Reusable Rockets to Orbit

Rune · 2012-02-02 20:04:00

Well, we now know a nominal abort firing is about 5 seconds long. We also know, at last, that the abort G's will be close to solid rocket escape systems. About 6 G's, give or take. So the landing sequence is not going to last less than that, since it will be done at a fraction of the thrust. It could last up to 6 times longer, if the chutes stopped the capsule in mid-air and neglecting air resistance. So less than 30s of landing sequence, more than... say 10. A half-minute controlled descent would take away the landing ellipse? That's the question. Of course, that's assuming this new system has its own fuel provision, and all of it is used in both maneuvers (it would make sense to empty the tanks on an abort, that's for sure). They don't really have to keep the orbital propellant and the abort/landing propellant separated for any reason that I can think of, but that is not going to stop me form extrapolating performance as if they did.

Rune. Let's speculate away! It's free, and sometimes you get it right and feel good about yourself.

SpaceNut · 2012-02-02 20:43:13

While the experiment did not fully work it did take a big step towards proving that its possible...

Armadillo Launches a STIG-A Rocket; Captures Awesome Image of ‘Ballute’

GW Johnson · 2012-02-10 15:42:37

Ballutes are very interesting stabilizers and decelerators. For use at high-subsonic and low-supersonic speeds, I found round ribbon chutes superior in stability. But chutes cannot be used successfully during re-entry. There is the possibility that ballutes can.

If some sort of ablative coating could be added to the ballute, a gas-inflated conical ballute could be used as a decelerator and stabilizer drogue during re-entry. On the end of a long ablative-protected line, it could add deceleration capability over and above what the basic object's heat shield does. How much, I dunno. Up close, none.

If the ballute is really big, part of it will protrude outside the object's wake into freestream flow, even up close. Then it could really produce a lot of decelerator effect, at the cost of very high surface pressures on the freestream-exposed surfaces. It just needs a high-enough inflation pressure. Don't count on ram effect for that, you won't get it.

GW

Rune · 2012-02-11 07:08:44

Interesting piece on Popular Mechanics about SpaceX in general and the reusable Falcon 9 in particular: Click here for it.

As very interesting data in there, straight from Musk's mouth, I especially notice the staging of the first stage at lower speed (Mach 6 for the reusable booster, 10 for the expendable). Also, a 40% reduction in payload over the expendable version. Not to mention crazy talk in there about turnaround times that no one will believe until the company is actually doing it, like using a lower stage several times each day, and upper stage at least once a day, and that only because it hasn't got the crossrange to come back to the launch site sooner.

Rune. The second stage is starting to look more and more like a SSTO reusable spacecraft. Good luck with making it perform as it should.

Rxke · 2012-02-11 09:36:47

wowowowow....

Lofty goals. I like that!

He could even use a second stage + Dragon for suborbital tourists to make a quick buck and p1ss off VirginGalactic

Terraformer · 2012-02-11 10:08:33

Hmmm. This suggestion that they could get $1000/kg with reusability is very interesting, even if it means F9H can only launch 30 tonnes to orbit rather than 54. If they can drop crew costs on Dragon to maybe $5 million/person as well with reusability, it significantly changes the possible profitability of a Lunar base...

GW Johnson · 2012-02-11 10:21:33

A dark horse nobody is talking about is XCOR. Their Lynx should be flying for the first time by year's end. Suborbital flights, single stage, HTO/L on simple tricycle gear at very ordinary lightplane speeds, with a simple delta-wing rocket plane. Ground crew of 3 or 4 to refuel and fly again, multiple times a day. Stick and rudder flying, like a piper cub, just very fast straight up and back down. I've sat in their mockup. Even I could fly the thing.

Mark my words, it's people like that who will solve the practical space plane-to-orbit problem. Extremely tiny logistical tail, just like Spacex. Combined with true reusability, just like Spacex is trying to do. Very impressive. In some ways, they're ahead of Spacex, as regards designed-in reusability and long life airframes.

That kind of thinking did not (I repeat, did not) come out of a government lab. The real smarts is out there in the innovative companies. And it's not all science, either. I'm fond of saying "rocket science" ain't just science, it's about 40% science, 50% engineering art, and 10% blind dumb luck. Because it's true. The government labs have the science, but little or none of the art.

GW

RGClark · 2012-02-12 07:39:14

SpaceNut wrote:

While the experiment did not fully work it did take a big step towards proving that its possible...
http://www.universetoday.com/wp-content … allute.jpg
Armadillo Launches a STIG-A Rocket; Captures Awesome Image of ‘Ballute’

Thank for that. I didn't know they they had gotten that high at 82 km or had tested this recovery system, the 'ballute'.

