GW Johnson Postings and @Exrocketman1 YouTube videos

RGClark · 2024-05-27 18:25:29

GW Johnson wrote:

…
You may want to use a bit higher. Fine. Do so. How about 440 sec? Corresponds to Vex = 4.315 km/s. We'll use that.
The mass ratio required to reach destination is exp(dV/Vex) = 8.630. The corresponding propellant mass fraction Wp/Wig = 1 - 1/MR = 0.884, or some 88.4% of the liftoff mass is propellant. THERE IS NO WAY AROUND THAT. My other numbers are right in that same ballpark. Go look for yourself. Go do the calculation for yourself. I just gave you the equations.
That leaves 0.116 (or 11.6%) of the liftoff mass to be both inert mass plus payload mass. THERE IS NO WAY AROUND THAT, EITHER! And my other numbers from the other studies are also right in that same ballpark.
Just remember, the sum of payload plus inert is your dry-tanks burnout mass. Add the propellant to it for your filled-tanks ignition mass. MASS MUST BE CONSERVED! Cannot play games here!
Now if you have 11.6% allowance, and you believe you can build an all-expendable stage with 4.5% inert, which has been done, then that leaves 7.1% max payload fraction…
GW

I’m more optimistic about a reusable hydrolox SSTO. Let’s say the gross mass was 200 tons, a little less than the Delta IV core, a little more than the Ariane 5 core.

At a mass ratio of exp(dV/Vex) = 8.630, that’s (dry mass + cargo) of 200/8.630 = 23.2 tons, with 9 tons dry mass and 14.2 payload. As you noted that’s 7.1% payload fraction. But 7.1% exceeds any rocket that now exists or ever existed, no matter how many stages or variations of propellants used. The Falcon 9 for example is only in the range of 4% payload fraction. On that basis alone it should be pursued even as expendable.

But I think even a reusable version is doable. Back in the day, there were discussions online by those in the industry on how to do it. Here it was suggested the landing gear weight would be in the range of 3%:

Landing gear weight.
(2) TPS (heat shield): the figures I hear for this are around 15% of the
>orbital mass

Could be... but one should be very suspicious of this sort of parametric
estimate. It's often possible to beat such numbers, often by quite a large
margin, by being clever and exploiting favorable conditions. Any single
number for TPS in particular has a *lot* of assumptions in it.

> (4) Landing gear: about 3%

Gary Hudson pointed out a couple of years ago that, while 3% is common
wisdom, the B-58 landing gear was 1.5%... and that was a very tall and
mechanically complex gear designed in the 1950s. See comment above
about cleverness.
https://yarchive.net/space/launchers/la … eight.html

Remember this is only for the dry mass that would have to be supported by the gear on landing. The fully loaded rocket would be supported by the launch pad on takeoff. So 3% of 9,000 kg for the landing gear is only 270 kg.

That passage I cited also mentions a thermal protection weight of 15%. That was Apollo era estimates. The PICA-X developed by SpaceX weighs about half as much so call it 7.5% of the landed mass. That’s 675 kg.

Assuming powered, vertical landing, almost all the reentry velocity is cancelled out aerodynamically for a vehicle entering broadside. You only need then to supply approx. 100 m/s to cancel terminal velocity. For engines like the SSME the sea level Isp is in the range of 366s. This requires about 250 kg of propellant for a 9,000 kg dry mass vehicle.

All together that’s about 1,200 kg subtracted off from the payload. That still leaves a 13 ton payload for the reusable vehicle.

Bob Clark

Last edited by RGClark (2024-05-28 01:05:33)

kbd512 · 2024-05-28 00:50:54

GW,

A quick article on high-temperature composites:

CAMX 2022 exhibit preview: Renegade Materials Corp.

REACH-compliant versions of all polyimides are available. Renegade Materials also offers RTM-1100 and MVK-10 polyimide RTM resins used for molding complex geometric parts seeing service over 600°F.

These parts are being incorporated into commercial airliners, and they routinely experience temperatures of 316°C. If these parts did not work as advertised, then you'd have a lot of dead passengers in very short order. These materials have been exhaustively tested at this point. They do work, believe it or not. 20 to 40 years ago, the composites technology you knew was exceptionally primitive compared to what is available today, both in terms of fiber and in resin matrices. Engine thrust structures for orbital class rockets are now being fabricated from composites.

Rocket engine thrust frame proves a strong candidate for composites conversion

Note how Airbus is using Hexcel IM7 and CYCOM 5320-1, a 400ksi composite. After Boeing and NASA spent the time and money to evaluate it's performance with LH2 tanks, ESA is now adapting that same fiber and resin matrix technology, which passed all tests, for their own upper stage applications.

