Getting to Mars with REAL technology, & what's currently missing.

SpaceNut · 2018-07-22 09:34:29

We have in the past once we knew that Shuttle was going to done there were a number of topics to what we could have been doing.

Space tourism & - a shuttle C hotel?

Space shuttle variants - Options?

Future of Space Shuttles - How could the Space Shuttles be used?

Shuttle derived revival - Space.com

4 SRB Shuttle C configuration

A new EELV v SDV - A new spacedaily opinion piece

Shuttle C - Bigger, better, badder

OSP: Capsule v. Wings - if you had to choose right now

Might Shuttle C - save Hubble?

Down sizing the shuttle Army

Passenger module for Shuttle

We knew with all of the topics that tossing the existing struture a do an apollo program on steriods from shuttle parts was going to be a bad Idea and tht start from stratch was about the same cost.

We did look at the clustering of solids, recycled engine pods, reuse of the external tank and so many other was to use the exisitng shuttle parts and all of the eelv vehicles at the time in 2004 knowing that it was going to all end in 2010.

GW Johnson · 2018-07-22 10:27:21

All I know is that there could be far better designs for a solid SRB than shuttle SRB technology. Better joints if you MUST segment, and better propellant grain geometry designs, too. I think the community is wrong to rule out using solids when they need them.

EDIT: I think solids are best used as add-on boosters to a liquid core, rather than the first stage itself, for manned vehicles. They are pretty rough. You get a smoother ride from a liquid. Keep the solids to maybe half to two-thirds thrust at takeoff, at most.

Keep them smaller by using more of them. Smaller is easier to build, given fixed max propellant mix size. You do NOT want to overcast from successive mixes! If you can stay within one mix, then you need no segment joints.

END EDIT

GW

Last edited by GW Johnson (2018-07-22 11:28:36)

kbd512 · 2018-07-22 12:02:20

Robert,

I read your post from 16 years ago and noted that variable cycle engines are just now entering serious development for our fighters since existing turbofan technology is tapped out, so far as subsonic to supersonic fuel economy is concerned. That's a fairly good representation of the glacially slow pace of development. Efficient hypersonic engines are still the stuff of science fiction. If an efficient air breathing engine was available to take a launch vehicle to Mach 6, that still represents just 2km/s of the roughly 9km/s total dV required to achieve orbit.

Like GW, I think that using advanced solids is the way to go for SLS, so long as we're limited to the favored contractors and engine designs. Lest we forget, Congress made that decision, not NASA. ATK's Dark Knight SRB's should drastically cut manufacturing costs and simplify the design. It's a reprise of what Thiokol wanted to do to cut the cost of the STS era SRB's. There's no technical reason why it won't work, especially if the SRB's are expendable. The basic materials and manufacturing technology is now available at affordable prices.

The Thiokol (now owned by Orbital ATK) Filament-Wound Casings (FWC) Solid Rocket Motors (SRM), or FWC-SRM's worked in 1984 and Orbital ATK demonstrated that it worked again in 2013. The cost reduction associated with the simplified motor is 40%, or $31.7M per pair of boosters. Between the Isp density improvement of the more advanced propellant and the lighter composite casings, the payload improvement is 15.1t. Orbital ATK's Omega rocket also uses composite FWC-SRM STS/SLS booster cores.

Boeing estimates that the external tank mass reduction, using composite technology to store propellants to feed the RS-25's, to be 30% lighter on the low end and 40% lighter on the high end. The test articles, which were of substantial size, were robotically fabricated in days, did not require autoclaving, and failed at pressures greater than the specified proof testing pressures.

The composite technology is very real and it's here to stay, so it might as well be incorporated into SLS.