Bob Clark

RGClark · 2012-07-16 15:31:19

GW Johnson wrote:

Shuttle leading edges went to 3000 F and required carbon-carbon composite precisely because the wing loading (vehicle weight divided by wing planform area) was up around 1 or 2 hundred pounds per square foot, just like a fighter jet, and all the space capsules. The white tiles on the sides and upper wing surfaces, and the black ones on the belly and lower wing surface, were low-density alumino-silicate, with a solid phase change that causes cracking at 2300 F. Those were restricted to peak 2000 F skin temperatures on the shuttle. Carbon-carbon is weak enough structurally, to be sure. Those tiles were far more fragile yet.
If the vehicle has a much larger aerosurface for its weight (low wing loading, say 10-20 pounds per square foot), peak skin temperatures reduce to under 2000 F, although total heat to be absorbed and disposed of actually increases. Skin temperature drives the material selection problem, the other is handled fairly easily. You will be decelerating at higher gees to make this happen. Plus, it's all a transient.
Myself, I rather like the idea of enduring a bit rougher ride in order to make my re-entry vehicle out of simple aluminosilicates, perhaps even plain old fire curtain cloth on a steel tube frame. I think it is very funny that the better, less fragile reentry vehicle might actually be built similar to the venerable old Piper Cub of the 1930's.
And, low density ceramics should be fiber reinforced as ceramic-ceramic composites, not that fragile stuff on the shuttle. I have done this, a quarter century ago. They are still extremely low density, yet fairly tough structurally. Really tough compared to that fragile nonsense they flew on shuttle. I used mine as a ramjet liner, which survived hours of burn and hundreds of excursions into very violent rich-blowout instability. The only reason I quit then was the project was done. Could have gone on for many more hours.
As it turns out, the materials I used then are still available. I checked just the other day.
GW

Thanks for that info, GW. There might be rocket stages that have such low wing loadings. First, a question though. If the vehicle is cylindrical should you use half the cylindrical surface area, representing the windward side, to calculate the wing loading or use just the rectangular cross-sectional area?
I'll use the more optimistic half-cylinder surface area to give an example of a rocket stage with such a low wing loading. Take a look at this diagram of the proposed DIRECT teams HLV:

Jupiter-241 Heavy-Lunar Cargo Launch Vehicle Configuration.
http://directlauncher.org/documents/Bas … 090608.jpg

The DIRECT team proposed using both common bulkhead design and aluminum-lithium alloy on the upper stage to get a highly weight optimized stage. The upper stage has the diameter of the shuttle ET at 8.4 m, with a propellant load of 199 mT and a dry mass of 12.8 mT using a J-2X engine. The J-2X engine weighs about 2,470 kg.
Suppose we swap out this engine for 2 SSME's at about 3,200 kg. Then our dry weight will be about 16,000 kg, 35,000 lbs. I'll show in a following post this will be a SSTO. But first about the wing loading. The density of hydrolox is 360 kg/m^3. Then for a propellant load of 199,000 kg, this corresponds to a volume of 199,000/360 = 550 m^3. Using the formula for the volume of a cylinder V = Pi*r^2*L with the radius at 4.2 m, for a diameter of 8.4 m we can calculate the length to be about 10 m. Then the half-cylinder surface area will be Pi*r*L = 130 m^2 = 1,400 sq.ft. Then the wing loading will be 35,000 lb/1,400sq.ft = 25 pounds per square foot. Fairly good but we can do better than that.
It turns out that for a given volume the cylinder side surface area gets larger as you make the radius smaller, which is what we want to get better wing loading. So let's take the diameter to be that of a Delta-IV at 5 m, but assume, unlike the Delta-IV, that we use common bulkhead design and aluminum-lithium alloy so that we keep our dry mass at the same low value of 16,000 kg, 35,000 lbs.
Note that the Delta-IV has about the same propellant load at 200 mT, but is not well weight optimized. But since the diameter of our new version is the same at 5 m, we could use the same tooling as used on the Delta-IV to produce this new weight-optimized, 2 SSME-powered version.
Again the volume is 550 m^3. We calculate now though the length for a 2.5 m radius to be 28 m. Then the half-cylinder surface area is 220 m^2 = 2,370 sq.ft. So the wing loading is 35,000 lb/2,370 = 15 pounds per square foot.

About the "rougher ride" at the lower wing loading, shouldn't the lighter weight at the same wing area allow you to get better lift resulting in lower gees?

Your post brought to mind some other possibilities that I'll discuss with you via email.

Bob Clark

GW Johnson · 2012-07-17 07:50:22

Most of the aerodynamics data for forces was correlated to flat planar areas. I'm guessing you are talking about dead-broadside return of these stages. If so, length x diameter is probably the reference area you want, and the broadside drag coefficients for circular cylinders apply. That stuff is a strong function of Mach between about 0.75M to about M3.