T1100G fiber tape / tow produces 502ksi composites with 60% fiber fill.
T1200 (a new fiber introduced last year by TorayCA) should produce 572ksi composites with the same resin and fiber fill ratios.
T1200 plus SWCNT resin matrix additive produces 1,001ksi composites.
CNT produces 957ksi to 5,221ksi composites with a 60% fiber fill ratio, although fill ratio is typically higher than that because the fibers are so much thinner. I have not included these composites in discussion because they are not "current technology" in the sense of having been used as parts of operational pressure vessels or major aerospace structures, although they are presently being tested for those applications. The use of T1100G and IM7 is now very common. T1200 is intended for the exact same applications as T1100 fiber. Hexcel IM7 fiber is 820ksi, IM9 fiber is 913.5ksi, and IM10 fiber is 1,000ksi (equivalent to T1100 fiber's tensile strength). I presume an "IM11" fiber will be introduced to compete with T1200.

It's not alien technology and it's definitely not magic. Time and technology marches on. It still looks crazy to anyone used to thinking about materials strength and stiffness in terms of steel or Aluminum or possibly Titanium (the "wonder metal" that doesn't do anything better than high strength steel). Composites are our new aerospace structural material technologies, and they greatly surpass the mechanical properties of all known metals, except for temperature resistance, but even there ceramic matrix composites (CMCs) are taking over where metals fail.

As I said before, for any practical SSTO, only pins, bearing surfaces, shear bolts, and certain high temperature parts will still be made from high strength steels. Nozzles for engines will be non-regeneratively cooled RCC with UHTC coatings. When we figure out how to 3D print CMCs for combustion chambers and turbine casings, metals will be applied sparingly to rocket components.

RGClark · 2024-05-28 01:36:35

GW Johnson wrote:

…
You may want to use a bit higher. Fine. Do so. How about 440 sec? Corresponds to Vex = 4.315 km/s. We'll use that.
The mass ratio required to reach destination is exp(dV/Vex) = 8.630. The corresponding propellant mass fraction Wp/Wig = 1 - 1/MR = 0.884, or some 88.4% of the liftoff mass is propellant. THERE IS NO WAY AROUND THAT. My other numbers are right in that same ballpark. Go look for yourself. Go do the calculation for yourself. I just gave you the equations.
That leaves 0.116 (or 11.6%) of the liftoff mass to be both inert mass plus payload mass. THERE IS NO WAY AROUND THAT, EITHER! And my other numbers from the other studies are also right in that same ballpark.
Just remember, the sum of payload plus inert is your dry-tanks burnout mass. Add the propellant to it for your filled-tanks ignition mass. MASS MUST BE CONSERVED! Cannot play games here!
Now if you have 11.6% allowance, and you believe you can build an all-expendable stage with 4.5% inert, which has been done, then that leaves 7.1% max payload fraction. …
GW

GW, you’re assumed 4.5% inert mass, i.e., dry mass, fraction may be too optimistic for a hydrolox stage, leading to that unexpectedly high 7% payload fraction.

Remember hydrolox is less dense than kerolox resulting in a lower mass ratio, equivalently higher dry mass fraction. Commonly a hydrolox stage might be at ca. 10 to 1 mass ratio, corresponding to 1/10 = 0.10 dry mass fraction, 10%. Recall the famous Centaur upper stage for example.

So let me redo that calculation I made above for a 200 ton hydrolox SSTO. Using exp(dV/Vex) = 8.630, the total final mass reaching orbit, dry mass + payload, would be 23.2 tons. But 20 tons of that would be dry mass of the vehicle, leaving only 3.2 tons payload mass. This means only a 1.6% payload fraction, a much more believable number, significantly worse for example than the Falcon 9.

It also means very little payload if any once you add on reusability systems.

Bob Clark

kbd512 · 2024-05-28 04:48:11

Dr Clark,

STS delivered 8.08% of its liftoff mass to a 200km orbit, ignoring the fact that the external tank was immediately discarded.

2,030,000kg - Space Transportation System wet mass at liftoff
110,000kg - Space Shuttle Orbiter wet mass
27,500kg - payload able to be carried inside the Space Shuttle, to a 200km circularized orbit
26,536kg - Space Shuttle External Tank dry mass
164,036kg - Total mass accelerated to orbital velocity sufficient to attain a 200km orbit

Total impulse for both STS SRBs was approximately 2,138MNs
Total impulse for both SLS SRBs is approximately 2,651MNs

Whether straight avg Isp (1.8795MNs) or 90% of vac Isp (1.8819MNs) is used for the RS-25's Isp, a 520s nominal burn time results in 2,932MNs or 2,936MNs total impulse. All together, STS total impulse was 5,070MNs to 5,074MNs of total impulse for all engines and solid motors. All the noise was provided by 1,003,492kg (585.127m^3) of APCP solid propellant for both motors, plus 735,601kg (2,050.798m^3) of LOX and LH2 to feed the orbiter's trio of RS-25s. 42.17% of the total impulse was provided by solids, 57.83% by liquids.

