LV Design to launch full-size DreamChaser on Strato-Launch

[Todd Martin]

It would be helpful for SNC & Stratolaunch if the Stratolaunch system was powerful enough to launch a full scale Dreamchaser.  A full size DC would be cheaper to complete development than working on a scale-down.  A full size DC could compete on commercial crew services to ISS and compete on ISS cargo.  A Stratolaunch system which could loft a full size DC would also be more capable in competing on commercial satellite business.

I've done some work on rocket design to see if this is feasible.  Below, I lay out the launch vehicle (LV) which I believe could do the job.

GLOW (Gross Liftoff Weight):  227,273 kg (500,000 lbs).
 
1st Stage: 
  3 J2X engines with nozzles optimized for ignition at 30,000 feet.  ISP:  419 seconds, total thrust 3683KN
  Dry Mass 22,900kg
  Prop mass 169883kg
2nd Stage:  Delta IV 2nd stage with 4 meter fairing
  2 RL10B engines:  ISP:  470 seconds, total thrust 220KN
  Dry Mass 2780kg
  Prop mass 20410kg
Dreamchaser (no fairing):  Mass = 11,300kg

Length of Launch Vehicle:  100 feet (1st stage) + 40 feet (2nd stage) + 30 feet (DC) = 170 feet.

Stratolaunch aircraft provides 250 meters/sec initial velocity.  In addition, the LV is launched at a 15 degree angle of attack instead of vertically, which provides another 600 meters/sec advantage.  Finally, StratoLaunch allows for launch closer to the equator (by 1500 km).

The question is whether Stratolaunch can launch a full-size DC with the appropriate LV mounted to its belly.  If so, what would be the best design to accomplish this requirement.  I've made my best shot at it, I invite others to review & comment.  Thanks!

[dror]

Is it even possible to carry and launch a cryogenic vehicle from this platform?

[Todd Martin]

Is it even possible to carry and launch a cryogenic vehicle from this platform?

That's a good question.  From what I understand, there are 2 issues which must be considered.  First, are the engines able to be lit while in flight?  How much boil-off of the cryogenic fuel will take place before LV ignition?

For the LV proposal, I selected the J2-X engine.  The J2-X can be lit while in flight, aka "Air-start"; it also has a higher ISP than the RS-68 engine or RS-68A engine as used on the Delta IV.  The SSME would have an even higher IPS, but the SSME is not an "Air-start" engine.

As for Boil-off, the rate depends on the insulation, tank diameter, and time.  It is an issue which would need careful analysis.  I can think of a few ways to mitigate boil-off, but at this time I am really interested in knowing whether the concept is feasible.

[Rocket Science]

Is it even possible to carry and launch a cryogenic vehicle from this platform?

That's a good question.  From what I understand, there are 2 issues which must be considered.  First, are the engines able to be lit while in flight?  How much boil-off of the cryogenic fuel will take place before LV ignition?

For the LV proposal, I selected the J2-X engine.  The J2-X can be lit while in flight, aka "Air-start"; it also has a higher ISP than the RS-68 engine or RS-68A engine as used on the Delta IV.  The SSME would have an even higher IPS, but the SSME is not an "Air-start" engine.

As for Boil-off, the rate depends on the insulation, tank diameter, and time.  It is an issue which would need careful analysis.  I can think of a few ways to mitigate boil-off, but at this time I am really interested in knowing whether the concept is feasible.

Those J2-X engines are pretty expensive to throw away, it would make more sense if this were a flyback stage...

[jongoff]

As for Boil-off, the rate depends on the insulation, tank diameter, and time.  It is an issue which would need careful analysis.  I can think of a few ways to mitigate boil-off, but at this time I am really interested in knowing whether the concept is feasible.