Nice to see such low numbers. That'll decelerate more, higher up in the thinner air, for sure. Ride might be rougher, just like a Piper Cub vs a jumbo jet. I'd worry about air pressure crush as it decelerates through about M1 about 20,000 feet. Structures that light are very flimsy. Air pressure crush after entry was over is what broke up both Skylab and shuttle Columbia's cabin section.

GW

RGClark · 2012-07-17 09:20:33

GW Johnson wrote:

Most of the aerodynamics data for forces was correlated to flat planar areas. I'm guessing you are talking about dead-broadside return of these stages. If so, length x diameter is probably the reference area you want, and the broadside drag coefficients for circular cylinders apply. That stuff is a strong function of Mach between about 0.75M to about M3.
Nice to see such low numbers. That'll decelerate more, higher up in the thinner air, for sure. Ride might be rougher, just like a Piper Cub vs a jumbo jet. I'd worry about air pressure crush as it decelerates through about M1 about 20,000 feet. Structures that light are very flimsy. Air pressure crush after entry was over is what broke up both Skylab and shuttle Columbia's cabin section.
GW

Yes, I was wondering about the air pressure effects too. But there are a couple of factors that made it worse for those two examples you mentioned, and also for the Falcon 9 first stage which also falls apart on reentry.
The main one is that these spacecraft or stages were tumbling, which puts severe stress on the structure. The second is that these structures were so heavy they came down so rapidly that they reached high density air while still traveling very fast. However, for a very light structure for the wing area, the lift should be enough to make the descent much shallower so it will have burned off much of the speed when it gets down to the higher density air.
We could perhaps also add some strengthening members to resist the bending loads that would not add too much to the weight made of carbon composites materials. I saw this new carbon composite isotruss bicycle frame when searching on lightweight composites:

That thing just looks like it would be light for the strength doesn't it?

I get now what you mean about a "rough ride" by your comparison of a Piper Cub to a jumbo jet. Perhaps the reaction control system could keep it stable considering its lightweight.

Bob Clark

Last edited by RGClark (2012-07-17 09:52:22)

GW Johnson · 2012-07-18 02:04:01

If you stabilize the object so it doesn't tumble, and build it stout enough so that it's tough enough to take the air loads, then your only real worry is the dings and dents from whacking the sea, even on a chute. Most of the time, shuttle SRB segments could be re-used, but not always. These were steel pressure vessels designed to operate routinely at about 900 psi chamber pressure. It was ocean impact that ruined the ones we couldn't re-use. Lightweight unpressurized liquid tankage gets smashed to rubble by ocean impact, even if the air loads don't crush it. That's why no Falcon stages have been re-used yet.

GW

RGClark · 2012-07-18 05:03:15

Hop wrote:

...
It seems to me one of the problems is achieving an FMR of 16:1 and having a spacecraft strong and temperature resistant enough to endure re-entry.
Watching Musk's Grasshopper video it looks like he hopes to use reaction mass to shed re-entry velocity in addition to aerobraking. If he hopes to achieve some re-entry delta V with propellant, this makes his FMR even more challenging. Get the FMR too high and you have a very tenuous, fragile vehicle even less able to endure re-entry. I'm not giving Musk's TSTO RLV even odds.
However, given propellant in orbit, I believe it's quite doable to decelerate the upper stage and land it intact on the launch pad. Thus I believe lunar supplied propellant depots would enable TSTO RLVs.

And not just TSTO's, and not just to orbit. But in fact SSTO's all the way to the Moon and back. It is a well known fact that if you have SSTO's, then with orbital refueling you can travel all the way to the Moon, land, lift off, and travel all the way back to Earth on that one single refueling. Another one of the many advantages of SSTO's. Note that this is not true for TSTO's whose second, orbital stage might get a delta v of, say, 6,000 m/s, insufficient for the round-trip to the Moon even with refueling in LEO.
I've been arguing that SSTO's are actually easy because how to achieve them is perfectly obvious: use the most weight optimized stages and most Isp efficient engines at the same time, i.e., optimize both components of the rocket equation. But I've recently found it's even easier than that! It turns out you don't even need the engines to be of particularly high efficiency.
SpaceX is moving rapidly towards testing its Grasshopper scaled-down version of a reusable Falcon 9 first stage:

Reusable rocket prototype almost ready for first liftoff.
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: July 9, 2012
http://www.spaceflightnow.com/news/n1207/10grasshopper/

SpaceX deserves kudos for achieving a highly weight optimized Falcon 9 first stage at a 20 to 1 mass ratio. However, the Merlin 1C engine has an Isp no better than the engines we had in the early sixties at 304 s, and the Merlin 1D is only slightly better on the Isp scale at 310 s. This is well below the highest efficiency kerosene engines (Russian) we have now whose Isp's are in the 330's. So I thought that closed the door on the Falcon 9 first stage being SSTO.
However, I was surprised when I did the calculation that because of the Merlin 1D's lower weight the Falcon 9 first stage could indeed be SSTO. I'll use GW Johnson's estimates for the Falcon 9 specs here:

WEDNESDAY, DECEMBER 14, 2011
Reusability in Launch Rockets.
http://exrocketman.blogspot.com/2011/12 … ckets.html

The first stage propellant load is given as 553,000 lbs, 250,000 kg, and the dry weight as 30,000 lbs, 13,600 kg. The Merlin 1C mass hasn't been released, but I'll estimate it as 650 kg, from its reported thrust and thrust/weight ratio. The Merlin 1D mass has been estimated to be 450 kg. Then on replacing the 1C with the 1D we save 9*200 = 1,800kg from the dry weight to bring it to 11,800 kg.
The required delta v to orbit is frequently estimated as 30,000 feet per second for kerosene-fueled vehicles, 9,144 m/s. When calculating the delta v your rocket can achieve, you can just use your engines vacuum Isp since the loss of Isp at sea level is taken into account in the 30,000 fps number. Then this version of the Falcon 9 first stage could lift 1,200 kg to orbit:

310*9.81ln(1 + 250/(11.8 + 1.2)) = 9,145 m/s.

Then the Falcon 9 first stage could serve as a proof of principle SSTO on the switch to the Merlin 1D.

Bob Clark

GW Johnson · 2012-07-21 13:26:05

Hi Bob:

Chemical SSTO does indeed seem possible once one can achieve stage inert mass fractions in the 5% range. I do have some very serious doubts about such stages ever being reusable in any way. That's the sort of inert weight of something like the Gossamer Condor, fragile beyond belief. No wood and fabric, metal and fabric, or all-metal airplane has ever been strong enough to fly routinely with a good safety record below an inert fraction of about 40%. The ones with the best survival-of-battle-damage records are closer to 50% inerts.

I know space launch vehicles are not airplanes, but you see my point. There is some sort of a lower limit for inert weight fractions that really are going to be reusable. I think they're way above 10% (shuttle SRB's were near that 10% figure, and were not actually fully reusable - there was a significant loss rate of motor case segments).

Last time I thought about space-faring airframes, I guessed about half and half composites and metals, and pivoted off that 40% all-metal aircraft figure with composites twice the strength-to-weight. You cannot use all-composite, they have a very low tolerance for heating.

I was using about 27% inert for an airframe that might conceivably fly into space reusably thousands of times (or more). At something in the neighborhood of 27%, no chemical rocket is capable of SSTO, but some nuclear rockets might be. The only other option was airbreathing propulsion, but I found scramjet a dead end, even assuming it could be made technologically ready (they've been trying since about 1960).

That was a non-intuitive dead end, too. Few people even today know it's a dead end. To get the thrust out of the engine to climb, you have to stay deeper in the atmosphere as you accelerate. But that's where drag is higher. Beyond about Mach 10 or 12 or so, you incur drag losses faster than you can maintain climb rate, and those drag losses eat up all your gains from the higher-than-rocket Isp. Dead end. I found that answer for almost nothing (pencil, paper, some hours of my time). The X-30 required billions of dollars to find the same answer (but they never actually publicly admitted what they found).

So, I think there's vertical launch rockets and there's TSTO spaceplanes (unless you allow nuke rocket spaceplanes). Any airbreathing assist occurs during the lower altitude phases early in the trajectory for both.

On a vertical launch rocket, you've pretty well left the usable atmosphere around 70,000 feet, at speeds around M 2 to at most 3. I could put ramjet strap-ons on the first stage, but they'd be simple pitot-inlet designs, running from about M0.5-ish to at most about M2.5-ish. That's 1945 technology, except that I would use a 1970-vintage dump combustor instead of a baffle stabilizer. Much better suited to variable-speed operation.

For the TSTO spaceplane, I would put both rockets and a high-speed ramjet engine in the first stage. Rocket off the runway to about M1.5-M1.6 takeover with an external-compression inlet. Ramjet climb at low Mach (1.6 to 2-ish), pullover and accelerate to about M5-6-ish at 60,000 feet, light the rockets, too, and pull up at 45 degrees to release the second stage. Cut rockets and chop ramjet throttle to decelerate in a 180 degree turn, and cruise back to base at M1.5-ish for a glide landing. Rocket power for emergency go-around or divert. The second stage is just a plain old rocket plane.

The ramjet would have the supersonic inlet, a C-D nozzle of throat area about 65% of engine flow area, and a dump combustor design. All fixed geometry, except possibly a translating inlet spike to maintain shock-on-lip operation. It likely would completely fill (be) the fuselage. The wings and fillets would contain the rockets and all the propellants.