Say I have a 1MN rocket engine that burns 301.5kg/s, so its avg Isp is 338.2s (311s sl to 365s vac). If this sounds suspiciously like a staged combustion variant of a Merlin-1D sized engine, then that's because it's exactly what it looks like. I can tweak Isp downwards to 316s and it doesn't really matter for purposes of this comparison.

301.5kg/s (per 1MNs) * 5,070MNs = 1,528,605kg of densified LOX/RP1 propellant consumed to supply 5,070MNs of total impulse. Obviously we will have multiple engines to get off the ground, but total impulse determines total force, and total force determines what you can push into orbit, with all the usual caveats about weight and thrust.

1,528,605kg / 1,224kg/m^3 = 1,248.860m^3 of densified LOX/RP1 propellant (1,224kg/m^3)

Now, then... Where were we?

Oh, right. The 1.5 stage to orbit Space Shuttle with its 1,739,093kg of propellant mass and 2,635.925m^3 of propellant volume provided the same total impulse as a notional SSTO with 1,387.065m^3 LESS propellant volume and 210,488kg LESS propellant mass. If SSTO is "bad" from a payload performance perspective, then solids plus LH2 is really bad. Pure LH2 is worse. To launch sizable payloads using pure LH2, you end up with giant paper-thin vehicle structures.

Total impulse is precisely the same as the Space Shuttle. If the Space Shuttle was viable, then so is a true SSTO with a lot less propellant mass and volume. Acceleration and therefore gravity losses won't be the same since RP1 provides a lot more thrust to work with.

Everything SSTO that is actually viable looks a lot like LOX/RP1, it's made from high tensile strength CFRP composites, and it's powered by 200:1 thrust-to-weight ratio RP1 burning engines.

Is it possible to do this some other way?

Maybe. I've made this as simple as I know how to make it. You cannot obtain better performance using LH2 if dry mass fraction matters at all (and it really does matter), in much the same way that Titanium is an absurdly poor substitute for a true high strength steel. There's a reason nobody makes landing gear for airliners out of Titanium, and it probably has a lot to do with strength-to-weight ratio and not fracturing like glass if you nick the metal. Titanium and LH2 are both "wow factor" materials with basic physics firmly set against them. You use them when you want mediocre performance results.

Space Shuttle pushed 164,036kg into a 200km orbit using more than double a notional SSTO's propellant volume and 210,488kg more propellant mass to do it. Space Shuttle's total dry mass was abhorrent as a result. That's how bad a LOX/LH2 plus solids alternative is to a true SSTO. That is why politicians shouldn't be in charge of selecting fuels or booster technologies for spacecraft. Stubbornness prevents people who should know better from actually doing better.

kbd512 · 2024-05-28 08:22:52

GW,

If we had a version of the Space Shuttle that was a LOX/RP1 powered and a true SSTO, with a dry mass fraction that was 5% of its propellant mass, it weighs 76,430kg. If it was capable of delivering the same 27,500kg payload as the historical Space Shuttle, the total mass to a 200km circular orbit in LEO is 103,930kg. That's still 60,106kg less than the historical Space Shuttle, which pushed the same 27,500kg to a 200km orbit, using the exact same total impulse. I think means we have another 60,106kg of payload performance, or we have another 60,106kg of mass to devote to the vehicle itself, so 136,536kg of total vehicle mass, and total liftoff mass is 1,692,641kg.

SSTO Space Shuttle Masses
Payload: 27,500kg
Vehicle: 136,536kg
Propellant: 1,528,605kg
Gross Liftoff: 1,692,641kg

Proposed SSTO vs Historical 1.5 Stage Vehicle and Payload Mass Percentages of Gross Liftoff Mass
Proposed SSTO Space Shuttle Vehicle: 8.1%
Proposed SSTO Space Shuttle Payload: 1.625%
Historical 1.5 stage Space Shuttle Vehicle: 5.42%
Historical 1.5 stage Space Shuttle Payload: 1.355%

GW Johnson · 2024-05-28 12:13:07

We in recent decades have built all-expendable stages with all-up inert mass fractions in the 4-5% range. That was not true early in the space age, but it has been for a couple of decades now. The key word there is "all-expendable". No heat protection, no means of landing.

SpaceX's Falcon cores with the landing legs are running in the 6-7% inert range. These are NEITHER SSTO's NOR do they have any real heat protection, seeing as how they hit the atmosphere about 8-10 times slower speeds than an entry vehicle.

I honestly do NOT believe that you can have a qualified entry vehicle also capable of landing, that does not have an inert fraction of 10% or more, probably significantly more, even to get a handful of trips, much less a long service life. And invoking "magic" materials that have yet to fly in space is NOT the answer here. Maybe a part of it, but not the whole answer.

