Orbital space is 25 times harder than suborbital

RGClark

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I am impressed that you went to so much effort to reply to a well known crank; you deserve credit for that effort, even if it is wasted on Smilin' Bob.

I'd like--if possible--to move this discussion toward actual mass fractions as achieved by amateur builders:

My six inch by 60" stages are coming in at just below 60% propellant (using a not fully aluminum loaded propellant). My 9 inch by 84" stage--for which I have actual weights--will come in at a slightly lower propellant fraction than the 6 inch motor (using the same propellant) because the required wall thickness jumps from 0.125" to 0.250". In both cases, the aluminum tube is the single largest contribution to dry mass. I have some parts (tubes, forward bulkheads) for a 12" by 120" motor; using CAD based weight estimates for the not yet built bits, I am showing slightly over 60% propellant fraction for that stage, largely because the tube wall remains at 0.250".

I've looked at both Titanium (Grade 5) tubing and carbon fiber. The former--for the six inch stage--will bring the propellant fraction up to just over 61%; however, that tubing in the required wall thickness in not available. Flow forming that tubing has an about $10K setup cost and then the cost of material and a bit more per tube; for the half dozen tubes that I probably would never get through, that appears to work out to about $1500 per tube...about 10x the cost of the current aluminum tube. In addition, the thin wall on the Titanium tubing would require going to button head fasteners rather than the current countersunk, meaning drag would be significantly higher.

A friend has privately and without any commitment suggested that he could make a carbon fiber tube in 6 inch by 60" dimension including an aerospace grade propellant liner for about $3000 per unit. That would get propellant fraction up to about 63% if no other piece parts changed. As with the Titanium tubing, the much thinner wall would present problems with fasteners; in this case that would most likely be met by using thicker walls at the required locations, those thicker walls would reduce the propellant fraction back toward 61%.

Can others comment on what sort of actual propellant fractions they are seeing in built stages?

Bill

Carbon fiber reportedly cuts the casing weight in half over standard aluminum. If the casing weight, is the greatest bulk of the dry mass why is that only improving the propellant fraction from 60% to 63%? That means the dry mass fraction is only reducing from 40% to 37%. Perhaps you can give a breakdown of the components weights of the dry mass of the rocket to see why the improvement is so minimal?

By the way there are other metals besides titanium that can improve the propellant fraction, some even better than carbon fiber:


Bob Clark
 
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RGClark

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My Hamster rockets come out to be right at 66% propellant mass of the complete rocket.
Should be able to push that over 70% for next year.

Nothin fancy or expensive used.

I’m curious. Just looking at the ratio of propellant weight to casing weight, what would that be? Remember for this application none of the stages would have recovery systems.

Bob Clark
 

wclaybaugh2

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YES - make the pressure vessel with the propellant already inside.
Embrace single-use.

Fred:

I admire your bravery in overwrapping cast propellant…not something I would do even for very small grains.

I did do my homework and now understand how you are getting the propellant fractions you report.

Thanks,

Bill
 

FredA

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bravery in overwrapping cast propellant…
I don't see it as brave at all - just logical.
APCP is pretty hard to light we don't do anything that would approach ignition temps.
I do NOT heat-cure - that I consider a LITTLE risky but really heat-cure temps are still only about half of ignition temp so good guard band on that should we go there in the future.

Nice thing is that rockets built this way cost next to nothing....and weigh next to nothing.
 

RGClark

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Odd, there have been many amateurs who have constructed their own carbon fiber casings. But I have been unable to find any who have given what propellant fractions they have achieved this way. That should be easy to get, just give the weight of the propellant grains without the casing then the weight with the casing. But not even in the case of the well-publicized flight to suborbital space by the USC RPL team using Traveler IV is this number given.

This video shows the launch attempts by the USC team of Traveler III and Traveler IV:



At about the 7:50 point it shows the USC team measuring the weight of the empty carbon fiber casing. It looks like it might be showing "36" on the scale in whatever units the scale is in but the video cuts away before we can see if that is the stable measurement, as it is fluctuating. I'm inclined to think the scale is in kilograms since several young men are holding it up to the scale, which wouldn't be so difficult if it was only 36 pounds.

