Viability of NASA SLS launch system; should it be cancelled?

SpaceNut · 2017-12-08 21:13:50

More slow building process to detail what not to do nah...

Protecting SLS from Fire and Ice – TPS foam application proceeding at MAF

Yah the foam is the same stuff used on the shuttle stack....

Nice triple image:

RobertDyck · 2017-12-08 22:04:13

SpaceNut wrote:

Yah the foam is the same stuff used on the shuttle stack....

You're surprised? The core stage is a Shuttle ET, with the pointy LOX tank made flat with a dome tank top and interstage to support an upper stage. It has 4 SSME mounted on the bottom. They destroyed the friction stir assembly used to assembly Shuttle ETs, so had to build a new one. Contractors are charging copious quantities of cash on the excuse it's a new stage, but it really isn't.

SLS block 2 is just the Ares launch vehicle from Mars Direct, but with main engines moved axially instead of a side-mounted pod. That's all. Block 1, 1B, and 2B will have 4 engines, but the design for block 2 used 5 engines. Ares was a launch vehicle as powerful as Saturn V built from Shuttle parts. Ares upper stage was to use a single J-2S engine, which was just the newest version of the J-2, the same engine as the 3rd stage of Saturn V. SLS block 2 was to use a single J-2X engine, simply the latest 21st century version of J-2.

However, good news is there are no orbiters heat shield tiles to knock off. The foam is safe without those tiles. And the tiles would have been safe without the foam. The foam wouldn't have been as dangerous to Shuttle tiles if it weren't for vibrations from SRBs. Put it all together and you get Challenger & Columbia. But the good news is that foam is not dangerous when there are no heat shield tiles to damage.

This is one reason I argue for block 2B to use advanced liquid boosters based on F-1B engines, not a new segmented solid. And advanced solids add RDX and/or HMX to the fuel mix. RDX is the active ingredient in C-4 plastic explosive, HMX is a more powerful explosive. I have been assured that military combat missiles already use this fuel mix, it doesn't explode. RDX and HMX burn to provide oxidizer, and adds a lot of energy (heat) as they burn. But it's never been tried with a solid rocket as large as an SRB. Yes, this was part of the original design for Ares. However, we now have a better proposal, replace the pair of solid rockets with a pair of liquid rocket boosters. Each booster will have a pair of F-1B engines, and RP1/LOX propellant mixture. That's the same propellant as the first stage of Saturn V, core stage of Atlas V, and all stages of Falcon 9.

I digress. Yes, the core stage of SLS will have the same foam.

GW Johnson · 2017-12-09 19:17:39

Solids can be made to burn smooth and safely. But not as segmented boosters with simple cylindrical grain designs with 4+ segments. The problem is instabilities from too long a motor chamber interacting resonantly with pressure-sensitive propellant combustion that is experiencing too-severe a bore pressure drop with 4 or more segments.

Having to pass 90+% of motor massflow through the downstream-most segment's bore raises bore Mach number, and introduces very serious erosive burning effects upon burn rate, which in turn affects the massflow and the pressure distribution. Positive feedback. Instability. Which takes the form of large pressure and thrust oscillations. "Surprise, surprise!" as Gomer Pyle would say!

As for inclusion of RDX and HMX, these particulate additives in place of some oxidizer improve performance at the expense of increasing hazardous response. That is well-known for decades. They must be relatively minor additives, or else you end up with a class 1.1 mass-detonable finished product, something very unattractive. Even worse, the hazard sensitivity of the uncured propellant to be cast may start to approach that of nitroglycerine.

Usually, when RDX or HMX are considered, it is uncured propellant hazard sensitivities that you bump into first. They are comparable to, but different in detail, from the hazards incurred with liquid explosives like nitroglycerine, one of the iron-bearing burn rate catalysts, and glycidyl azide polymer (GAP).

I recommend avoiding the liquid explosives, period. I recommend against more than trace amounts of powder explosives like RDX or HMX unless you are absolutely forced into using them.

You can use pelletized nitrocellulose up to a point, more if pre-blended with a portion of the otherwise-inert binder polymer, than if incorporated neat (another reason to stay away from GAP). "Wet" processing of dangerous powder ingredients with an inert liquid ingredient is almost always the better bet.