GW

RGClark · 2012-07-26 14:40:59

RGClark wrote:

...
The first stage propellant load is given as 553,000 lbs, 250,000 kg, and the dry weight as 30,000 lbs, 13,600 kg. The Merlin 1C mass hasn't been released, but I'll estimate it as 650 kg, from its reported thrust and thrust/weight ratio. The Merlin 1D mass has been estimated to be 450 kg. Then on replacing the 1C with the 1D we save 9*200 = 1,800kg from the dry weight to bring it to 11,800 kg.
The required delta v to orbit is frequently estimated as 30,000 feet per second for kerosene-fueled vehicles, 9,144 m/s. When calculating the delta v your rocket can achieve, you can just use your engines vacuum Isp since the loss of Isp at sea level is taken into account in the 30,000 fps number. Then this version of the Falcon 9 first stage could lift 1,200 kg to orbit:
310*9.81ln(1 + 250/(11.8 + 1.2)) = 9,145 m/s.
Then the Falcon 9 first stage could serve as a proof of principle SSTO on the switch to the Merlin 1D.

I think we can probably do better than that first estimate of the Falcon 9 first stage with Merlin 1D's as a SSTO. The Merlin 1D has a 147,000 lb sea level thrust:

Modified Merlin engine completes full duration firing.
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: June 25, 2012
http://www.spaceflightnow.com/news/n1206/25merlin1d/

The gross mass of the Falcon 9 first stage with the Merlin 1D's and a 1.2 mT payload would be 250 + 11.8 + 1.2 = 263 mT, 578,600 lbs. This could be lofted by just 4 of the Merlin 1D's. But the thrust would be just a little over the gross mass resulting in high gravity loss. So let's use 5 Merlins. Subtracting off 4 Merlins makes the dry mass 11,800 - 4*450 = 10,000 kg.
The number of 30,000 fps delta-v for LEO is assuming a T/W ratio common for liquid fueled rockets, in the range 1.1 to 1.2. With all 9 Merlins the T/W ratio would above 2.2. This would result in a much reduced gravity loss. So the required delta-v would be less than the 30,000 fps number, and so actually higher than 1.2 mT could be sent to LEO even in that case.
But let's look at the case of using 5 Merlins. SpaceX has given a vacuum Isp of the Merlin 1D as actually 311 s. Then we could send 3.1 mT to LEO:

311*9.81ln(1 + 250/(10 + 3.1)) = 9,152 m/s.

SpaceX has said though they want to move to a larger version of the Falcon 9 called the Falcon 9 v1.1, in accordance with the Merlin 1D's larger thrust. The Falcon Heavy will use this version's first stage for its core stage and side boosters. SpaceX expects the Falcon 9 v1.1 to be ready by the end of the year.
Elon Musk has said the version 1.1 will be about 50% longer:

Q&A with SpaceX founder and chief designer Elon Musk.
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: May 18, 2012
http://www.spaceflightnow.com/falcon9/003/120518musk/

I'll assume this is coming from 50% larger tanks. This puts the propellant load now at 375,000 kg. Interestingly SpaceX says the side boosters on the Falcon Heavy will have a 30 to 1 mass ratio. This improvement is probably coming from the fact it is using the lighter Merlin 1D engines, and because scaling up a rocket actually improves your mass ratio, and also not having to support the weight of an upper stage and heavy payload means it can be made lighter.
So I'll assume for this SSTO version of the Falcon 9 v1.1 the mass ratio is 30 to 1, which makes the dry mass 13 mT. Then this version can lift 6.7 mT to LEO:

311*9.81ln(1 + 375/(13 + 6.7))

Bob Clark

RGClark · 2012-07-31 17:52:33

RGClark wrote:

...
So I'll assume for this SSTO version of the Falcon 9 v1.1 the mass ratio is 30 to 1, which makes the dry mass 13 mT. Then this version can lift 6.7 mT to LEO:
311*9.81ln(1 + 375/(13 + 6.7))
Bob Clark

I didn't include the answer to that last calculation:

311*9.81ln(1 + 375/(13 + 6.7)) = 9,145 m/s.

Dr. John Schilling has produced a payload estimation program:

Launch Vehicle Performance Calculator.
http://www.silverbirdastronautics.com/LVperform.html

It gives a range of likely values of the payload. I've found the midpoint of the range it specifies is a reasonably accurate estimate to the actual payload for known rockets.
Input the vacuum values for the thrust in kilonewtons and Isp in seconds. The program takes into account the sea level loss. SpaceX gives the Merlin 1D vacuum thrust as 161,000 lbs and vacuum Isp as 311 s:

FALCON 9 OVERVIEW.
http://www.spacex.com/falcon9.php

For the 9 Merlins this is a thrust of 9*161,000*4.46 = 6,460 kN. Use the default altitude of 185 km and the Cape Canaveral launch site, and a 28.5 degree orbital inclination, to match the Cape's latitude.
Input the dry mass of 13,000 kg and propellant mass of 375,000 kg. Then it gives an estimated 7,564 kg payload mass:

Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit:    185 x 185 km, 28 deg
Estimated Payload:    7564 kg
95% Confidence Interval:    3766 - 12191 kg

This may be enough to launch the Dragon capsule, depending on the mass of the Launch Abort System(LAS).