SpaceX's "Starship" prototypes (which still bear little resemblance to any operational concepts!!!) are running about 120 tons inert and 1320 tons fully loaded but with ZERO payload. That's about 9% inert. And it has yet to survive entry even once! Although hopefully that will soon change. But, even if it does successfully enter and land successfully in an upcoming flight, that is one hell of a long way from an operational capability, and from a long service life. Light years away.

Plus, "Starship" is not an SSTO, and it is not going to be an SSTO. It is a reusable upper stage for a TSTO. It is NOT designed to provide 9.3 km/s with any useful payload at all. Something closer to 7 km/s, yes. But that difference makes a huge difference in the mass ratio and propellant fraction!

The requirement that bites you first for a reusable SSTO is the rocket equation. The dV requirement (and your Isp in the guise of Vex) determine the mass ratio your stage MUST have! Period! That mass ratio is essentially the propellant mass fraction = 1 - 1/MR. 1 minus that propellant mass fraction is the sum of your inert and payload masses.

And THAT "allowance" is all you have to work with! Period! End of issue! Density impulses and "magic" materials notwithstanding!

If you cannot meet the "allowance" left over from the rocket equation, then none of the rest of that stuff means a damn! Sorry, but that's just the ugly truth of it.

For an ascent-averaged Isp of 450 sec with 9.3 km/s demanded, I show an 88% propellant fraction, leaving 12% to cover inert + payload. (At 400 s, it is 10%. At 360 s, it is 7%. Run the damned numbers for yourself!!!)

If you really believe you can build a qualified entry vehicle that includes a landing capability for only (!!!) 10% inert fraction, then you can have 2% payload at 450 s Isp, but that's if, and ONLY if, you do everything else "just right".

Myself, I think 10% inert in a qualified reentry and landing vehicle is utter BS. I probably always will.

But you can have that 12% allowance with LOX-LH2 at 450 s Isp, and with a 10% inert, carry 2% payload, as a reusable SSTO. If you believe in 10% inert for a reusable entry-capable vehicle (and I clearly do NOT). Which utterly rules out LOX-LCH4 at 360 s (zero payload even at 10% inert), and most certainly rules out LOX-RP1 at 330 s (7% inert at zero payload).

I'll believe it when I see it! But not one minute before!

GW

Last edited by GW Johnson (2024-05-28 12:24:36)

tahanson43206 · 2024-05-29 11:20:07

For GW Johnson...

Just FYI ... here is what ChatGPT4o "saw" when I uploaded your Word document...

Weight statement, metric tons GWJ 5-29-2024
Payload 13
Inert 0 (ridiculous!)
Burnout 13
Propellant 200
Ignition 213
Mass ratio = ign/bo = 213/13 = 16.3846, ln(MR) = 2.7963
Propulsion models (Vex = 9.80667*Isp/1000)
Isp, s Vex, km/s
350 3.432
400 3.923
450 4.413
Rocket equation prediction at indicated mass ratio
dV = Vex ln(MR):
Isp, s dV, km/s
350 9.597
400 10.970
450 12.340
Bear in mind these are ridiculous predictions:
It is NOT possible to build anything that has zero inert mass!
Look at the problem the other way around:
Mission (LEO eastward at low inclination, rendezvous, deorbit)
Required dV capability ~ 9.3 km/s
Propulsion
Isp, s Vex, km/s (dV/Vex)
350 3.432 2.70979
400 3.923 2.37063
450 4.413 2.10741
Rocket equation (reversed as MR = exp(dV/Vex))
Isp, s MR
350 15.022
400 10.706
450 8.227
Weight statement-related:
Propellant mass fraction Wp/Wig = 1 – 1/MR
“allowance” = sum of payload and inert fractions = 1 – propellant mass fraction
Isp, s Wp/Wig allowance (= Wpay/Wig + Winert/Wig)
350 0.9334 0.0666 (everything that is “not propellant”)
400 0.9066 0.0934
450 0.8784 0.1216
Design considerations: the “allowance” is your “design space”
Whatever you think your inert fraction is, subtract that from the allowance for payload fraction
Size the weight statement masses from the fractions and the payload you want to deliver
Size the thrust requirement using T/Wig = 1.5+
Size inert masses for engines, tanks, and equipment, then iterate until it matches inert fractions

tahanson43206 · 2024-05-29 11:35:32

For GW Johnson re conversion of word format to bbcode

AISE was unable to handle the content of this post.
Meanwhile, the post went into the clone with no hiccups
This is one more confirmation that installation of the new updated software will be helpful

Here is a link to a post with links to a set of documents and a spreadsheet about the SSTO option(s)

https://newmars.com/forums/viewtopic.ph … 63#p223663

(th)

tahanson43206 · 2024-05-30 07:22:21

For GW Johnson re kbd512 concept for a launch vehicle ...

Recently kbd512 published a study he had created using a new set of numbers.

ChatGPT4o has evaluated the work and reported that it seems to hold up.