GW

Last edited by GW Johnson (2017-12-09 19:18:28)

kbd512 · 2017-12-16 17:52:11

GW,

Why is it that multiple smaller solids can't provide the same functionality as these behemoth 5 segment SRB's? For example, what if NASA had used 4 of the 3 segment SRB's from ATK with a 10m diameter LOX tank atop a 8.4m diameter LH2 tank, much like the ICPS configuration?

SpaceNut · 2017-12-16 18:35:04

Length is used to create burn time at pressure for the exhaust velocity in a single tube. Adding small boosters like Boeing and Atlas do only gives the initial push and stops far short due to short burn times.

kbd512 · 2017-12-16 18:41:13

SpaceNut,

I looked at the 2016 OA Motor Catalog PDF from ATK.

3 segment RSRM burn time is 133.7s, 1.67Mlbf

5 segment RSRM burn time is 131.9s, 2.89Mlbf

The thrust profiles look pretty similar to me.

SpaceNut · 2017-12-16 19:45:30

I think the fuel grains were altered for the 5 segment GW would know this better than I, but it was the way they could control the time of burn....

kbd512 · 2017-12-16 19:55:43

SpaceNut,

Yes, that's what they have to do. My point is that 4 of the 3 segment RSRM's should present fewer technical challenges than 2 5 segment RSRM's. Probably no longer the case, but when they were designing SLS, this seems like it would've been an easier problem to solve.

SpaceNut · 2017-12-16 20:09:46

The 3 unit would have required a larger launching pad from what I remember but then again thats just what they have done any how for the new rocket called SLS....just more make workfare...

Oldfart1939 · 2017-12-17 13:43:51

Here's a link to a YouTube presentation, wherein it raises questions about the SLS versus Musk's BFR.

https://youtu.be/wePA66dk5KQ

SpaceNut · 2017-12-17 16:10:51

They way to go with Nasa should be to convert what we know into a reuseable rocket like kbd512 is suggesting using the parts in another manner by some other business.
I would favor 4 x 3 segments wrapped around a 8 or 10 meter wide first stage that is recoverable to allow for a greater mass to orbit...but then again sticking with Lox/LH2 makes for a large rocket....

GW Johnson · 2017-12-17 19:29:56

To address the questions raised in posts 29-34 vis-a-vis SRB design: for a segmented booster comprised of simple cylindrical-bore grains, one to a segment, length has nothing to do with burn time and everything to do with exposed burning surface and thrust levels. These are radial burners. It is diameter that is related to burn time, and not exactly proportionally. Length is related to thrust, and not exactly proportionally.

That simple cylindrical-bore grain design has a "rainbow-neutral" set of burning surface vs web, pressure vs time, and thrust vs time curves. In principle, you stack up increasing numbers of grain segments to increase expelled mass and thrust levels, at otherwise the same pressure levels, and burn times. Except that you cannot really do this, there are severe limits. Surfaces are roughly within 20% of truly neutral. The variation in thrust and pressure is higher.

One limitation is the massflow that must pass the aftmost grain bore when the motor first lights. The more segments, the more the massflow that must pass through that last grain's bore, and the higher the Mach number of the hot flow in that bore. Higher bore Mach number is greater pressure drop down the bore, and a higher pressure difference between the forward and aft ends of the motor. This leads to two very serious and fundamental problems.

(1) A non-uniform motor pressure distribution puts a distribution of burn rates among the segments, leading to differing burn times for the forward segments (shorter) versus the aft segments (longer). Not having a definite common burn time leads to long tailoffs, and drastically lowered specific impulse for the significant quantity of propellant expelled during the reduced-pressure tailoff. (“Tailoff” refers to that portion of the pressure-time trace between when the pressure starts to fall drastically, and when it finally zeroes.)

(2) Higher bore Mach number past an entirely-experimental empirical threshold (that is very propellant-dependent) leads to erosive burning effects that significantly amplify burn rates (much higher aft than forward). These further disrupt the distribution and pattern of burn rates and surface regression from segment to segment, and also act to significantly impact equilibrium motor pressures by raising them, a positive feedback effect that can easily over-pressurize the motor.