Bob Clark

Last edited by RGClark (2012-07-31 23:34:13)

RGClark · 2012-08-04 11:37:34

RGClark wrote:

...
Dr. John Schilling has produced a payload estimation program:
Launch Vehicle Performance Calculator.
http://www.silverbirdastronautics.com/LVperform.html
It gives a range of likely values of the payload. I've found the midpoint of the range it specifies is a reasonably accurate estimate to the actual payload for known rockets.
Input the vacuum values for the thrust in kilonewtons and Isp in seconds. The program takes into account the sea level loss. SpaceX gives the Merlin 1D vacuum thrust as 161,000 lbs and vacuum Isp as 311 s:
FALCON 9 OVERVIEW.
http://www.spacex.com/falcon9.php
For the 9 Merlins this is a thrust of 9*161,000*4.46 = 6,460 kN. Use the default altitude of 185 km and the Cape Canaveral launch site, and a 28.5 degree orbital inclination, to match the Cape's latitude.
Input the dry mass of 13,000 kg and propellant mass of 375,000 kg. Then it gives an estimated 7,564 kg payload mass:
Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit:    185 x 185 km, 28 deg
Estimated Payload:    7564 kg
95% Confidence Interval:    3766 - 12191 kg

This may be enough to launch the Dragon capsule, depending on the mass of the Launch Abort System(LAS).

Bob Clark

According to this report from 2010, ESA was considering plans to use the Orion on the Ariane 5 to get a European manned spaceflight capability:

French govt study backs Orion Ariane 5 launch.
By Rob Coppinger
on January 8, 2010 4:45 PM
http://www.flightglobal.com/blogs/hyper … rench.html

This would cost several billion dollars to man-rate the Ariane 5. I have to believe the solid rocket boosters, which can not be shut down when started, play a significant role in that high cost.
However, the ESA has now given up on an indigenous manned spaceflight capability because of the estimated billion dollar cost to man-rate the full Ariane 5 system:

WSJ: Europe Ends Independent Pursuit of Manned Space Travel.
"LE BOURGET, France—Europe appears to have abandoned all hope of
independently pursuing human space exploration, even as the region's
politicians and aerospace industry leaders complain about shrinking
U.S. commitment to various space ventures.
"After years of sitting on the fence regarding a separate, pan-
European manned space program, comments by senior government and
industry officials at the Paris Air Show here underscore that budget
pressures and other shifting priorities have effectively killed that
longtime dream."
http://www.orbiter-forum.com/showthread.php?t=23006

In contrast, the Ariane 5 core stage with an added, second Vulcain engine could serve as a SSTO to carry a manned capsule to orbit. JAXA was able to add a second cryogenic engine to their H-IIa rockets first stage, which is about the same size as that of the Ariane 5, for less than a $250 million dollar development cost:

Rocketing to the future.
http://www.gov-online.go.jp/pdf/hlj_ar/ … /05-07.pdf

Mitsubishi Heavy To Invest In Next-Generation Rocket.
by Staff Writers
Tokyo, Japan (AFX) Jun 14, 2006
http://www.spacedaily.com/reports/Mitsu … ocket.html

The $250 million cost number I'm getting from the exchange rate from yen to dollars used in the second article. It is important to note about $50 million of this was used to widen the tanks, which wouldn't be needed in the Ariane 5 case. So we can estimate the development cost without widened tanks as less than $200 million. This is about the amount of the subsidy that the ESA gives to ArianeSpace every year. But this would be for a 4 year development, judging by the JAXA case, so would only be $50 million a year.

In the calculations for this multi-Vulcain Ariane core stage, I used this page for the specifications on the Ariane:

Space Launch Report: Ariane 5 Data Sheet.
http://www.spacelaunchreport.com/ariane5.html#config

For the Vulcain 2 specifications, I've seen different numbers in different sources, though close to each other. I'll use this source:

Vulcain 2.
http://www.astronautix.com/engines/vulcain2.htm

I'll also use the earlier Ariane 5 "G" version that is lighter than the current "E" version to be lofted by two Vulcains without side boosters. According to the SpaceLaunchReport page it had a 170 mT gross mass for the core at a 158 mT propellant load, giving a 12 mT dry mass.
According to the Astronautix page, Vulcain 2 has a 434 s vacuum Isp and 1350 kN vacuum thrust. So two will have a 2700 kN vacuum thrust. The Vulcain's mass is listed as 1,800 kg. So adding another will bring the stage dry mass to 13,800 kg.
Now input this data into Schilling's calculator. Select again default residuals and select "No" for the "Restartable Upper Stage?" option. Select the Kourou launch site for this Ariane 5 core rocket. For the orbital inclination, I input 5.2 degrees. I gather Schilling uses this for Kourou's latitude since deviating from this decreases the payload. I chose also direct ascent for the trajectory.