The video also helps us to estimate the volume, and so mass, of propellant since knowing the diameter is 8" we can estimate the casing height.

Bob Clark
 
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LithosphereRocketry

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Carbon fiber reportedly cuts the casing weight in half over standard aluminum. If the casing weight, is the greatest bulk of the dry mass why is that only improving the propellant fraction from 60% to 63%? That means the dry mass fraction is only reducing from 40% to 37%. Perhaps you can give a breakdown of the components weights of the dry mass of the rocket to see why the improvement is so minimal?

By the way there are other metals besides titanium that can improve the propellant fraction, some even better than carbon fiber:


Bob Clark
I don't think that's been "reported" by anyone except yourself.
 

RGClark

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I don't think that's been "reported" by anyone except yourself.

This member of the forum got to 70% propellant fraction with a 2” motor and believes 75% should be doable with a larger motor:

Even 75% would be adequate for a staged orbital rocket composed of solid motors.

And this amateur got to 87%:

EDDE35DE-C709-40A7-9260-27CF6162D4AB.png

And to quote a line from a well known movie,

“What one man can do, another man can do!”

Bob Clark
 

Antares JS

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So once you build that and successfully accomplish staging at high altitude, near vacuum, you’ve accomplished the most difficult parts of reaching orbit.
This is so unbelievably wrong it's hard to even know where to start.

Okay, you put a lot of emphasis on having a high mass fraction. Having a high mass fraction means nothing if your rocket does not impart enough energy, and making a motor large enough to impart enough energy is expensive and difficult, especially if you are an amateur.

The energy requirement to cross the Karman line is not even CLOSE to the requirement to reach orbit. Not only do you have to ascend to a point where atmospheric drag isn't going to bring your would-be satellite back down, you have to boost it horizontally to 17,500 mph. You say "once you have a first stage and a good separation system, you've done the hardest part," while underestimating how hard it is to build a first stage capable of delivering an orbital-class upper stage to the needed altitude and orbit. And that's not even the hardest part because...

Guidance. You say "oh we don't need guidance unless we are going to a specific orbit" and this is also wrong. Unless you are towing a massive amount of extra propellant/energy to make up for any imprecision, which is unlikely, your vehicle needs to follow a very precise course that spends its energy in precisely the amount and direction needed to reach a stable orbit. Failing at this means you crash before completing a single orbit. Active guidance is a must, especially on your upper stage, and you have by no means accomplished the hard part once you have a first stage booster.

I work in the space launch industry and am very intimate with what it takes to reach orbit. Your belief that reaching orbit is easy after you manage to reach the karman line is utterly false. I'm not saying no amateur will ever do it, but it's not going to happen anytime soon and it's far harder than you seem to think.
 

kramer714

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I'm really interested in carbon casings because of the possibilities they offer for better integrating the motor into the rocket. Instead of having a separate fin can, you could seamlessly integrate composite fins into the motor casing, with no step in the OD from tip to tip. You can also have a tailcone integrated into the motor casing and nozzle area, and run the fins down onto that for a drag reduction.
um, nope.

If the tube is optimized for a perfect pressure vessel straight tube (lets ignore end domes for now) adding fins with fibers not optimized for a pressure vessel and not in the vector of the pressure loads would locally weaken the structure. You can compensate for that by adding thickness - doable but not without a weight and drag penalty.

Integrating tail cones and end domes also comes with a penalty. Hoop wraps (near 90 degree wraps) can not go down slopes like tail cones or end domes. Additional material needs to be added off axis to make up for that. Another problem is the fiber angles change as you go down a slope, again can be accounted for but everything has a price. If you look at most composite rocket motors there is usually a skirt attached over the ends with an increased OD. The hardest part of a composite pressure vessel is the transition from the cylinder into the domes. This is where we spent the most time designing, and typically where we would see a failure.