Here is a link to the numbers: https://newmars.com/forums/viewtopic.ph … 91#p223791

I'm hoping your spreadsheet system will match the output of the hand calculations.

My goal here is to try arrive at a common set of numbers everyone can agree upon.

If we can get everyone to agree upon one set of numbers, then we can focus intently on design of the vehicle to match the vision.

AISE has blocked my attempts to add the actual numbers to this post, or they would be here.

Meanwhile, the clone accepted the post without a quibble or a hiccup. This is one more indication that we will benefit from installation of the updated software as soon as that becomes possible.

(th)

tahanson43206 · 2024-05-31 20:14:36

GW Johnson provides a spreadsheet to help NewMars members who are studying SSTO vehicles...

https://www.dropbox.com/scl/fi/f566nkvz … n1cwt&dl=0

This line is where I'm planning to place a link to the image showing results.

(th)

GW Johnson · 2024-06-02 10:51:25

That spreadsheet is named "comparisons". It has multiple worksheets. The one you want for doing what is in post 310 is "rev roc eq", for reversed rocket equation.

GW

GW Johnson · 2024-06-02 10:52:28

I’m unsure where to post this. It and some of the preceding posts may well belong somewhere in the new SSTO topic.

Composite materials are tricky, and the advertising hype about them is very, very misleading.

The advertised high strengths are always ONLY in the fiber lay direction! Strength in the other directions is very much lower, and can be near-zero out of the fiber plane, depending essentially only upon the matrix properties. You have to “play God” and know all the necessary fiber lay directions to handle all the loads your structure might ever encounter. The odds of actually doing that successfully, first time up in a new design, are actually quite low.

These composite materials have strength only in the direction of the fiber lay, and even then ONLY so long as the matrix holds them in place! If the matrix fails for any reason at all, your material essentially turns into loose rags or threads or yarns. You can get strength in two directions, for some variable-with-angle level of strength in-plane, by using real woven cloth as your fiber.

The strength between two layers of fiber laid in different directions, or two layers of woven cloth oriented differently, is ONLY that of the matrix gluing them together, which is usually down in the 5-10 ksi range, and even then usually only by shear! If your stresses are 3-D, the material will fail at very low stress levels, in precisely the direction where fibers do not lay. It is inherent.

There is no practical way around that, other than fully-3-D fiber preforms, the manufacture of which is extremely difficult and horribly expensive. And the infiltration of those 3-D preforms with a matrix is also extremely difficult and horribly expensive! This is usually only done with metal matrix composites.

The tightly-woven tapes in particular are very strong in the long direction, when used in composite work. But if you try to use that by itself, in multiple layers of spiral wrap in alternating directions to make a cylindrical vessel, the only thing connecting one strip of tape to the next is matrix shear, with a low shear-bearing area from one tape strip to the next tape strip! Pressure test results of the vessel will be extremely disappointing! You might have considerable hoop strength, but you will inherently have almost zero longitudinal strength, and you will have essentially no bending strength as a tubular bending-loads-resisting structure (as in a cylindrical tank that is also an airframe component). These are best made with a few layers of woven fabrics, or many, many layers of yarn spin-wrapped in multiple different directions.

You cannot use composites, especially woven composites, in arbitrarily thin layers! The thinnest thickness of a one-layer woven cloth composite is the thickness of that cloth layer. Period. Set by the threads or yarns from which it was woven, but a bit larger due to the weave geometry. You are extremely unlikely to get the strengths you need, unless there are actually multiple layers with different fiber lay directions! This stuff simply cannot be arbitrarily thin! Formable metals can be arbitrarily thin, except that alpha-phase titanium cannot, only beta phase titanium can be formed into thin sheet (which ages at room temperature into uselessness, with big, weakly-bound grains).

The ceramic and carbon fibers that are now popular are indeed “good” for substantial exposure temperatures, by themselves. Generally speaking, the organic polymer matrices you must use with them are not! Most of the hydrocarbon-type polymers start charring at around 300-400 F, and even the silicones start charring at about 600 F.

Everything organic or silicone is fully carbonized by no more than 1000 F, into something soft and crumbly, resembling the charcoal in your BBQ grill, except that it may still be reinforced by the fibers, but only if those fibers were ceramic or carbon. That is not a particularly reusable material for heat exposures, but if the charcoal “hangs together” better with the fiber reinforcement, the hype can truthfully claim their material “survived” the exposure, as long as NOT A WORD was said about ever using it again!

None of the polymers, or other plastics that I have ever heard of, have more than utterly-trivial elongation capability soaked out cryogenically cold (way under 1%). Using such materials in composite cryogenic propellant tanks thus presents two extremely serious problems: (1) very brittle behavior conferring extreme fragility in any kind of handling, and (2) leakage of propellant through the inherent porosity of the composite, unless you can successfully add some sort of internal sealing layer (not a trivial exercise in and of itself).