RobertDyck · 2018-07-22 13:00:21

GW,
I respect your experience. My concern with solids is the segment seals. And the proposed advanced solids use RDX or HMX or both. Those are military explosives. RDX is stabilized with a plasticizer to make C4 explosive. I realize the "advanced" fuel has been used for missiles before, but never for a rocket as big as a Shuttle/SLS SRB. We've seen far too many catastrophic failures. F-1B is just an F-1A with 21st century electronics, and certain parts made with 3D printers. F-1 is the engine of the Saturn V first stage. It's reliable, and RP1/LOX is a very well demonstrated propellant mix for first stages.

kbd512,
Why are you bringing up something from 16 years ago? Boeing's website made wonderful claims, and that was soon after X-33 / VentureStar was cancelled. America was looking for a new shuttle. Has it really been that long? Wow.

kbd512 · 2018-07-22 13:51:10

Robert,

I was merely trying to illustrate how long it took for variable cycle engine development to come to become a priority. From time to time, I go back and re-read old posts. SpaceNut posted some links and I went back and read through them.

I still don't understand what the aversion to explosives happens to be. All rocket propellants are explosives. The best propellants also happen to be the best explosives. In chemical rocketry, there's no way around that.

RobertDyck · 2018-07-22 14:31:41

kbd512 wrote:

I still don't understand what the aversion to explosives happens to be.

SpaceNut · 2018-07-22 14:50:16

Bad Oring seal....

I think that there is only 18 srb casing remaining to make up for about 9 flights before they are out of them since these are not being recovered.

https://www.nasaspaceflight.com/2015/10 … en-flight/

“As part of our proposal for the NASA Advanced Booster Engineering Development and Risk Reduction (ABEDRR) solicitation, Orbital ATK generated a preliminary design of a solid booster with a graphite composite case, new propellant, and other component improvements that would enable NASA to achieve the 130 mT mass goal.”

https://www.nasaspaceflight.com/2013/01 … d-for-sls/

One can see that segmented srb due to size and location of where they are made are still a possible issue in cold weather.

As far as the SLS goes pg 8 & 9 slide of this Game Changing: NASA’s Space Launch System and Science Mission Design

Tonnage in one shot seems to be the driving force for SLS that and keeping a standing army employed.

GW Johnson · 2018-07-22 18:09:00

AP is a mass-detonable monopropellant explosive all by itself. RDX and HMX have more yield used neat, but present little added risk if used as small-percentage ingredients (modifiers), not main ingredients (like AP). These have decades of experience in safe-to-use rocket motors. I have used them myself, when I worked in that business.

The Challenger "explosion" wasn't what it looks like to the amateur observer. It WAS NOT a propellant explosion! If you look, both SRB's are flying free under their own thrust, including the one that leaked and caused the "explosion". The solid propellant explosion risk occurs during propellant mixing. The main risk during the motor burn is a case overpressure burst, pure and simple.

The most common cause is excess exposed burning area from a crack or void in the propellant grain assembly. Been there and done that many times investigating why a test motor exploded. Detail causes can vary widely, but nearly all produce excess burning surface. You get all that resolved long before you go into production.

In Challenger's case, the leak at the failed joint spit a tongue of flame that cut a lower retention strut, before the joint itself could part. The booster hinged about the upper strut, and pushed its nose cone through the side of the center tank, which then immediately collapsed and let its liquid propellants loose to "explode", meaning deflagrate, not detonate. THAT is what looks like an explosion. It was not a cause, it was an effect.

The center tank's collapse freed the orbiter and the two SRB's. The orbiter pitched up broadside to the airflow due to its main engine thrust angle going through a cg that was no longer there. All airplanes break up at 90 degree AOA, especially at anything above landing speeds, and this occurred supersonic at fairly low altitudes (around 30-40 kft). The force levels were not fatal to the crew, it didn't even knock them cold. They died on impact with the sea at 200 mph, after arcing up to about 60 kft.

Had the strut not failed first, the joint would have parted. The PA kick load on 120 inch diameter at 900 psig is around 10 million pounds. That would have also broken up the vehicle at something under average 5 gees. Might not have knocked them cold. Certainly not immediately fatal.

GW

ps -- I have even used pelletized NC in rocket propellant as a modifier additive, up to 5-10%. Met all the class 1.3 tests, and got rid of all the smoky metal, too.

Last edited by GW Johnson (2018-07-22 18:21:53)

kbd512 · 2018-07-22 18:28:57

GW,

What's the best way to keep the casings or joint seals from bursting, since that seems to be the only logical objection to their use?