It is difficult to mal-design a cylindrical-segment motor design to the point that you choke the bore at the outlet to the aftmost segment, but such has actually happened. It ALWAYS causes an immediate motor explosion upon ignition. The rule of thumb to prevent that is that bore min area must be at the very least twice the nozzle throat area. Far, far more than “twice” gets the more acceptable results, by far. Numbers in the 4-10 range are more typical of “good” designs. Although, bore/throat area ratio also limits bore diameter relative to insulated case diameter, and thus motor propellant loadout, by any measure conceivable. And there are several.

The other problem with stacking-up too many segmented grains is combustion stability. Stability is far "iffier" with reduced-smoke non-aluminized propellants, and far better with aluminized propellants. Best stability potential is obtained with 20% aluminum in composite propellants. More doesn't help. Aluminum also raises effective specific impulse, by increasing combustion temperatures and chamber c* velocities.

The usual stability problem in big motors relates to the first longitudinal mode, the quarter-wave organ pipe mode, of vibration. This is a closed-end / open-end type situation, with the nozzle the open-end. That model isn’t perfect, but it’s a start. Smaller motors also fall prey to radial modes and to circumferential modes. These vary quite strongly with the kind of grain design. Usually only the first couple of available modes in each of those three directions is really susceptible to excitation. The lowest frequencies are always associated with the longitudinal modes.

What was “sort-of stable” in a shuttle SRB with 4 segments can very easily be unstable in an SLS SRB at 5 segments, in otherwise the same diameters and grain design! We’ve already seen this. It's a full 25% change in the cavity length and the natural oscillation frequency. Such change magnitudes are historically very, very significant! The "theory" for this works a lot better at explaining experimental test results after-the-fact, than it does predicting them before the test. "Surprise, surprise", as Gomer Pyle would say.

Yes, 3 or 4 4-segment SRB's could do the same (or better) job of adding thrust compared to 2 5-segment SRB's, for SLS. What makes those top-level design decisions has a lot more to do with pork-barrel politics, than anything to do with solid rocket engineering. Unfortunately.

If you get the impression that solid ballistics is a whole lot more complicated in real life, than as depicted in the physics (and even the engineering) textbooks, then, yes, you have gotten the right impression. It is not a job for amateurs.

GW

kbd512 · 2017-12-19 01:43:56

GW,

As it pertains to rocketry, it seems as if every time we attempt to redesign something to function in a way it was never originally intended to function that we end up spending every bit as much time and money as if we'd just started from scratch and built whatever it was that we really needed to begin with. There's no possible way that this project could've taken longer to complete than if we'd simply rebuilt Saturn V using F-1B and J-2X engines. We're using the most expensive reusable rocket hardware in the world on an expendable launch vehicle that fails to deliver the payload that Saturn V delivered.

Why could we not have simply developed the SLS concept in stages?

The J-2X ground test program was completed before Ares V morphed into SLS, so the upper stages could've been designed first. After we determined what the second and third stages would weigh, we could've designed the first stage around the F-1B engines. The basic engine designs were iterative developments of flight proven technology. The new composite tanks that Boeing has developed have come to be ground proven technology for LH2 storage, so that technology could've been incorporated into the upper stages and inter-stages to reduce upper stage mass to contend with the inevitable weight gain that launch vehicles acquire during the development process. In the end, we'd get a 150t+ super heavy lift vehicle that would not require hundreds of millions of dollars for MLP upgrades to deal with the crushing weight of those massive 5 segment SRB's that SLS uses.

I don't think it's possible to screw this up any worse than NASA managed to, but if SLS gets cancelled then we'll know for sure. That's as bad as it gets. It's just irritating waiting forever for something that'll cost so much that we can't afford to use it, assuming Congress doesn't pull the plug first.

Oldfart1939 · 2017-12-19 10:45:28

GW & SpaceNut-

The cost so far for development of the SLS has been estimated to be roughly $11 Billion. To make it useable for a moon mission would probably add another $6 Billion to the price tag. No one has even made a WAG regarding the DSG using this monstrosity of a rocket. But if the ISS is any indicator, we can figure on at least doubling what has been spent on the SLS.