Taken together, those cryo-cold fragility troubles, the weight of multiple fabric layers to get strength in more than one direction, and the inability to sustain high temperature exposures, plus the manufacturing costs, are EXACTLY why SpaceX abandoned its original composites-construction concept for its Starship, in favor of simply welding-up 300-series stainless steel sheets!

Not having to put a heat shield on the lee side surfaces with stainless steel saves a great deal of heat shield weight, and also saves some rather large manufacturing costs, which made stainless steel the far better choice for a stage that must also qualify as a re-entry vehicle from orbit, plus endure descent and landing loads. You can take 304/304L SS to 1200 F as many times as you like, although it will start corroding and scaling if you let it get any hotter than that. Nothing re-radiates efficiently enough for self-cooling until it reaches temperatures in that 1000-1200 F class. Organic-matrix composites never will survive that exposure, and still be usable a second time.

There is also the issue of stiffness, which is partly the moment of inertia of the geometry, and partly the modulus of elasticity of the material. Steels have a modulus of elasticity in the 30 million psi range. Aluminum is about 10 million psi. Carbon-epoxy gets close to aluminum, and the other things like glass-vinyl ester (or glass-polyester) are far less stiff than that. Even Kevlar-vinyl ester, while quite tough against impacts and punctures, is far less stiff. Depending upon what the structure has to do, the achievable stiffness may (or may not) be crucial. But carbon-epoxy (and the others I mentioned) is limited to under about 300 F exposure, if the epoxy matrix is to survive in a reusable condition!

And then there is detection of impact damage, to which any composite is critically vulnerable. Unless of catastrophic magnitude, it’s often hidden damage, you see nothing on the surface. Ignoring the advertising hype otherwise, the only way known to reliably detect hidden damage is by comparing before and after x-ray images, looking for differences. There has to be one taken before the impact, and the other after, for this to work.

Every square inch of the composite has to be in an as-built x-ray somewhere, so that after a suspected impact, a post-impact x-ray can be taken for direct comparison to the as-built x-ray. The advertising hype is wrong, there are NO general criteria for evaluating post-impact x-rays only! Nothing is reliable enough to permit that! These materials are simply too variable from one square inch to another, to permit that sort of thing. This requirement for as-built x-rays during manufacture acts to further drive-up costs.

There is also the question of joints. Everyone is familiar with the way the fiberglass fenders on a Corvette eventually tear out around the machine screws holding them in place. You simply cannot attach composite parts to other structures that way, it is not reliable, and it won’t last, especially if highly-loaded. Fiberglass or something else, it does not matter! Composites and conventional fasteners are simply incompatible.

What is required is a hybrid of alternating layers of the composite and shim-stock thin metal, with enough layer bond area to carry the loads by only matrix shear, from the metal layers holding the fasteners, into the composite layers that extend into the rest of the composite structure. There must be enough layers of the metal shim material to take the fastener loads as bearing loads without failing, which in turn means there must also be many layers of composite, in the region of the joint.

That’s very labor intensive (and therefore expensive) to accomplish, which is EXACTLY why I do not trust Boeing to have done that job correctly on the B-787. Neither do I trust Airbus to have done this job correctly, not after that Airbus lost its composite vertical fin and crashed in NYC many years ago. The fin tore out around the bolts holding it to the aluminum airframe (it was not a hybrid joint, surprise, surprise). But this hybrid technique worked like a charm with composite solid rocket motor cases and metal end closures, up to 4-5000 psi, as tested and well-proven.

If you get the idea that designing and building with these materials is difficult, and fraught with fatal pitfalls, then you understand the main point here. If you do this job wrong, you will achieve better weight reductions and more cost savings (although still more expensive than metal construction). However, the product WILL INEVITABLY FAIL unexpectedly in service, causing very serious and expensive legal issues of the “whose fault was this?” type. Such can be quite catastrophic.

On the other hand, if you do the design correctly, and avoid those unexpected failures in service, your product will cost even more, and it will save only a little weight, compared to equivalent metal construction. Top managers hate that. Rightly or wrongly (wrongly in my opinion), they tend to discount the legal troubles for unexpected failures in service, and will push you very hard to do the job wrong. (Which in turn is partly why the advertising hype lies so egregiously.)

I’m probably obsolete and out-of-date, regarding all the latest composite materials. But none of these older ones (or the new ones I am not familiar with) will have any different extreme directionality of their properties, or any different severe limitations on exposure temperatures for their organic matrices (including the silicones). And the metal matrix composites are still horribly expensive, and probably always will be. As well as exceedingly difficult to manufacture.

There are no “magic” materials!

GW

SpaceNut · 2024-06-02 11:31:40

This current forum does not allow me to split and merge a post to the correct topic, but I can quote, and copy paste it to what I believe is the correct one.