SpaceNut · 2018-07-22 19:47:52

https://www.nasaspaceflight.com/2013/10 … ter-drive/

Orbital ATK optimistic about proposed KSC rocket

Still digging for the composite casings images but not much data so far,

kbd512 · 2018-07-22 20:43:43

SpaceNut,

Google "orbital atk filament wound casing". There's plenty of images of the casings.

Edit:

One-on-One with ATK's Charlie Precourt about composite materials and NASA's Space Launch System

Last edited by kbd512 (2018-07-22 20:46:40)

SpaceNut · 2018-07-22 21:50:20

So far none of them show the sealing methods of the joints only the casing sidewall. but thanks
The old srb went from a single joint to a multi tongue slot with orings and firestop materials.

GW Johnson · 2018-07-23 11:40:33

To answer Kbd512’s question, there are two different solid propellant motor problems here that lay persons often conflate, because of the Challenger disaster. One is sealing failures, the other is over-pressure burst. Actual explosion of the propellant itself is a problem during mixing, not the motor burn. I think I have written about this issue before on these forums, but here it is again.

There are many openings in a rocket motor case that must be sealed reliably. Usually, there is some sort of forward closure, either at full case diameter, or at some reduced diameter with part of the forward closure integral with the case. There is almost always an aft nozzle closure , usually at full case diameter. There can be other closures for igniters and instrumentation, usually small, but still requiring effective sealing. If you must segment the motor, there are segment joints.

What you must deal with is a very hot gas that is polluted with a very significant solids content, largely a mix of carbon black and metal oxides, called slag. The thermal radiation from these solids utterly dwarfs the thermal radiation from the hot gas itself, by around a factor of 100. Plus, there is the sandblast erosion effect of these solids, increasing as flow speed increases. These particles do not cool as the pressure drops, they are incompressible! They are very nearly chamber temperature-hot.

You don’t seal this kind of dirty gas with multiple O-rings “for redundancy”. That doesn’t work, because the upstream O-ring is always breached at a single point, collimating the leak stream as a concentrated blast upon the second O-ring. The oil field perforation-explosive guys have long known this, as have the solid rocket motor manufacturers. That is the fundamental design mistake NASA made with the Shuttle SRB joints that killed Challenger’s crew, and their redesign made that problem worse by going to 3 O-rings from the original 2. The ONLY reason it never failed again was they never flew that cold again.

You use ONE AND ONLY ONE O-ring in each joint, and you whole-motor pressure-leak-check it with air at very modest pressure. If it holds at low pressure, it will hold at high pressure, given a proper O-ring seal design. You find that out (“did you design it right?”) in hydrotesting, before you ever fire a motor, and then you verify it again in your live motor test firings, long before you ever go into production.

Now, there’s joints in the insulation inside the motor where these closure and any segment joints are located. You DO NOT “pooky” these insulation joints with any sealant, you leave those small gaps wide open! Otherwise, the pressurizing hot gas wormholes through the “pooky” at a single point, creating a sandblast jet attack on your O-ring.

NASA made that mistake as well, shown definitively by the chunks missing from many of the upstream O-rings upon SRB disassembly. The wide-open gap will pressurize evenly at much lower gas velocities all around the O-ring: that simply eliminates any jets that might cut it! Once pressurized, there IS NO BETTER INSULATOR than a static gas column. The surviving pooky just increases the chamber heat transferred to the O-ring.

NASA’s joint designs were leaked checked BETWEEN the O-rings with air at full motor pressure. This ALWAYS drives the upstream O-ring to the WRONG SIDE of its groove! That O-ring CANNOT seal during motor pressurization, until it has been driven to the correct side of its groove. If you “pooky” the insulation joint, the wormholing jet hits the O-ring at a single point, until the pressure buildup can later move the whole O-ring to the downstream side of its groove. That concentrates the strain as a sharp V-shape bend at the jet impact point. If the O-ring is cold, it will snap there, and it did at 29 F the morning of the Challenger disaster.