It's probably possible to make this work from a boiloff standpoint using something like what the Quest Thermal guys are developing (http://www.questthermal.com/products/launch-vehicle-mli). Apologies to those who've already seen me comment on this before. But basically they have a way of making a lightweight vacuum-jacketed MLI insulation solution that can work on the outside of launch vehicles. Cuts the heat leak way down, and could theoretically be combined with a vapor cooled shield of some sort (Quest did some experiments in this direction previously) to take what boiloff you do get, and use it to intercept heat leaking into the tanks. If pure insulation isn't enough you could have either a small topping tank that you disconnect shortly before launch (it can be insulated a ton better), or subcool one or both propellants. Lots of ways to do this, but the first key is good insulation.

I think a LOX/LH2 air-launch is doable, though I'm not endorsing (or dissing) this specific approach.

~Jon

[jongoff]

Is it even possible to carry and launch a cryogenic vehicle from this platform?

That's a good question.  From what I understand, there are 2 issues which must be considered.  First, are the engines able to be lit while in flight?  How much boil-off of the cryogenic fuel will take place before LV ignition?

For the LV proposal, I selected the J2-X engine.  The J2-X can be lit while in flight, aka "Air-start"; it also has a higher ISP than the RS-68 engine or RS-68A engine as used on the Delta IV.  The SSME would have an even higher IPS, but the SSME is not an "Air-start" engine.

As for Boil-off, the rate depends on the insulation, tank diameter, and time.  It is an issue which would need careful analysis.  I can think of a few ways to mitigate boil-off, but at this time I am really interested in knowing whether the concept is feasible.

Those J2-X engines are pretty expensive to throw away, it would make more sense if this were a flyback stage...

With air-launch, it's a lot easier to get booster RTLS without having to pay a big penalty: http://selenianboondocks.com/2008/09/orbital-access-methodologies-part-vi-air-launched-glideforward-tsto/

You just have your drop point be up-range of your landing point, so your booster glides forward to somewhere near the airstrip you took off from (after factoring in any retro-burns for deceleration, etc). As for how to do the landing, you could have some dedicated landing engines, and land either vertically or horizontally. Or you could use wings, or helicopter blades, or possibly even mid-air recovery. Lots of options, though my personal faves in this case would be horizontal rocket powered landing, or helicopter landing.

How heavy would the first stage be dry?

~Jon

[Todd Martin]

Is it even possible to carry and launch a cryogenic vehicle from this platform?

Those J2-X engines are pretty expensive to throw away, it would make more sense if this were a flyback stage...

How heavy would the first stage be dry?

~Jon

Dry mass of the 1st stage would be 22,900 kg (50,380 lbs) with a 5 meter diameter tank.

[Comga]

Those J2-X engines are pretty expensive to throw away, it would make more sense if this were a flyback stage...

With air-launch, it's a lot easier to get booster RTLS without having to pay a big penalty: http://selenianboondocks.com/2008/09/orbital-access-methodologies-part-vi-air-launched-glideforward-tsto/

You just have your drop point be up-range of your landing point, so your booster glides forward to somewhere near the airstrip you took off from (after factoring in any retro-burns for deceleration, etc). As for how to do the landing, you could have some dedicated landing engines, and land either vertically or horizontally. Or you could use wings, or helicopter blades, or possibly even mid-air recovery. Lots of options, though my personal faves in this case would be horizontal rocket powered landing, or helicopter landing.
~Jon

This isn't a really serious disucssion, but for it, Jon has a good point, which we knew he would bring up again.  It answers the question "What Would Elon Do?" in that it puts affordability over performance.
Rather than fly to the equator to gain some percentage in payload to orbit (Think Sea Launch sailing two giant ships to the equator) the launch is optimized for economy, in this case recovery of expensive hardware.

Then you have to work on the cost of two RL-10s.  If we are discussing "pie in the sky" you might want to hypothesize another hydrolox engine, like from Blue Origin or Xcor.

[jongoff]

Dry mass of the 1st stage would be 22,900 kg (50,380 lbs) with a 5 meter diameter tank.

Yeah, I saw that after I posted the comment. 50klb is fairly heavy. You'd probably be talking about 2-4x RL-10 class engines for landing.

~Jon

[jongoff]

Those J2-X engines are pretty expensive to throw away, it would make more sense if this were a flyback stage...