Other issues are how dynamic a pressure vessel is, they grow in length and diameter when loaded Things like fins and skirts can actually cause stress concentrations that result in local failures. All needs to be factored in.

Optimization is about dealing with what you have in the 'least bad ways' it is very difficult to have a structure, even one as simple as a pressure vessel, fully using the material properties everywhere. It is just a matter of how best can you make them work.

Not trying to complicated it, but simple to make it work, much harder to make it work at minimal weight / cost.

Mike (actual composite engineer) K


1669846002799.png
 

kramer714

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um, nope.

If the tube is optimized for a perfect pressure vessel straight tube (lets ignore end domes for now) adding fins with fibers not optimized for a pressure vessel and not in the vector of the pressure loads would locally weaken the structure. You can compensate for that by adding thickness - doable but not without a weight and drag penalty.

Integrating tail cones and end domes also comes with a penalty. Hoop wraps (near 90 degree wraps) can not go down slopes like tail cones or end domes. Additional material needs to be added off axis to make up for that. Another problem is the fiber angles change as you go down a slope, again can be accounted for but everything has a price. If you look at most composite rocket motors there is usually a skirt attached over the ends with an increased OD. The hardest part of a composite pressure vessel is the transition from the cylinder into the domes. This is where we spent the most time designing, and typically where we would see a failure.

Other issues are how dynamic a pressure vessel is, they grow in length and diameter when loaded Things like fins and skirts can actually cause stress concentrations that result in local failures. All needs to be factored in.

Optimization is about dealing with what you have in the 'least bad ways' it is very difficult to have a structure, even one as simple as a pressure vessel, fully using the material properties everywhere. It is just a matter of how best can you make them work.

Not trying to complicated it, but simple to make it work, much harder to make it work at minimal weight / cost.

Mike (actual composite engineer) K


View attachment 548471
Oh and fiber angles, For an 'optimized' motor tube, ignoring the end domes, you would want as a minimum two wind angles. A high angle (nearly hoop) and a low angle. The angle used for the low angle in a motor case with integral domes has a lot to do with the ratio of the port diameter (hole at the end) and the tube diameter. Again, the optimization is the least worse compromise.

Mike K
 

plugger

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um, nope.

If the tube is optimized for a perfect pressure vessel straight tube (lets ignore end domes for now) adding fins with fibers not optimized for a pressure vessel and not in the vector of the pressure loads would locally weaken the structure. You can compensate for that by adding thickness - doable but not without a weight and drag penalty.

Integrating tail cones and end domes also comes with a penalty. Hoop wraps (near 90 degree wraps) can not go down slopes like tail cones or end domes. Additional material needs to be added off axis to make up for that. Another problem is the fiber angles change as you go down a slope, again can be accounted for but everything has a price. If you look at most composite rocket motors there is usually a skirt attached over the ends with an increased OD. The hardest part of a composite pressure vessel is the transition from the cylinder into the domes. This is where we spent the most time designing, and typically where we would see a failure.

Other issues are how dynamic a pressure vessel is, they grow in length and diameter when loaded Things like fins and skirts can actually cause stress concentrations that result in local failures. All needs to be factored in.

Optimization is about dealing with what you have in the 'least bad ways' it is very difficult to have a structure, even one as simple as a pressure vessel, fully using the material properties everywhere. It is just a matter of how best can you make them work.

Not trying to complicated it, but simple to make it work, much harder to make it work at minimal weight / cost.

Mike (actual composite engineer) K


View attachment 548471
Not to argue with anything you've said specifically here, but I'd point out that on page one I linked to Mike Passaretti's attempts to do this, which was successful (in terms of holding the N5800 together and surviving boost) in the first instance.

Page 1 of that thread starts here.
 

kramer714

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Not to argue with anything you've said specifically here, but I'd point out that on page one I linked to Mike Passaretti's attempts to do this, which was successful (in terms of holding the N5800 together and surviving boost) in the first instance.