Richard Feynman demonstrated the effect quite conclusively with a piece of O-ring and a glass of ice water during the Rogers Commission hearings, on live TV. Made idiots of the analysts who modeled the O-ring as pressurized on a broad front, while managers insisted on putting “pooky” in the insulation joints. Analyzed one way, built differently. Worthless analysis results. BAD management. So, why should loss of Challenger be a surprise? It was almost pre-ordained by this idiotic management.

In Challenger’s case, that upstream O-ring failure collimated a jet onto the second O-ring and cut it, too. That hole was plugged by some debris eroded from the first seal and the case material around it. Then the debris plug shook loose 73 seconds into the flight, with the results we all saw. BAD, BAD, BAD joint design that was insisted-upon by NASA, when Thiokol knew better. However, you cannot tell the government customer he is wrong, without losing your contract. This idiocy is what killed that crew.

The final thing to worry about is that intense thermal radiation from the motor environment. As long as there is a 90 degree bend in the gas path between insulation joint and the case joint, that radiation cannot get to the vulnerable O-ring. You will need some extra case material to act as a heat sink during the pressurization transient, but you usually need that anyway to hold the structural loads at the joint. This heat sinking effect works better in metal cases than composite cases, because the metal has better thermal conductivity. With composites, you have to do very careful transient thermal modeling to ensure you got the design right.

That takes care of sealing. Overpressure burst can come about from one of three causes: (1) something increased the available burning surface, (2) something plugged the nozzle throat, or (3) the case insulation failed, letting flame eat through the case.

Equilibrium solid motor pressure can be expressed as P = (rho S a k exp / CD At gc)^(1/(1-n-m)), where rho is propellant density, S the burning surface, a and n come from the burn rate model r = a P^n, k and m come from the characteristic velocity model c* = k P^m, exp is the motor expulsion efficiency, CD is nozzle discharge coefficient, At is the nozzle throat area, and gc the gravity constant for inconsistent units. The exponent can be quite large, since n is usually in the 0.2 to 0.7 range, with m being only a slight effect at something like 0.01. For n = 0.5, this exponent exceeds 2. At n = 0.7, it exceeds 3. A 10% surface change is over a 20% motor pressure change, maybe well over 30%.

Cause (1) extra surface is by far the most common. It comes about from voids or cracks in the propellant grain assembly, usually traceable to quality control problems in propellant cast processing, or sometimes to inadequate thermal cycling design when going too cold in your soakout, which can crack the propellant or tear loose its case bond. The higher the solids during mixing, the harder this stuff is to cast. High solids confers significantly better performance, but requires pressure-pack casting under very hard vacuum. And it requires full-motor X-ray to confirm every single motor’s quality. Soakout design adequacy is usually confirmed in development test firings that are soaked out fully cold.

Cause (2) plugged throat has almost been eliminated by nearly all manufacturers, by no longer using inhibitor-surface blankets as part of the propellant grain design. There is usually nothing in a modern design that could come loose inside a motor, and not pass the nozzle throat easily. It is more reliable to just design the propellant to burn as-cast. The only inhibition is the case bond joint. That gets verified in development test firings long before production. Usually cold soak is also most critical for the bondline.

Cause (3) insulation failure is quite rare, and nearly always identified and corrected during the development test fairings. For such a thing to ever happen in a production motor is an indicator of very serious quality control problems in case insulation processing.

All in all, if you have done your design well enough, verified it adequately in testing, and maintained stringent quality control throughout all production, you will NEVER see an overpressure burst event. The only sealing failures I have seen were bad joint designs: NASA’s SRB’s and an overlap weld on a USN-designed Tartar-Terrier gas generator cartridge case. If the overlapping pieces were made and fitted too tightly precise, you’d get the gas wormhole effect between them, which cut the weld. Lesson: wide open pressurization gaps, just like with O-rings.

Hope all that helps.

