


The Lunar Module (LM) 3 "Spider", still attached to the Saturn V third (S-IVB) stage, is photographed from the Command/Service Module (CSM) "Gumdrop" on the first day of the Apollo 9 earth-orbital mission. This picture was taken following CSM/LM-S-IVB separation, and prior to LM extraction from the S-IVB. The Spacecraft Lunar Module Adapter (SLA) panels have already been jettisoned.  - Photo date 03 March 1969


Nice overview of Apollo program at - http://www.apolloexplorer.co.uk/bymission.htm

Just to show how complicated the Saturn V was:

The S-IC first stage was built by Boeing and measured 138 feet tall by 33 feet wide with a 63-foot finspan. The powerful first stage employed five Rocketdyne F-1 engines which burned liquid oxygen/RP-1 (kerosene) liquid fuel and produced a combined 7,500,000 pounds of thrust at liftoff.

The S-II second stage was built by North American and measured 81 feet, 6 inches tall by 33 feet wide. It employed five Rocketdyne J-2 engines which burned liquid oxygen/liquid hydrogen and could produce a combined thrust of 1,000,000 pounds.

The S-IVB third stage, also used as a second stage on the Saturn IB, was manufactured by Douglas Aircraft. It measured 58 feet, 8 inches tall by 21 feet, 8 inches wide. The S-IVB employed one Rocketdyne J-2 engine which could produce a thrust of 200,000 pounds.

A NASA-designed and built Instrument Unit (IU) attached to the top of the S-IVB by special adapter measured 3 feet tall by 21 feet, 8 inches wide. The IU housed equipment which controlled all electronic commands for Saturn V control and guidance during ascent.

In a typical Saturn V Apollo flight profile, the five F-1 first stage engines were ignited six seconds before liftoff. The center F-1 engine was shut down 135 seconds after launch. The outer four F-1 engines were shut down 15 seconds later.

One second following cutoff of the four outer F-1 engines, the first stage separated. Simultaneously, eight retro-rockets were fired for less than one second to slow the forward speed of the first stage, thus keeping it from bumping into the second stage.

These first stage retro-rockets were located in pairs at the base of each of the outer four F-1 engines. They provided a total thrust of 88,500 pounds. Following separation, the spent first stage fell back into the Atlantic Ocean about 400 miles downrange.

One second after first stage separation, eight solid-fueled motors mounted on the first/second stage adapter ring were fired for four seconds. These provided a combined thrust of 181,000 pounds.

In addition to maintaining the positive motion of the rocket, these motors performed an ullage maneuver, forcing the second stage fuel to the bottom of its tanks in order to feed the engines. The five J-2 second stage engines were fired during this ullage burn.

Thirty seconds following second stage ignition, the first/second stage adapter ring separated and slid past the second stage engines for a tumble back toward Earth. Six seconds following this, the Apollo spacecraft escape tower was jettisoned.

The second stage engines burned for 365 seconds prior to separation from the third stage. At separation, four solid-fueled second stage retro-rockets were fired to keep the second and third stages from hitting one another.

These four retro-rockets were located in a conical adapter on the front face of the second stage. They provided a total thrust of 140,000 pounds. The second stage began its tumble, eventually impacting the Atlantic Ocean about 2,500 miles downrange.

At this point, the Saturn V had achieved a speed of 15,700 m.p.h. and an altitude of 115 miles.

Two solid-fueled ullage motors located 180 degrees apart on the third stage aft skirt were fired for four seconds to settle the liquid fuel. These motors produced 6,800 pounds of thrust.

Three seconds after second stage separation, the S-IVB third stage J-2 engine was ignited. Nine seconds later, the third stage ullage motors which fired at separation and their cases were jettisoned.

The third stage J-2 engine was fired for 142 seconds before being shut down. This initial S-IVB burn was sufficient to carry the Apollo spacecraft into a 118-mile orbit at a speed of 17,500 m.p.h.

from - http://www.spaceline.org/rocketsum/saturn-V-apollo.html

Now, I was wondering about the altitude of burnout for the second stage and more critically the ignition of the third stage.  It appears from this article that the second stage burns out at 115 miles, and the third stage kicks in at about 117-118 miles and its primary function is the get the third stage, with Apollo modules up to escape velocity (17,500 mph).  So the question would be, would an HF rocket be reasonable?  Otherwise, the incremental boost in performance and reduction in mass just do not outweigh the risk of having an H₂/F₂ system.

Also from - http://www.spaceline.org/rocketsum/saturn-V-apollo.html

At the end of this first S-IVB burn, two ullage motors were fired to settle the remaining fuel and provide spacecraft stabilization. These ullage motors were housed in two Auxiliary Propulsion System (APS) modules located 80 degrees apart on the third stage aft skirt.