Page 1 of that thread starts here.
Nice what Mike did, but not optimized. Wrapped cloth is ok for this but the ratio and orientation of the fibers isn't optimized. Rule of thumb for composite pressure vessels, you want twice as many fibers in the hoop direction as the axial direction. Cloth is 50/50. You can buy cloth with a bias to get closer.

The 2/1 ratio is why in aerospace you frequently see a metal 'bottle' with a hoop wrap of composite. Have the metal take all of the axial loads and half of the hoop stress, let the composite take the other half of the hoop stress. This simplifies sealing too, let the metal be the balloon in a composite 'birdcage'. Have the end fittings integral into the metal. Even if you have a straight tube, the smooth metal ID makes sealing easier than trying to seal to composite.

Typical construction is to have a straight tube and weld end fittings onto it. This gets over-wrapped with a composite only in the hoop direction. The end fittings thicken to handle the stresses where the hoop wrap ends.

I have used this method over the years for pressure vessels and hydraulic accumulators. Depending on what you need to do to the ends, this can be lighter than a composite tube with attached metal fittings.

Mike K
 
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kramer714

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Scaling and Mass Fraction

Some (well intend) bad assumptions people are making on scaling pressure vessels. It doesn't matter if they are composite or metal for this discussion.

There are a lot of factors on pressure vessel design and optimization that are part of the calculations. From a parametric standpoint, the Space Shuttle SRB are 900 psi, 12 foot in diameter and a wall thickness of .5 inch (steel). Just scaling that down to 98mm gives a ratio of 36/1 (12 foot / 4 inches). Using that ratio the wall thickness would be .014 thick for steel. Would a .014 wall thickness case work? (in reality, with manufacturing and materials variability, impact resistance in handling, etc... ) um - nope. Plus remember the SRB were designed with fatigue in mind for multiple launches and impact / handling loads during recovery. Plus safety factors for man rating. They are thicker than they would need to be for a single use non man rated.

Ok so, what about the AJ-260 22 foot diameter with a wall thickness of .6 inch (built using submarine welding methods!) , almost double the diameter of the SS-SRB but only 20% thicker...

These are extreme examples but scaling for a 54mm to a 98 or a 150 mm has the same type of effects. You cant just parametrically scale them. Big diametrs act like plates, small diamters act like arches from a stress standpoint. (really strain driven) even attachements dont scale, a few large bolts don't act like bunch of smaller ones.

Other problems with the way people are scaling here, if the goal was ONLY to optimize mass fraction, larger diameter win. but at what drag cost? What is the cost from an internal ballistics standpoint?

Last comment on this, fun to 'what if' by scaling, but seldom works once you put reality, engineering and physics to the problem the end result can change.. a lot.


Mike K
 

plugger

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Nice what Mike did, but not optimized. Wrapped cloth is ok for this but the ratio and orientation of the fibers isn't optimized. Rule of thumb for composite pressure vessels, you want twice as many fibers in the hoop direction as the axial direction. Cloth is 50/50. You can buy cloth with a bias to get closer.
Sure, it's not optimised in comparison to your OP, but it's certainly a major improvement when compared to COTS casings and their comparatively inferior mass fraction.
 

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I use a small plywood bulkhead/drogue-bay for my forward closure and an RCS nozzle for the aft.
Winding over each of these "just enough" to hold onto the mostly hoop-wound [loaded] casting tube.
Ends are already epoxied in place - winds over them just help secure them and on the nozzle-end pushes the eventual dog-bone down the nozzle shoulder for less aerodynamic impact.
 

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I'd like--if possible--to move this discussion toward actual mass fractions as achieved by amateur builders:

My six inch by 60" stages are coming in at just below 60% propellant (using a not fully aluminum loaded propellant). My 9 inch by 84" stage--for which I have actual weights--will come in at a slightly lower propellant fraction than the 6 inch motor (using the same propellant) because the required wall thickness jumps from 0.125" to 0.250". In both cases, the aluminum tube is the single largest contribution to dry mass. I have some parts (tubes, forward bulkheads) for a 12" by 120" motor; using CAD based weight estimates for the not yet built bits, I am showing slightly over 60% propellant fraction for that stage, largely because the tube wall remains at 0.250".......