GW

Last edited by GW Johnson (2018-07-23 11:53:36)

RobertDyck · 2018-07-23 16:30:52

GW,
What I said is not because I don't understand what you explained, it's because I do. You have pointed out serious issues with segment joint seals. I am still concerned. The new proposed advanced boosters will use carbon fibre composite casings instead of steel, but they will still be segmented. New SRBs for SLS are 5-segment, but the manufacturer ATK has said their design for advanced solids go back to 4 segment. That means they still have segment seals.

Propellant mixture is my greatest concern. You posted "RDX and HMX have more yield used neat, but present little added risk if used as small-percentage ingredients (modifiers), not main ingredients (like AP). These have decades of experience in safe-to-use rocket motors. I have used them myself, when I worked in that business." ... "ps -- I have even used pelletized NC in rocket propellant as a modifier additive, up to 5-10%. Met all the class 1.3 tests, and got rid of all the smoky metal, too." By "NC" I assume you mean nitrocellulose, also known as guncotton. Your statements assume trust that all propellants will be blended evenly, without any pockets of explosive. NASA expressed distress when the Senator from Utah insisted segmented solid rockets based on those of a Titan III be added to Shuttle. One concern is there's no way to throttle a solid rocket, nor abort should something go wrong. The Shuttle could not abort until solids separate. The design before that was Two-Stage-To-Orbit, with a piloted fly-back booster. That meant the fly-back booster would have an aircraft skin over the insulation, so no chance of ice-filled foam breaking off and striking heat shield tiles. Not only did the external tank expose insulation to supersonic air stream, it also exposed that foam to extreme vibrations of SRBs. STS-1 resulted in white tiles shaken off the Shuttle orbiter entirely. What causes the vibration? Explosion of AP (ammonium perchlorate)? Then best case scenario, explosion of RDX or HMX would cause greater vibration. Remember, RDX is the explosive in C4, HMX is more powerful. Worst case scenario, that I'm worried about, is propellant is not evenly mixed, resulting in a pocket RDX. That causes explosion of the SRB, causing a hole in the casing which creates shrapnel that blows a hole in the LH2 tank of the core stage, then total vehicle loss similar to Challenger. I expect explosive within a carbon fibre composite casing greater damage and consequently faster loss of vehicle than Challenger.

SpaceNut · 2018-07-23 19:28:27

Provide the srb with venting capped tubes (plugged such that it can not burn through) that run into the core of combustion with a means to explode the cap to allow for the exhaust to exit and the booster to lose thrust. The size and number of these depends on the internal pressures seen that needs to be reduced. The vent tubes would angle upward on the outward sides of the srb's such as to face away from the core of the rocket after they are open one can then jetison the boosters and launch the escape from launch protocal to finish to a safe landing.

GW Johnson · 2018-07-23 19:47:16

Robert:

Propellant mixing is done under vacuum starting with the liquid polymer and one of the solid powder ingredients. This mix goes on for several minutes before stopping, breaking vacuum, and adding the next powder ingredient. Vacuum is restored and mixing resumes for several minutes.

This process repeats until all solid ingredients are well-incorporated. The last ingredient added is the polymer’s curative agent, after which mixing under vacuum proceeds for something on the order of half an hour. Then the mix is ready to cast. It will cast from a valved opening in the bottom of the mix bowl.

No reputable manufacturer would ever short the mix times or mixer blade power and rpm. Clumps of unmixed stuff leads to irreproducible burn rates at the very least, as well as irreproducible properties of all kinds. This stuff can be seen in quality control X-rays of the motor, as well as voids and cracks. Clumps of solids in finished propellant is prima facie evidence that the mix procedure is quite wrong. That is a serious quality control issue best addressed very up front in the propellant development process that precedes motor development.

Many manufacturers mix at lower solids content so that apparent viscosities are lower, allowing gravity casting down a flexible sleeve into the motor case and cast tooling, like concrete. This is done under vacuum so that voids can slump closed without any air film forming a crack line or bubbles. Lower solids content (~80% or so) leads to lower performance, primarily achievable Isp.

High-solids propellants (up to 86-88%) have much higher apparent viscosities, dominated by fully thixotropic effects. These definitely have higher Isp performance, but ABSOLUTELY CANNOT be gravity cast down a sleeve! You must physically connect the mix bowl to the case-with-tooling, and put a big piston into the top of the mix bowl resting directly on the propellant, plus a pressure closure over the top of the bowl.