Each APS module housed three attitude control motors and one ullage motor. The attitude control motors could each produce 150 pounds of thrust, while the ullage motors could each produce 70 pounds of thrust. All burned nitrogen tetroxide/hydrazine liquid fuel.

During two or three checkout orbits, the S-IVB attitude control motors could be fired in sequence to make any necessary on-orbit corrections. Following these checkout orbits, the ullage motors were fired for 77 seconds to settle the fuel and provide forward spacecraft momentum.

The third stage J-2 engine was then re-ignited for 345 seconds to achieve a speed of 25,000 m.p.h.. This second J-2 firing was necessary to carry the Apollo spacecraft out of Earth orbit and place it on a proper trajectory toward the Moon.

Once the Apollo spacecraft was on its way to the Moon, the Saturn V had completed its job. The S-IVB third stage separated from the Apollo CSM/LM combination. The third stage ullage motors were fired for 280 seconds to move the S-IVB away from the CSM/LM.

Now here is something interesting -

The third stage J-2 engine was then fired for the last time until its remaining fuel was spent. Depending upon the specific Apollo mission profile, the S-IVB was either sent toward deep space or the Moon.

So those Saturn IVB third stage boosters are still out there!  Somewhere.

-------------------------
Nice summary of rockets

http://www.spaceline.org/rocketsum.html
------------------------

Hi Astronuc, sorry I've been absent from the discussion for a bit-too much pressure with work, travel, and so on.

A couple of things. Different things were done with the burned out S-IVBs. There was a lot of publicity recently about how this object orbiting the Sun inside the Earth's orbit was probably the burnt out S-IVB from the Apollo 12 mission. On the other hand, the stage from Apollo 13 was deliberately crashed into the Moon's surface as a seismic test. I wonder what happened to the stage from Apollo 9. That was just an Earth orbit mission so when the stage was fired the second time it had no spacecraft to push, and must have picked up well more than escape velocity.

I haven't heard any discussion about fluorine additions to LOX in recent times. I wonder if there is a corrosion issue. Gaseous fluorine, of course, is pretty nasty stuff.

I remember reading that the F-1 engines were originally rated at 1.5MM lbs of thrust each but I also believe that later in the Apollo program they were uprated by several percent. It seems that increasing the burn rate, even though it shortens the burn time, can increase the payload. There was even talk of being able to get 1.8MM lbs out of each F-1 eventually, but the Saturn program was terminated in 1972.

Hey Retrospector, glad you could drop in.  I have the same pressure on lots of fronts.

I did not really follow Apollo back then, so it was news to me about the disposition of the upper stages.  I remember the seismic test though.

As far as I know, the only discussion of H/F systems was from the standpoint of thermodynamic and Isp.  No one would seriously consider an HF plume in the atmosphere - at least I hope not.  But out in space, that would be different.  On the other hand, F₂ is nasty stuff and leak or spill would be a big problem.

I seem to remember the first stage (5 F-1's) at 7.5 MM lbs thrust or 1.5 MM lbs from each F-1.

Back again, it's been a crazy month. I wonder if we lost some posts from this thread?

It still amazes me that the Centaur upper stage is still in use, in a lot of use as a matter of fact. Of course it has been considerably improved since the early 1960s. The much bigger LH2/LOX stages for the Saturns have long since come and gone.

The F-1s used on the "J" mission Saturns were uprated to slightly over 1.55 million pounds of thrust each by simply re-orificing them for higher mass flow.  No major pump work was required.  Liftoff thrust on the Apollo 15 Saturn V was officially rated at 7.765,852 lbs, increasing to 9,155,147lbs just prior to center engine cutoff.

There were later model F-1s tested on the stands at Edwards at 1.8 milliion pounds of thrust, and they apparently demonstrated stable combustion and adequate chamber cooling at this thrust level.  I don't have any specific impulse numbers - could've gone either way.  If they went up on chamber pressure and fiddled a bit with mixture ratios, they might've eked out a little more.  If they were close to critical thermally, they might have actually had slightly less Isp if they had to burn a little more fuel-rich to keep temperatures under control.  There was some talk of a stretched-tank S1C variant to take advantage of the available thrust, but of course it came to nought.

I saw a late test in early 1972 at Edwards.  I have no idea if it was a 1.55 or 1.8 million lb thrust engine, but I CAN tell you that a person 3/8 mile away from a single F-1 felt the start overpressure in their chest cavity!     During mainstage operation, we stood there glorying in the resonances in our chests, and after shutdown we realized that the skin on our upper bodies that had been under fabric was all slightly numb from the beating of our shirts on our skin while the engine was firing.  I can't imagine what all five sounded like - recordings simply can't duplicate the sheer presence that those overpressures in the air create!