Can others comment on what sort of actual propellant fractions they are seeing in built stages?

Bill
Bill, as far as motors themselves go, the 54/4000 M-1378 is at 59.9% propellant mass fraction, full up loaded with RTV, grease and all.
Case wall is .094' and liner .052"

A 98mm O motor I've been developing is currently at 62.4% and I hope to push that a few percent higher in the next iteration.
 

RGClark

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This is so unbelievably wrong it's hard to even know where to start.

Okay, you put a lot of emphasis on having a high mass fraction. Having a high mass fraction means nothing if your rocket does not impart enough energy, and making a motor large enough to impart enough energy is expensive and difficult, especially if you are an amateur.

The energy requirement to cross the Karman line is not even CLOSE to the requirement to reach orbit. Not only do you have to ascend to a point where atmospheric drag isn't going to bring your would-be satellite back down, you have to boost it horizontally to 17,500 mph. You say "once you have a first stage and a good separation system, you've done the hardest part," while underestimating how hard it is to build a first stage capable of delivering an orbital-class upper stage to the needed altitude and orbit. And that's not even the hardest part because...

Guidance. You say "oh we don't need guidance unless we are going to a specific orbit" and this is also wrong. Unless you are towing a massive amount of extra propellant/energy to make up for any imprecision, which is unlikely, your vehicle needs to follow a very precise course that spends its energy in precisely the amount and direction needed to reach a stable orbit. Failing at this means you crash before completing a single orbit. Active guidance is a must, especially on your upper stage, and you have by no means accomplished the hard part once you have a first stage booster.

I work in the space launch industry and am very intimate with what it takes to reach orbit. Your belief that reaching orbit is easy after you manage to reach the karman line is utterly false. I'm not saying no amateur will ever do it, but it's not going to happen anytime soon and it's far harder than you seem to think.

I have been impressed by the levels of accomplishment by amateurs also with the avionics which would be necessary for rockets reaching useful orbits.

For example, the BPS team has been able to gimble their engines so that spin stabilization is not needed and their rocket can even be landed:



Bob Clark
 

wclaybaugh2

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Bill, as far as motors themselves go, the 54/4000 M-1378 is at 59.9% propellant mass fraction, full up loaded with RTV, grease and all.
Case wall is .094' and liner .052"

A 98mm O motor I've been developing is currently at 62.4% and I hope to push that a few percent higher in the next iteration.

Scott:

That is certainly consistent with my experience if I strip out the fins from one of my motor / airframe assemblies.

I imagine some additional propellant fraction could be achieved by going to optimal Aluminum loading of the propellant but experience indicates that lowers Isp due to external burning because of low residence time for the last grain. If I were being compulsive about this I might try going to an optimal propellant mix for the top four grains and use my current lower aluminum content mix for the last grain but that means field mixing two different propellants for what appears to be an at most percent or two gain in propellant fraction. Ease of field operation wins on that trade, in my view.

Bill
 

kramer714

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Bill, as far as motors themselves go, the 54/4000 M-1378 is at 59.9% propellant mass fraction, full up loaded with RTV, grease and all.
Case wall is .094' and liner .052"

A 98mm O motor I've been developing is currently at 62.4% and I hope to push that a few percent higher in the next iteration.
Scott,

Roughly how much of the non propellant weight are the end fittings (forward closure / nozzle / nozzle retainers, the things that you would need to have if you changed from a metal tube to a composite tube. Reason im asking is, to help people analyze the potential mass fraction of a composite case vs a metal one. Assuming for a quick calculation that the weight of these items is the same (for an initial sizing) and the only weight change is the composite vs aluminum for the tube.