Putting shop air pressure on top of the piston forces it down, extruding propellant into the motor case-with-tooling, which is also under hard vacuum. When the sight glass shows “full”, vacuum is broken at the bowl-case connection, and weep holes opened, so that pressure-packing can be done by forcing the piston down with even higher pressure. This forces closed any vacuum-filled voids in the stiff material. Tooling is Teflon-coated steel. Cases are insulated before casting.

From there, in either solids content range, carts of freshly-cast motors go to the cure ovens for a day or two at about 250-300 F or thereabouts. The idea is to vulcanize the rubber binder at a higher temperature than it will ever see in service, just like a tire. These are big room-sized ovens.

You must X-ray inspect every single high-solids motor made this way. The kind of quality control this requires is quite extraordinary. But, you get a motor with significantly higher performance that way, something the sleeve casters simply cannot achieve. Only a few manufacturers ever did this. I worked at one, actually the very best in the business. That plant was closed more than 2 decades ago.

What I have described here and in the other posting is the way very high-quality, safe, and reliable composite solid propellant motors are made, including those that perform the best. Millions upon millions of these things have been made for military weapons, and with never a leaky seal or a motor explosion in service, since the 1970’s, anyway.

Most of what I describe you will not find in textbooks, and not in very many reports, either. This is mostly engineering art, passed-on on-the-job from the experienced hand to the newbie. It doesn’t get passed-on in outfits that do not want expensive oldsters on their payrolls. That kind of mismanagement is why so many companies end up reinventing so many wheels.

Double base propellant processing is different, and I know far less about it. Pelletized NC is mulled with the other solid powder ingredients into a uniform dry mix, and packed into cases fitted with cast tooling. These are flooded from the bottom with neat NG until everything is wetted to the top. During cure, the NG gelatinizes with the NC into a plastic that encapsulates the other dry powder ingredients. Handling neat NG is extremely dangerous. Composite-modified double-base propellants use an oxidizer as part of the dry powder ingredients: AN or AP. That adds to friction sensitivity during processing. And that’s about all I know about that.

As to whether you WANT a solid motor, that is a different question entirely.

There is an inherent physical chemistry link between combustion energy release from the burning propellant, and the fluid mechanics of the hot gas flow toward the nozzle. It shows up as oscillations in pressure (and thrust), that are usually a small percentage of the average pressure (and thrust) level. These are most accurately seen with analog trace recordings at megahertz frequency. The modern digital stuff usually obscures what is really going on.

When these percentages get larger, we call it “combustion instability”. If there is a natural oscillation mode inside the ever-changing motor geometry that matches the band of natural oscillation frequencies, then that resonance drives the oscillations to very large percentages of average levels. These can blow a motor up. Happens often in early development testing.

This is analogous to, but not the same as, combustion instability in a liquid rocket engine. Those are usually traceable to vortices of mixed but unburnt propellants breaking away from a solid surface and then sucking in combusted gas as an ignition agent to suddenly explode the vortex. That process is more fluid mechanical in nature, and most often driven by incorrect injector plate geometry.

The 4-segment Shuttle SRB’s had significant (but still small-percentage) pressure and thrust oscillations. Not something we would classify as unstable, but certainly able to excite structural resonances in the vehicle. As best I can tell, the frequency corresponded to the first longitudinal mode for the motor volume from forward dome to nozzle contraction. It’s an open end/closed end quarter-wave mode.

The 5-segment designs initially all had the same diameter but a longer length until recently, first in Ares, and now in SLS. The longer length is apparently more resonant, leading to much stronger oscillations, even dangerous ones. I did notice that the most recent proposed SRB designs are now larger diameter than before. That is one way to possibly reduce resonant response. Possibly. No guarantees.