GW Johnson · 2024-06-02 15:56:59

I see that you did exactly that. Thanks, Harold.

-- GW

tahanson43206 · 2024-06-02 19:14:31

For gW ... the Zoom meeting is open ... use the link at the top of the Zoom topic

tahanson43206 · 2024-06-09 07:53:27

For GW Johnson...

Void mentions your work in this post: http://newmars.com/forums/viewtopic.php … 20#p224320

The video is about Blue Origin patent applications with a focus on aerospike engine design.

(th)

GW Johnson · 2024-06-09 10:14:25

I went and saw, and thanked Void there. -- GW

tahanson43206 · 2024-06-11 08:35:54

For GW Johnson re Video "direct landings of big things on Mars"

At 29 minutes into the video, the slide title is: Easy Access to Unload.

Your proposal is to design the exterior wall of the landing vehicle so it can swing down for use as a ramp.

In the case of the SpaceX vehicles, the walls are designed to bear weight. I'm assuming the doors will not bear weight along the longitudinal axis of the ship. Can you confirm that impression? I'm assuming the walls of the vehicle were doors/ramps are not present would bear all the load. For your design there may not be much load to worry about, but this is a detail that may be of interest to engineers or those with an interest in engineering.

However, my reason for creating this post is to offer a suggestion for your ramp idea....

The gravity is lower on Mars, but those ramps need to be as light weight as possible, since they are serving no purpose other than as ramps.

Thus, the components of the payload that weight a ton on Mars are going to be directing that force as they move along the ramps. It seems likely to me that the ramps will be unable to withstand the load, in light of the dead mass you would have to allocate to them.

I have a possible solution for your consideration.

The regolith itself can be collected under the ramp to provide support when the heavy payload items are brought out. This could be done using robotic equipment at the landing site, if it is available.

An alternative is to run cable from the nose of the ship to anchors in line with the ramps, so that part of the force exerted by gravity can be borne by the cable. This system would require more dead weight on the ship, but it might be prepositioned, so that it is simply called up when the ship lands.

For those NewMars readers (as well as members) who may be interested, the video is part of the series at exRocketman1 at YouTube.

(th)

GW Johnson · 2024-06-11 09:04:10

The unload ramps are cylindrically-curved panels, 4 of some 8 up each aside. These are conceived as metal shells exposed to the environment, covered internally with a layer of something like mineral wool insulation, and a thin layer of plastic sheet to cover up and hold the insulation in place. There might be a couple of stringers that add extra bending strength, just to construct the things.

The curved shape offers quite a bit of extra stiffness in bending compared to a flat panel. However, since this is a one-shot lander, it does not matter if the panels get bent a bit during unload operations. Their real value after unload is their salvage value. Both the metal and the insulation have reuse value. The plastic sheet, not so much. Maybe as dust door mats or something.

GW

tahanson43206 · 2024-06-11 17:58:06

For GW Johnson ... two items this time...

First, thanks for your reminder that the lander would be a one-and-done, so the ramps would be expendable.

Second, kbd512 just posted a ** really ** interesting idea in the Vertical Launch Assist topic ...

I am wondering if you think a ram jet might work at the low velocity coming out of the Launch Assist?

I ran a quick check with omnicalculator.com, and found a velocity of 343 meters/second as the velocity of an object accelerated at 5 G for 7 seconds, which is about a kilometer.

I'll be interested to see what might be done with this idea.

****
We're in the GW Johnson topic, so we are not constrained by the topic title ...

Do you think kbd512's idea, of using atmosphere oxygen to help a hybrid SSTO to reach orbit might work?

This might be combined with the inclined track he and Calliban have been discussing...

A cargo carrier might be accelerated to some velocity on an inclined track, and it might use a ram jet to pass through the atmosphere where that makes sense, and then use an ordinary rocket to reach orbit?

This would be a winged vehicle so it can return to Earth for reuse.

(th)

GW Johnson · 2024-06-26 09:21:30

From AIAA’s “Daily Launch” for Wednesday 6-26-2024:

About Starliner docked to ISS ---

THE WASHINGTON POST
Astronauts’ delayed return reflects high stakes for Boeing and NASA

Before Boeing’s first flight with humans on its Starliner spacecraft earlier this month, the company and NASA said repeatedly that a rigorous testing program following years of delays and costly setbacks meant it was finally ready to fly astronauts. Instead of coming home after about eight days, the spacecraft remains docked to the station, its return delayed indefinitely while teams continue to troubleshoot a series of problems in the capsule’s propulsion system.

About NASA’s new spacesuits ---

ARS TECHNICA
NASA’s commercial spacesuit program just hit a major snag

Two years ago NASA chose a pair of private companies to design and develop new spacesuits. Now, that plan appears to be in trouble, with one of the spacesuit providers—Collins Aerospace—expected to back out,

And ---

SPACE
ISS astronauts conduct 'spacewalk review' after spacesuit coolant leak

ISS astronauts are reviewing spacesuits and spacewalking procedures after a leak during a spacewalk on June 24. NASA's next spacewalk is still scheduled...