Oldfart1939 · 2017-12-19 10:58:20

IMHO, NASA should get out of the business of designing rockets, and should become the mission designer; design the mission, and then ask for private quotes to build the system. It could include milestone payment for achievement of specific goals. And NO "cost plus" contracts which allow Lockheed Martin engineers to state that "overhead is their most important product."

One of the features they should include in the request for proposals is maximum reusability of the purchased components.

RobertDyck · 2017-12-19 11:29:56

According to Robert Zubrin in The Case for Mars, cost estimate for Mars Direct in 1990 was $20 billion for research, development, construction of infrastructure, and the first human mission to Mars. And $2 billion more for each mission thereafter. That $20 billion included development of Ares. The proposal for Ares was 8.4 metre diameter core stage, 5 SSME, a pair of advanced solids, and an upper stage with a single J-2S engine. J-2X is just the updated version of the same engine, so the only difference between Ares and SLS block 2 is position of main engines. Mars Direct also included developing a capsule for the ERV. Applying inflation $20B in 1990 works out to $37.75B today. How much have they already spent on SLS plus Orion? And EM1 will launch SLS block 1, EM2 will be block 1B. How much more before block 2 or 2B flies?

Oldfart1939 · 2017-12-19 19:18:33

As Robert points out, these figures are also subject to inflationary pressures. But the biggest fault in the SLS system is lack of reusability. Each one of these very expensive vehicles is a one-shot throwaway. Totally unacceptable in light of SpaceX's performance.

RobertDyck · 2017-12-19 19:52:31

You realize the mission architecture I came up with in 2002, and have been arguing for on this forum, is based on a reusable interplanetary vehicle.

Oldfart1939 · 2017-12-19 21:44:34

Yes. But this discussion is the SLS, not your system which actually DOES make sense.

elderflower · 2017-12-20 05:12:01

Even Spacex will have to write off their boosters to achieve maximum possible payload as there will be no fuel left over for landing them.

Oldfart1939 · 2017-12-20 09:49:02

Elderflower-

Not necessarily, since the payload capacity is a design feature. The rocket should be designed for the desired payload, and inclusive of the fuel reserves required for re-landing of the booster. A reduced payload is a small price to pay for recovery of a piece of expensive hardware.

kbd512 · 2017-12-20 10:23:40

Why can't we just send everything to ISS for assembly and use NTO/MMH for in-space propulsion, just as we've always done on all real missions involving humans? That sort of stuff is ready to go right now. The actively cooled cryogenic propulsion and high-powered electric propulsion concepts are still experimental. They may both work quite well, but we're not betting any lives on it in the near future. It just seems like we could do this mission without resorting to enormous rockets. Where is it written in the rule book that everything has to be delivered to wherever it's going aboard one rocket? We have lots of tech and know-how for orbital assembly. Mating a kick stage to a module is child's play for NASA and ROSCOSMOS. We do this every few months or so.

RobertDyck · 2017-12-20 11:24:48

There were proposals in 1961 for Earth orbit assembly. The problem was they didn't have orbital assembly technology then. We do now.

NACA was founded in 1915 because manufacturers did not want to invest in new technology. They wanted to only use what was already proven. Maybe a tiny tweak to improve performance incrementally. That resulted in extremely slow development. NACA was a government agency specifically mandated to research high-risk/high-payoff technologies. On October 1,1958, NACA was dissolved, NASA was founded. The personnel of NACA were the first staff of NASA, and NASA inherited the mandate of NACA. NASA is still doing aircraft technology research.

My point is current contractors for NASA claim that anything new is just wrong. You have to only use existing proven technologies. Industry had that problem in 1915, that's what NACA was founded to fix. But NASA is now listening to those corporate executives, letting them tell them what to do. This idea that going to the Moon just *has-to* be done the way Apollo did it in the 1960s is part of that attitude.

However. My initial mission design in 2002 was based on Robert Zubrin's Mars Direct, but tweaked. Mars Society members at the time were big on the idea of using Russia's Energia launch vehicle, and Robert Zubrin himself first proposed it in his book. The Case for Mars was published before the founding convention of the Mars Society. So my initial mission design used Energia as well. But relations with Russia are not as good now. And Falcon Heavy is reusable! Reusable!