Then the result I got was 7,456 kg(!) to orbit:

================================
Mission Performance:
Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Guiana Space Center (Kourou)
Destination Orbit: 185 x 185 km, 5 deg
Estimated Payload: 7456 kg
95% Confidence Interval: 4528 - 10898 kg
================================

Interestingly, the payload capability of the Falcon 9 v1.1 first stage and of this two-Vulcain Ariane 5 core stage would be about the same as SSTO's.

Bob Clark

Rune · 2012-08-04 12:49:07

...And I'm sorry to say, you committed similar conceptual errors in both. Like assuming doubling the thrust in the core and taking our the rigid heavy boosters the whole thing hangs from does nothing to the structural weight in this last Arianne case, or like working out numbers assuming a 1.2mT payload weight to come up with a 7.5mT payload in the Falcon example. At first sight your analysis are full of correct data, but you should take care how you add and subtract things, because making changes to a subsystem impacts the whole system.

Rune. Look twice, then give it to someone to double check. You will probably miss something anyhow, that's my philosophy.

Antius · 2012-08-17 06:56:24

I based one of my nuclear engineering MSc projects on an SSTO vehicle. The conclusion was that a H2/O2 propelled SSTO was unlikely to be economic.

The basic problem is that with hydrogen/oxygen propellant (Isp = 440s), some 90% of the vehicle weight needs to be fuel. This requires a lightweight structure, very high performance engines, lightweight heat shield, etc. When all is added up, the payload fraction ends up no greater than 1-2%. When one considers the engineering cost of such an optimised vehicle, between-flight refurbishment costs and site related launch costs, it is difficult to see how such a vehicle would achieve much cost saving over a conventional rocket. Like the shuttle, it could easily end up being more expensive.

Using some seriously unpleasant chemical fuels (H2/F2, or H2/F2/Li) improved its physical performance, but introduced other hazard related costs and pollution difficulties.

The best solution would appear to be to use a nuclear thermal rocket engine for ground launch. Payload fraction then increases to about 30% and all costs are spread over much larger orbital payloads. A sea launch avoids the need for a ground based launch site and cuts costs further. The problem of crash related accident risks is mitigated by using element fuels that drain fission products into a carbon bottle that can be removed after each use. The vehicle can then be used dozens of times without the need for a core change.

GW Johnson · 2012-08-20 10:30:30

I am unconvinced that the financial benefits of launch rocket reusability will ever be dramatically large. The real cost-reducer has been the transition to smaller logistical tails. But I could be wrong.

That being said, I don't see the desirability of going for reusability and SSTO in the same design, since both of these depend critically upon inert mass fractions, but with opposite signs on their trends. I rather suspect that practical SSTO designs will always be throwaway items, driven there by 4-5% inert fractions to get 2-3% payload fractions. KeroLox, LH2-LOX, no difference. It's chemical, all you can get is 310-470 sec Isp.

A reusable launch rocket will be 2 or maybe even 3 stages, so that inert fractions in the 10-20% range can be tolerated and still get 2-3% payload fractions. I'm not at all sure that the second and thirds stages should be reusable, but I think a reusable first stage is feasible, if it's not too large.

We have enough trouble landing million pound airplanes without driving the landing gears up through the airframe. You can't whack the air broadside hypersonically, or the sea at around 50 mph, without damage to lightweight structures. So they cannot be lightweight. Metals and composites are only so strong.

Mass scales as dimension cubed, while part cross-section area scales as dimension squared. Better to stay smaller with reusable items.

I agree with Antius that a nuclear thermal rocket provides the Isp for a practical surface-launch SSTO, probably as a winged spaceplane of medium size. I'm not sure the public could be sold on reactor safety in the event of an abort or crash, not with solid core. A gas core lightbulb engine is a better sell for that application. The safest version for crash or abort is open-cycle gas core, but the somewhat radioactive plume will be a hard sell. Those things being true, I'm not so very sure the best application for nukes is surface launch. It would be a lot easier to sell them as orbit-to-orbit propulsion to and from Earth. Not trivial, but an easier sell.

GW

Last edited by GW Johnson (2012-08-20 10:35:49)

RGClark · 2012-08-20 19:06:06

Rune wrote:

...And I'm sorry to say, you committed similar conceptual errors in both. Like assuming doubling the thrust in the core and taking our the rigid heavy boosters the whole thing hangs from does nothing to the structural weight in this last Arianne case, or like working out numbers assuming a 1.2mT payload weight to come up with a 7.5mT payload in the Falcon example. At first sight your analysis are full of correct data, but you should take care how you add and subtract things, because making changes to a subsystem impacts the whole system.