About Falcon-Heavy being successful yet again ---

SPACENEWS
Falcon Heavy launches GOES-U weather satellite

A SpaceX Falcon Heavy rocket lifted off June 25 carrying the final spacecraft in a series of geostationary weather satellites that also features several firsts. The rocket’s payload, the GOES-U weather satellite, successfully deployed from the Falcon Heavy’s second stage four and a half hours after liftoff, after the stage completed a sequence of three burns to place the satellite into a geostationary transfer orbit.

And, about the upcoming first launch of Ariane-6 ---

SPACE
Europe's new Ariane 6 rocket on track for long-awaited 1st launch on July 9

Following its wet-dress rehearsal on the launch pad in French Guiana, Europe's new Ariane 6 rocket is on schedule for its inaugural launch. The European Space Agency (ESA) announced during a June 25 press conference that the rehearsal was a "full success."

GW

Last edited by GW Johnson (2024-06-26 09:22:29)

tahanson43206 · 2024-06-26 18:08:44

For GW Johnson!

In the SpinLaunch topic, SpaceNut has provided links to reports on traditional solid fuel rockets. I have not taken the time to follow the links, because I am almost certain (98.9% sure) that the links will be about traditional solid fuel rockets, which would fail if spun in a SpinLaunch system. I have asked SpaceNut to keep looking, because he may find a report of someone developing a solid fuel rocket that can survive 10,000 G's of lateral force before ignition.

However, since SpaceNut does not appear to be convinced by your assertion that solid fuel rocket material will flow like water at 10,000 G's, I'm wondering if you might recall where you posted the information we need. If you are willing to accept a search tag for your post, we can add one to help future readers who might want to read the post for themselves.

An alternative is for you to repeat your post, although I recognize that is an imposition on you and a waste of your limited time. The problem at hand is .... we did not tag your original post, because (I for sure) did not realize it's importance.

A problem I am struggling with is how to tag your information. This is a report from your experience, showing that solid fuel rocket engine material flows like water at G forces far less than 10,000.

So far, I have seen no evidence that NewMars readers other than one moderator have seen your presentation on this subject, but it is critical that engine designers for SpinLaunch must accept the reality you have observed, so that success might be achieved.

(th)

tahanson43206 · 2024-06-28 08:20:02

For GW Johnson!

In the SpinLaunch topic, SpaceNut has provided a very interesting insight. He observes that if the rocket motor for the SpinLaunch application is designed with the nozzle in the center, then any fuel that is above the opening of the nozzle will flow out the nozzle.

There are two possible solutions that I can think of, and each requires your evaluation:

1) the nozzle might be fitted with a metal plug that can hold the grain inside the motor during acceleration.
2) the nozzle might be set at the top of the cylinder where the air cavity will form as the fuel flows into the cylinder shape under acceleration.

Please comment upon this two ideas, and add additional insights.

I see NewMars >> the Mars Society as in direct competition with any other organizations on Earth that are interested in this problem.

(th)

GW Johnson · 2024-06-28 15:21:55

In any solid propellant motor, there must be a free volume in which to collect the gas coming off the burning surface, and then funnel it into the nozzle entrance. This same free volume affects c* efficiency (the "combustion efficiency"), because the stuff coming off the surface is not yet fully mixed or fully burnt. It's a residence time thing.

As for extreme gee, Spacenut is exactly right, if there is a path for the solid propellant to flow (like a liquid) out the nozzle, then it will flow out the nozzle under extreme acceleration. There is no such thing as a true solid. Under extreme temperature and extreme applied force, all materials flow like liquids. Only the detailed numbers vary from material to material.

As for asymmetry, that affects the moments applied, for forces off axis. This can be quite the severe effect.

GW

Last edited by GW Johnson (2024-06-28 15:23:25)

tahanson43206 · 2024-07-01 15:58:04

For GW Johnson re #324

Thank you for this further clarification of the limitations of solid fuel for the SpinLaunch application.

SpaceNut attended the most recent NewMars Google Meeting (via Smartphone) and he heard the discussion about the SpinLaunch problem of achieving LEO. SpaceNut then provided a link to video that showed that the orbital version of the system will be canted toward the East as kbd512 had recommended. Equally importantly, the video that SpaceNut found appears to show that the rockets for stage 2 and 3 will be liquid and not solid as I had assumed.

Given the recent discussion of hypergolic propellants in other topics, please evaluate the potential of hypergolic propellants for this application. My guess is that these propellants will reside in spherical containers (as shown in the video) and that they will not have a problem withstanding 10,000 G's.

Given that, do you think these propellants can deliver the payload to LEO, if the launch angle is canted toward the East?

(th)

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