There is another concern. Each module requires a propulsion stage to rendezvous with the assembly. And docking mechanism to attach it. And those docking mechanisms can form weak points. So you don't want to break it into too many small pieces. Keeping pieces few and large has several advantages. But still, the cost saving of Falcon Heavy vs SLS is so extreme that it makes it worth it.

And there's the fact that during the Shuttle era, America didn't have technology for automated rendezvous and docking. Russia did it with Progress starting with Salyut 7, then Mir, now ISS. But America has had that capability since Dragon and Cygnus.

Last edited by RobertDyck (2017-12-20 11:28:55)

Oldfart1939 · 2017-12-20 18:03:30

I really suspect that the reason OldSpace wants to keep doing things the way they were don in the 1960's & 70's is they don't really require any innovative technology and they are insanely profitable to them.
I'm all for orbital assembly, and strongly supportive of MMH and NTO as the deep space fuels of choice. Check several of my earlier proposed mission architectures which are inclusive of both concepts.

SpaceNut · 2017-12-20 18:18:15

SLS is another of Nasa's over engineered and cost plus built here designs to which they said no to using the original parts just the way they were before they started to insist on updating all of them, re-egineering them and saying we are using human rated parts...
Untill Nasa and its contractors are booted from the gravy bowl nothing will change. If anything Nasa could be broken up into categories of work and funded seperately for each with the space policy group steering each operation rather than doing nothing. To say nothing about copying the previous policy and changing only a paragraph from it and calling it new.....

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#26 2017-12-08 21:13:50

Re: Viability of NASA SLS launch system; should it be cancelled?

#27 2017-12-08 22:04:13

Re: Viability of NASA SLS launch system; should it be cancelled?

#28 2017-12-09 19:17:39

Re: Viability of NASA SLS launch system; should it be cancelled?

#29 2017-12-16 17:52:11

Re: Viability of NASA SLS launch system; should it be cancelled?

#30 2017-12-16 18:35:04

Re: Viability of NASA SLS launch system; should it be cancelled?

#31 2017-12-16 18:41:13

Re: Viability of NASA SLS launch system; should it be cancelled?

#32 2017-12-16 19:45:30

Re: Viability of NASA SLS launch system; should it be cancelled?

#33 2017-12-16 19:55:43

Re: Viability of NASA SLS launch system; should it be cancelled?

#34 2017-12-16 20:09:46

Re: Viability of NASA SLS launch system; should it be cancelled?

#35 2017-12-17 13:43:51

Re: Viability of NASA SLS launch system; should it be cancelled?

#36 2017-12-17 16:10:51

Re: Viability of NASA SLS launch system; should it be cancelled?

#37 2017-12-17 19:29:56

Re: Viability of NASA SLS launch system; should it be cancelled?

#38 2017-12-19 01:43:56

Re: Viability of NASA SLS launch system; should it be cancelled?

#39 2017-12-19 10:45:28

Re: Viability of NASA SLS launch system; should it be cancelled?

#40 2017-12-19 10:58:20

Re: Viability of NASA SLS launch system; should it be cancelled?

#41 2017-12-19 11:29:56

Re: Viability of NASA SLS launch system; should it be cancelled?

#42 2017-12-19 19:18:33

Re: Viability of NASA SLS launch system; should it be cancelled?

#43 2017-12-19 19:52:31

Re: Viability of NASA SLS launch system; should it be cancelled?

#44 2017-12-19 21:44:34

Re: Viability of NASA SLS launch system; should it be cancelled?

#45 2017-12-20 05:12:01

Re: Viability of NASA SLS launch system; should it be cancelled?

#46 2017-12-20 09:49:02

Re: Viability of NASA SLS launch system; should it be cancelled?

#47 2017-12-20 10:23:40

Re: Viability of NASA SLS launch system; should it be cancelled?

#48 2017-12-20 11:24:48

Re: Viability of NASA SLS launch system; should it be cancelled?

#49 2017-12-20 18:03:30

Re: Viability of NASA SLS launch system; should it be cancelled?

#50 2017-12-20 18:18:15

Re: Viability of NASA SLS launch system; should it be cancelled?

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