Point is, solids are always a rough ride, and if the instability isn’t addressed early on, a dangerous one, and an expensive one to fix. You have to do the motor design “right” from the get-go to minimize the roughness, and that very often conflicts with the other shape requirements the customer may have for his motor. Most notably, customers hate it when you want to change length and diameter to settle out the resonant instability. Best done at the paper vehicle design and development motor testing stage. Unless you mismanaged your program and froze the vehicle design before you had a reliable motor.

Going to metallized propellant helps damp some of this oscillation out. The swarm of particles in the chamber seems to interfere in some way with the propagation of the waves. Models for this are ENTIRELY empirical. This is NOT something reliably predicted with CFD. Low or no-metal propellants are much more prone to instability problems.

Hope that helps.

GW

GW Johnson · 2018-07-23 19:51:59

Spacenut:

I have to think about what you said. You are proposing a thrust-termination rig of some kind. None of those that I ever saw worked. The pressure x area kick loads are just too high (around 10 million pounds for a Shuttle SRB). The vent area has to be essentially full case open area. Typical small motor pressures are ~2000 psi. Big motor pressures are nearer 900-1000 psi, still very high. Even if you are successful, there is still some thrust, if the propellant does not extinguish (some do, some don't).

GW

SpaceNut · 2018-07-23 20:08:22

A booster has a nozzle opening area to which in order to reduce that we are adding at various locations openings up the length on the outward side new exit nozzles in a drag arrangement or reverse thrust orientation as compared to the nozzle at the end of the srb. Each vent is a portion of the original area and if the number of them is great enough the pressure exiting each will be lower than the internal pressure to be releaved.
So if we only have a single opening at the opposite end of the srb then the pressure is 50% of the total to exit out of both, increase the number along the length and we get a reverse thrust that is greater than the forward thrust.

kbd512 · 2018-07-23 20:14:39

GW,

Is SpaceNut talking about rapidly de-pressurizing the combustion chamber to extinguish the motor?

GW Johnson · 2018-07-23 20:17:02

Yep. P x A kick loads are usually unacceptable in that approach.

GW

SpaceNut · 2018-07-23 20:28:11

Not trying to extinguish but to slow forward motion as the boosters ran for 70 seconds burning through to the external tank.

So a slow venting that times opening vent area tubes in a balancing progression down the length will force it to extinguish in what ever interval of time is required.

What I am thinking of would be simular to what happens with a musical recorder or tin whistle.

GW Johnson · 2018-07-24 08:31:52

Well, like I said, I saw several efforts aimed at either extinguishment or thrust neutralization during my work in the industry. These were feasibility demonstrations, not system development efforts. Some of this was for the USAF ICBM programs, some was for the USN SLBM programs. Most of the experiments yielded unattractive results, the few were quite spectacular failures.

Your venting will produce thrusts of its own, and if opened suddenly, very impact-like force damage results. In any case, you must deal with the vent thrust or forces you produce: magnitude, direction, and point of application. Flight vehicles are extremely sensitive to all three of these.

GW

GW Johnson · 2018-07-24 08:43:46

Getting back to the other technologies were were discussing, where NASA is with electric propulsion is described in the following excerpt. They seem to be exploring the 100 KW class. At Kbd512's 54 N/MW, that is 5.4 N of thrust they are talking about, or about 1.2 lb. Applied to, say, 10 metric tons of spacecraft, that's in the 0.00005 gee acceleration class. That's still way too low, and I bet a spacecraft capable of generating 100 KW will weigh more than 10 tons.

Considerable scaleup still remains to be done. My fear is that the power supply weight will grow much faster than the thrust will. That leaves you with unacceptably low accelerations regardless, for interplanetary travel.

My hope is that power supply weight grows more slowly with power than thrust increases. In that case, there is a size at which spacecraft accelerations on the order of 0.01 gee can be achieved. That actually is practical for interplanetary travel.

I looked at NASA because what they est is a lot closer to flight hardware than any sort of grant work in academia. NASA is still 2 or 3 orders of magnitude away from practical acceleration values.

GW

From AIAA “Daily Launch” email newsletter for 7-24-18:

NASA Prepares For “Key” Space Electric Propulsion Tests.