Mike K
 

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Next, you chose a vacuum isp of 285s for your motors. This is an unreasonable assumption. As stated previously, the O8000 doesn’t use a high isp propellant. I ran some simulations in Openmotor with a motor similar to the ones I was evaluating, and when changing the exit pressure to a vacuum, and enlarging the nozzle exit, I saw about a 20% increase in isp. This seems to match what I’ve been able to find for performance on motors used in industry. Here is the first part of my spreadsheet, with the isp and delta V calculated for each motor, at the stock hardware mass, half hardware mass, and a 0.8 mass fractio

I’m still learning to use OpenMotor. What do you get for the vacuum Isp for the motor used on the UP Aerospace SpaceLoft suborbital rocket. Specifications here:

EE21F6B3-1E51-4DFB-BBC2-72159B791175.jpeg

Thrust curve of the motor here:

Cesaroni Booster motor for UP Aerospace.
; @File: CTI_UPA-264-C.eng, @Pts-I: 42, @Pts-O: 31, @Sm: 0, @CO: 5%
; @TI: 438207.0, @TIa: 437890.0, @TIe: -0.05%, @ThMax: 52454.9, @ThAvg: 37051.4, @Tb: 11.827
; Exported using ThrustCurveTool, www.ThrustGear.com
S37029 265 3018 P 186.8801 242.672 CTI
0.037 540.773
0.061 1406.008
0.073 13086.69
0.086 48561.4
0.147 52454.9
0.257 50616.3
0.355 49859.2
0.698 48453.2
0.918 46073.8
1.175 44667.8
1.591 43586.3
2.338 42829.2
4.199 42288.4
4.701 41855.8
8.006 35907.3
9.658 32230.0
10.234 26606.0
10.368 23577.7
10.637 21414.6
11.176 21414.6
11.262 21847.2
11.335 21090.1
11.445 19575.96
11.543 17737.33
11.629 14600.85
11.69 11464.37
11.764 6813.73
11.825 4001.72
11.911 2163.09
12.021 648.927
12.376 0.0
 

wclaybaugh2

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Scott,

Roughly how much of the non propellant weight are the end fittings (forward closure / nozzle / nozzle retainers, the things that you would need to have if you changed from a metal tube to a composite tube. Reason im asking is, to help people analyze the potential mass fraction of a composite case vs a metal one. Assuming for a quick calculation that the weight of these items is the same (for an initial sizing) and the only weight change is the composite vs aluminum for the tube.

Mike K

Mike: I won't speak for Scott but I can observe that the aluminum tube in my 6" by 60" vehicle is about 13.4 lbsm. If I were to directly substitute carbon fiber (which is a bit less than half the density of aluminum) then a "black aluminum" tube of the same 0.125" wall would come in at a bit over 6.6 lbsm. This would result in the stage propellant fraction going from around 0.59 to about 0.64.

Simulation indicates that would change the ground launch performance from about 75 k feet to about 80K feet, payload included....

Bill
 

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Mike: I won't speak for Scott but I can observe that the aluminum tube in my 6" by 60" vehicle is about 13.4 lbsm. If I were to directly substitute carbon fiber (which is a bit less than half the density of aluminum) then a "black aluminum" tube of the same 0.125" wall would come in at a bit over 6.6 lbsm. This would result in the stage propellant fraction going from around 0.59 to about 0.64.

Simulation indicates that would change the ground launch performance from about 75 k feet to about 80K feet, payload included....

Bill

Because it is stronger the carbon fiber casing could also be thinner.

Bob Clark
 

kramer714

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Because it is stronger the carbon fiber casing could also be thinner.

Bob Clark
not stronger... higher specific strength (strength / weight)

Yield Strength 7075-T651 70 ksi 2.8 SG
Yield strength 301SS Hardened = 140ksi 8.0 SG
Yield strength 4140 hardened steel 135 ksi 7.9 SG

Ok for composite - T700 bidirectional cloth at 60%fv has a tensile ultimate Ftu1 of 115ksi and a Ftu2 of 85ksi (A Allowable RTD) at a 0/90 orientation, and a specific gravity of 1.57

Even a unidirectional composite at 60% fiber by volume gives a tensile ultimate of 270 ksi (RTD B ALLOWABLE)
Composite pressure vessels are typically THICKER wall than the metal ones they replace, plus will require some type of inner seal / liner for gas permeability. Uni properties are a great number if you are making tensile rods but doesn't just 'factor' into a composite multi axis pressure vessel with curved sides.