It's all theoretical until we know if the Falcon Heavy side boosters really wind up having a 30 to 1 mass ratio. The new version of the Falcon 9, version 1.1, is due to be introduced by the end of the year. We'll have a better idea then.

Bob Clark

Antius · 2012-08-25 04:02:51

One compromise solution is to combine the benefits of the BDB with the SSTO in a way that reduces the limitations of both. Truax considered something similar in the early 1960s, if memory serves. Produce a two-stage to orbit, sea-launch vehicle with a take-off mass perhaps 10 times that of a Saturn V.

The lower stage would be constructed from alloy steels similar to those used for submarine pressure hulls and would have a fuelled mass of perhaps 20,000 tonnes. This would make it suitable for submarine-yard engineering techniques and construction cycles would fit in between sub-build programmes.

The fuel would be LoX/Heavy Oil with tanks pressurised with nitrogen and helium. The single pressure fed engine would include a fibreglass ablative lining that would be replaced after each use. The lower stage would achieve a delta-V of about 3km/s and would be used in pure booster mode - i.e. high acceleration, intermediate ISP, no transverse velocity increment. Following use, it would seperate and fall into the ocean, where it would be collected for reuse. The advantage of using steel tanks and engine bells is a good repeatable stress-cycle, so minor refurbishment would be needed between launches, i.e. engine liner replacement.

The upper stage would be constructed primarily from fibreglass composites and fuelled with LoX & liquified natural gas. Its delta-V would be 6km/s and its engine bell optimised for building up orbital velcocity in vacuum. This would be expendable in the traditional sense, but will be engineered to allow canibalisation upon reaching orbit. Its material construction would principally be polymers, which contain chemically valuable carbon and hydrogen for high-orbital manufacturing industries. Electrical components will be designed as modular units that can be slotted out and reused for other applications. Other components such as s-glass fibres, would be ground and used as mass-driver propellants.

In orbit, the the upper stage would be intercepted by a mass driver vehicle, which would carry it to the L5 manufacturing centres where it would be deconstructed. The upper stage will therefore be 'sold' upon reaching orbit. So in fact we have a completely reusable vehicle.

RGClark · 2012-08-27 22:24:14

GW Johnson wrote:

Most of the aerodynamics data for forces was correlated to flat planar areas. I'm guessing you are talking about dead-broadside return of these stages. If so, length x diameter is probably the reference area you want, and the broadside drag coefficients for circular cylinders apply. That stuff is a strong function of Mach between about 0.75M to about M3.
Nice to see such low numbers. That'll decelerate more, higher up in the thinner air, for sure. Ride might be rougher, just like a Piper Cub vs a jumbo jet. I'd worry about air pressure crush as it decelerates through about M1 about 20,000 feet. Structures that light are very flimsy. Air pressure crush after entry was over is what broke up both Skylab and shuttle Columbia's cabin section.
GW

I looked up the dynamic pressure of the space shuttle during reentry and found it was in the range of 300 psf (pounds per square foot). This is only 300/144 = 2.1 psi, 1/7th bar. This would not seem to be too difficult to deal with especially for a pressurized structure flying in a streamlined fashion.

Bob Clark

RGClark · 2012-08-27 23:40:45

GW Johnson wrote:

I am unconvinced that the financial benefits of launch rocket reusability will ever be dramatically large. The real cost-reducer has been the transition to smaller logistical tails. But I could be wrong.
GW

Watching the retrospectives on Neil Armstrong on NASA TV, I saw he was a X-15 pilot. This reminded me we actually had reusable rocket engines from the very earliest days of manned rocket-powered flight. The XLR-99 engine used on the X-15 was reusable for 20 to 40 times before overhaul, after which it could be reused again:

XLR-99.
http://www.astronautix.com/engines/xlr99.htm

The 3 copies of the X-15 aircraft flew for a total of 199 flights. Can you imagine how expensive that program would have been if an entire new X-15 aircraft had to be used for each flight?

Bob Clark

GW Johnson · 2012-08-28 19:06:35

Excluding the B-52 it dropped from, the inert mass fraction of the basic X-15 was right at 40%, somewhat like many other very fast aircraft. Even in the X-15-A-2 configuration with the big external tanks, it was still near this inert figure.

Kinda takes structural weight to confer the strength to be tough as an old boot at Mach 6, don't it?

Supersonic USAF bombers were about 48-53% inert, and supersonic USN bombers like the A-5 were about 58% inert. Even the subsonic carrier-based A-7 was 50% inert. 17 US tons dry, with 17 more US tons of fuel and ordnance. Tough old birds.

GW

Last edited by GW Johnson (2012-08-28 19:07:57)

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