Aviation Week (7/24) reports that NASA teams are preparing for ground tests of two “high-power electric propulsion systems” this November in a “key milestone in NASA’s plans for the exploration and commercialization of space beyond Earth’s orbit.” Under the NextSTEP program, Aerojet Rocketdyne and Ad Astra Rocket will “conduct tests in which their thrusters are planned to run for 100 hr. at a power level of 100 kW. “

GW Johnson · 2018-07-24 08:46:10

Here is a piece of useful technology we have not discussed before: space lubricants. Looks like they have something vacuum-qualified that will go hotter than my 30-year-old can of C-5A. Goes colder, too. That's actually good.

GW

From AIAA “Daily Launch” email newsletter for 7-24-18:

NASA Cites Space Lubricant As Invention Of The Year.

KAKE-TV Wichita, KS (7/23) reports that NASA’s Inventions and Contributions Board has named the PS/PM400 lubricant its “2018 Government and Commercial Invention of the Year.” The lubricant is a “fourth-generation material” developed by the agency that is “able to continue lubricating in temperatures that range from negative-300 degrees to 1,700 degrees Fahrenheit.” According to NASA Senior Technologist Christopher DellaCorte, who helped develop the material, the additional temperature capability “opens up the design space for machines and mechanisms to place moving parts right inside the hottest environments, like on the surface of Venus or deep inside aircraft engines, without the expense or problems affiliated with providing active cooling.” NASA’s PS/PM400 could potentially be used in rocket engines, aircraft turbines, and with a “wide range of other industries, such as with drone makers, large equipment manufacturers, and even in heat treatment furnace conveyor systems.”

kbd512 · 2018-07-24 10:40:08

GW,

The 100 hour ground test has already been completed at 98kW to 102kW, the test results were published, and I believe I provided a link in another thread on the best propellants and propulsion systems for Mars. The output thrust doesn't actually scale linearly with input power, but I used a linear output projection from a 100kW input to illustrate the thrust levels achievable if such were the case. A larger 1.2MW to 1.4MW version of the NHT (Nested Hall Thruster) system with more concentric channels than X3 is expected to produce 216N of thrust and weigh 320kg. Test observations have very closely matched actual thrust achieved, so I expect that to be the achievable case in the future with larger thrusters capable of accepting more input power.

I presume Ad Astra's system is VASIMR. Contrary to what Dr. Zubrin asserted, Ascent Solar already has a lab demonstrator of a variant of their thin film solar array technology that's producing 2.25kW/kg and I think I read something about design work on a 2.75kW/kg thin film technology. The 1.4kW/kg array already flew in orbit for more than 6 months for JAXA, as a proof of concept demonstrator for their upcoming Jupiter probe. Therefore, Dr. Chang-Diaz's assertion that he could get us to Mars in 39 days if he had a 1kW/kg power source may be proven correct, but not by using nuclear power. A speculative aneutronic fusion power source was projected to achieve 1kW/kg, but current generation flight proven solar arrays have already surpassed that power-to-mass ratio at 1AU. Thus, it would appear that the business of predicting the future is fraught with faulty assumptions based on current technology.

The thin film solar arrays, combined with "Hoytether" wiring (originally designed for decidedly higher amperage electric sail applications), will ensure that the combined mass of the solar arrays and connection wiring remains quite small. In LEO, I expect that the total mass of the solar arrays and wiring for the panels themselves will remain at 1MW/t using existing flight proven technology, and dropping as you move away from the Sun. Even so, the mass of the arrays and wiring remains quite small for multi-MW-class systems. The PMAD is another story. Maximum voltage minimizes the mass of the PMAD. The thin film arrays help there, too, achieving up to 270V per module. Until dielectric breakdown of the insulator occurs, high voltage keeps the diameter of the wiring small and the mass within tolerable limits.

The point is that high power solar electric propulsion systems are rapidly approaching practicality as the solar array technology rapidly outstrips the power-to-mass ratio of other speculative power sources. The only nuclear technology of practical use for mass reduction would be gas core nuclear thermal engines. Since we're not developing those, I see no point in not using SEP for what it was intended for.

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