  I was excited but didn't really realize what a rare privilege this was at the time - I was a 16 year old high school student.  I simply could not imagine that we had spent all the treasure and blood to develop this fantastic capability and would piss it away.  If somebody had told me that we were just a little more than a year away from the last flight of an F-1 powered vehicle, I would've laughed.   

Speaking of HF, we used about 35% HF as an immersion etching solution for titanium turbopump components for SSME.  Nasty, nasty stuff!  When you pulled a titanium fuel inlet assembly, a big "snail shell" volute, out of the solution, there was this thick, greasy-looking vapor coming off the surfaces, which we cheerfully referred to on my crew as "the brown cloud of death".  We all knew that contact with hydrofluoric acid was truly bad news -  either by inhalation or skin contact.  The stuff has a fairly strong affinity for calcium and eats right through soft tissues very quickly, so it is capable of doing real damage to bone.  Inhale it, if memory serves, in significant quantities, and you're going to be drowning in the reaction products filling your lungs.  If you work around strong acids and caustics, there are some which you come to regard as simple irritants.  Others, like hydrofluoric acid, must obviously be treated with the utmost respect.  HF is one of those uncommon compounds that you can't handle in normal pyrex labware - it merrily eats glass.  I once saw a foolish person pull some HF solution out of a tank with a 200 ml pyrex beaker.  He was in apron, gloves and respirator, but it was only dumb luck that saved him when the bottom fell out of the beaker 20 feet from the tank.  He didn't get splashed on his vulnerable lower legs or upper arms. and was still over the grated walking surface of our "pit" area, so the spill was caught for neutralization.  Nobody realized it wasn't one of the plastic beakers suitable for HF until it was too late.

And I thought it was fun working in there!

Jim

That shirt flapping sensation I recall from the Skylab launch. The sound was not so much a roar as much as an incredibly loud series of high-speed snare drum rimshots.

I need to look it up, but I seem to recall one flight had a premature F-1 shutdown, which was compenstaed for with a longer burn on the other engines. Lemme check.

yale

There was at least one premature J-2 shutdown (Apollo 13) on a crewed flight.  SA-502, the second test flight in April of 68, had multiple shutdowns of J-2s in the second stage - they had liquid air forming in some critical line sleeves and hydraulically creating ruptures.

I don't recall any premature F-1 shutdowns on crewed missions.

(Added more to the previous post, fond memories of working with HF..)

Jim

That s what I googled. J-2 pogo problems.

yale

The RS-68 engine is the first new large liquid-fueled rocket engine to be developed in the United States in 25 years. Designed for the Boeing Delta IV (http://www.boeing.com/defense-space/space/delta/delta4/delta4.htm) family of evolved expendable launch vehicles (EELV), the bell-nozzle RS-68 is a liquid hydrogen - liquid oxygen booster engine utilizing a simplified design philosophy resulting in a drastic reduction in parts compared to current cryogenic engines. This design approach results in lower development and production costs.

http://www.boeing.com/defense-space/space/propul/RS68.html

See also

http://www.boeing.com/defense-space/space/delta/delta4/journey/rs68engine.htm

Well, we are still using them!

ESA's page on the Vulcain/Vulcain-2.

http://www.esa.int/SPECIALS/Launchers_Access_to_Space/ASELVQI4HNC_0.html

Top photo - Test firings of the Ariane-5 cryogenic Vulcain engine in Vernon (France) and Lampoldshausen (Germany).

Credits: SNECMA/SEP/DLR/DASA

Bottom photo - DLR ( auf Deutsch - http://www.dlr.de/ )
( in English - http://www.dlr.de/en/desktopdefault.aspx )

RELEASE: 06-226 - http://www.nasa.gov/home/hqnews/2006/may/HQ_06226_RS-68_ENGINE.html

NASA has chosen the RS-68 engine to power the core stage of the agency's heavy lift cargo launch vehicle intended to carry large payloads to the moon.

The announcement supersedes NASA's initial decision to use a derivative of the space shuttle main engine as the core stage engine for the heavy lift launch vehicle.

The cargo launch vehicle will serve as NASA's primary vessel for safe, reliable delivery of resources to space. It will carry large-scale hardware and materials for establishing a permanent moon base, as well as food, fresh water and other staples needed to extend a human presence beyond Earth orbit.