When you compare weight, you need to look at the entire 'ready to fly' system with end attachments, fin attachment, forward attachment, gas barrier / insulator etc. Many of these don't scale down well from a weight standpoint.

What is the design constraint? If it is overall thickness, a steel vessel without a gas barrier and welded on ends is the smallest and lowest cost. If the overall weight is the constraint, the composite MAY be lighter depending on the design and all the constraints.

Mike (Composites engineer with pressure vessel experience) K
 

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one more thing,

'it worked....' is great, but that isnt the same as 'it's certified... and tested... to a known standard... with a procedure.." Nar/ Tripoli say you cant use steel cases. plus they call out a factor of safety.

From NAR motor testing,

4.1.2 For metal-casing reloadable motors, the casing rupture pressure is at least twice the peak pressure
normally expected to be developed by the reloadable motor system to be tested in that casing.

So, what is the HIGHEST pressure of any reload, then test it to twice that number.

I was doing some 'un-official' consulting to a college team a bunch of years back, and talked about testing (this was for a metal lined filament would motor tube). I wasn't sure what to do when the 'composite tube person' on the team told me that they dont want to test the tube before loading it and flying it because 'what if it fails?', got worse when I asked what is the required proof pressure and operating pressure required for the tube.

So, are 'cots' motor tubes over weight, it depends on what the requirement is. To compare two designs are they the same.....
  • With the same margin - NAR requires burst >2x MAX operating pressure
  • With the same case operating pressure?
  • WIth the same fatigue limit stress, max principal strain, (going to use them more than once right?),
  • With manufacturing variations designed in (what is the thinnest wall thickness i can get that is still 'in tolerance' - did i test the thinnest or thickest case... remember on a .050 thick case +/-.005 inch is a high percentage!
  • With material allowables that are determined and predict the LOWEST possible value I may see with certified materials (Mil-Handbook 5 - Hdbk 17)
  • WIth normal defects from routine handling / manufacturing (is a scratch on the case ok? how deep?)
  • With the same mounting requirements - end skirt or attachment lugs...
  • with the same grain loading method? Rotomolded liner and cast in place, with an integral end dome is very different than a snap ring case with both ends open
  • with a design tolerant of routine handling (if it falls from the table to the floor is the case ruined?)
  • with a design that limit shrapnel distance in a failure (NAR calls this out!)
Are COTS motor cases the absolute lightest that they can be.. nope. But are they light weight (Certainly a reasonable weight) for the requirements on them, yup. .

Apples to apples and engineering not luck....

Mike K
 

OverTheTop

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Guidance. You say "oh we don't need guidance unless we are going to a specific orbit" and this is also wrong. Unless you are towing a massive amount of extra propellant/energy to make up for any imprecision, which is unlikely, your vehicle needs to follow a very precise course that spends its energy in precisely the amount and direction needed to reach a stable orbit. Failing at this means you crash before completing a single orbit. Active guidance is a must, especially on your upper stage, and you have by no means accomplished the hard part once you have a first stage booster.
No, not a must. Not wrong and not mainstream, but orbit without guidance is achievable. There is a Japanese launch vehicle that achieves orbit, without guidance, by setting a launch angle and using the gravity turn to determine the final orbit at burnout. I suspect the actual tolerances on the orbit would be a bit wide and probably need finessing with the payload stage if a more particularly accurate orbit was a requirement.

 

jsdemar

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No, not a must. Not wrong and not mainstream, but orbit without guidance is achievable. There is a Japanese launch vehicle that achieves orbit, without guidance, by setting a launch angle and using the gravity turn to determine the final orbit at burnout. I suspect the actual tolerances on the orbit would be a bit wide and probably need finessing with the payload stage if a more particularly accurate orbit was a requirement.

No, the final stage Japan Lambda-4S had an IMU, controller, and RCS.
https://open-aerospace.github.io/Lambda-4S/overview/gnc/
 
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