Recent studies examining life-cycle cost showed the RS-68 is best suited for NASA's heavy-lift cargo requirements. The decision to change the core stage engine required an increase in the size of the core propulsion stage tank, from a 27.5-foot diameter tank to 33-foot diameter tank, to provide additional propellant required by the five RS-68 engines.

The RS-68 is the most powerful liquid oxygen/liquid hydrogen booster in existence, capable of producing 650,000 pounds of thrust at sea level. In contrast, the space shuttle main engine is capable of producing 420,000 pounds of thrust at sea level. The RS-68, upgraded to meet NASA's requirements, will cost roughly $20 million per engine, a dramatic cost savings over the shuttle main engine.

The prime contractor for the RS-68 engine is Pratt & Whitney Rocketdyne of Canoga Park, Calif. Pratt & Whitney Rocketdyne is the same company that manufactures the shuttle main engine.

The RS-68 is used in the Delta IV launcher, the largest of the Delta rocket family developed in the 1990s by the U.S. Air Force for its evolved expendable launch vehicle program and commercial launch applications.

The cargo launch vehicle effort includes multiple project element teams at NASA centers and contract organizations around the nation and is led by the Exploration Launch Office at NASA's Marshall Space Flight Center in Huntsville, Ala.

http://www.nasa.gov/mission_pages/exploration/main/index.html

I got a personal kick out of the RS-68 decision for the Cargo vehicle, or , apparently, the Ares V.  I wish they would quit trying to borrow Apollo nomenclature - it seems a bit self-conscious.  Crew launch Vehicle is Ares I, cargo Ares V, and the official release on the names called it a tribute, I think, to the Saturn IB and Saturn V.

I personally think Von Braun would be spinning in his grave if he knew that somebody was comparing that SRB pencil to his Saturn IB.  Even though the Saturn I family was an expedient that was built around clustered Jupiter and Redstone tankage as a quick and dirty way to gain heavy lift capability, I always found the unusual first stage contours rather elegant...

Anyway, back to the RS-68 decision.   I am fairly familiar with both the RS-68 and the SSME from manufacturing experience on both and long personal interest led me to become fairly familiar with operational characteristics. 

When they first announced the intention to power the upper stage of the crew vehicle with SSME, I recall wondering about implementing an air-lit SSME variant, given the rather extensive prestart conditioning requirements of the current engine - there's a great deal of ground equipment whose functions would have to be served by a flight-mass system, a real challenge if you want to preserve your Isp advantage for payload rather than trading it for subsystem mass.  Well, folks, it turns out that J-2X is looking like a better idea now, but it took a very long time to get to that decision - TOO long IMHO.

I was also frankly  incredulous about the idea of trying to make an lower-cost expendable SSME variant for the cargo vehicle, and I think I posted a speculation that they HAD to be talking about a recoverable engine module - there is just no way to build a truly cheap SSME variant.  I believe I opined at the time that RS-68 would make more sense for the cargo vehicle from a life-cycle cost AND operational flexibility standpoint. 

Now, I'm fairly bright, but certainly no engineering genius - yet my antennae were twitching on these issues over a year ago if memory serves.  It disturbs me that NASA took so long to reach these conclusions.

I fear that the vast inertia of the shuttle program once more reared its head in this process.  SSMEs are expensive, complex, very high performance machines, and they require all sorts of TLC to turn around between flights.  This employs a large number of people at KSC, Marshall, and Stennis (though not many at Rocketdyne these days - no more overhauls being done there...), and this interests the congressional representatives of these places.  An entire generaton of NASA program people have made their whole careers on shuttle, ET, SRB, and SSME.    They have not had to actually develop new major booster systems designs, just refine the one system they operate.  That's a mite unsettling to me if I think about it too long!  The idea that "shuttle-derived" equated to lower cost vehicles was sheer fantasy when one looked at STS program costs - yet the NASA folks clung to this straw as long as possible.  They wouldn't jump to RS-68 because the lower Isp meant either larger tanks or less payload than required for the CaLV baseline lunar missions.   Larger diameter tanks meant shuttle ET tooling couldn't be adapted at Michoud, and this was resisted until the absurd cost of tossing 5 SSMEs a launch forced them to reconsider - and perhaps recall that Michoud produced 10 meter diameter S-IC tankage 40 years ago and is quite roomy enough to do so again!

Anyway, I hope they move out now that the last vestige of drop-in shuttle hardware is off the table.  Even the 5 segment SRB is a new piece, though it is derivative.  They tried to hang on to the 4 segment RSRM for the crew vehicle, but they couldn't make that happen without a lightweight, air-lit SSME for the second stage. I'm hoping for good things, but I'm a concerned observer - we've already wasted more than a year squirming around with shuttle-legacy hardware designs.

Jim