With air-launch, it's a lot easier to get booster RTLS without having to pay a big penalty: http://selenianboondocks.com/2008/09/orbital-access-methodologies-part-vi-air-launched-glideforward-tsto/

You just have your drop point be up-range of your landing point, so your booster glides forward to somewhere near the airstrip you took off from (after factoring in any retro-burns for deceleration, etc). As for how to do the landing, you could have some dedicated landing engines, and land either vertically or horizontally. Or you could use wings, or helicopter blades, or possibly even mid-air recovery. Lots of options, though my personal faves in this case would be horizontal rocket powered landing, or helicopter landing.
~Jon

This isn't a really serious disucssion, but for it, Jon has a good point, which we knew he would bring up again.

Thanks, I think. :-)

Re: the 2x RL-10 engines, those aren't a showstopper if you're not recovering them, but going either with XCOR or Blue Origin engines or upper stage recovery/reuse (working on it!) would definitely help from an economics standpoint.

But the underlying concept even without reuse wouldn't be a non-starter, though I don't know if it'd be any cheaper than say an Atlas V. Especially once you factor in the development costs.

~Jon

[Todd Martin]

An RL10-B has a thrust of 110KN in vacuum, so you would need at least 3 to provide 1G of thrust to 22,900kg of dry mass + some prop at the ground.

Alternatively, a J2X engine can be throttled down to 82%.  If a center placed J2X was given a ground level type nozzle, it could do a vertical landing burn via "hover slam" at around 4 G's.

According to this link, http://depletedcranium.com/we-need-a-new-cheaper-upper-stage-engine/ the price of an RL10 is $38 Million while a J2-X would be $25 Million each.  So, my configuration for the LV of 3 J2X & 2 RL10's has a total engine cost of $151 million which is perhaps why a hydrolox concept is not being pursued for Stratolaunch.  The Atlas V / CST-100 combo will be using 2 RL10's also, so it isn't completely out of the realm of possibility. 

It seems to me that if commercial price pressures were to be brought to bear on hydrolox engines, these prices could be substantially reduced and become competitive with kerolox engines.  If Stratolaunch went this route, such savings could be realized.

 

 

[Patchouli]

An RL10-B has a thrust of 110KN in vacuum, so you would need at least 3 to provide 1G of thrust to 22,900kg of dry mass + some prop at the ground.

Alternatively, a J2X engine can be throttled down to 82%.  If a center placed J2X was given a ground level type nozzle, it could do a vertical landing burn via "hover slam" at around 4 G's.

According to this link, http://depletedcranium.com/we-need-a-new-cheaper-upper-stage-engine/ the price of an RL10 is $38 Million while a J2-X would be $25 Million each.  So, my configuration for the LV of 3 J2X & 2 RL10's has a total engine cost of $151 million which is perhaps why a hydrolox concept is not being pursued for Stratolaunch.  The Atlas V / CST-100 combo will be using 2 RL10's also, so it isn't completely out of the realm of possibility. 

It seems to me that if commercial price pressures were to be brought to bear on hydrolox engines, these prices could be substantially reduced and become competitive with kerolox engines.  If Stratolaunch went this route, such savings could be realized.

 

 

Replace the three RL-10s with a single BE-3 the ISP is lower but it's thrust is much higher so you only need one and it's a lot cheaper.
Though on the cost of the J-2X 25 million was with a low production for Ares would have needed.
If it was used in a rocket that flew more often it might be cheaper.

As for return  the stage already has wings and is launched horizontally  here glide back sort of landing and using landing gear may be more idea then transitioning the stage to vertical and landing it.

I wonder could the hybrids in DC be replaced with two Chase-10 or Xcor methane engines allowing it to effectively be it's own upper stage?

[ChrisWilson68]

Replace the three RL-10s with a single BE-3 the ISP is lower but it's thrust is much higher so you only need one and it's a lot cheaper.
Though on the cost of the J-2X 25 million was with a low production for Ares would have needed.
If it was used in a rocket that flew more often it might be cheaper.
I wonder could the hybrids in DC be replaced with two Chase-10 or Xcor methane engines allowing it to effectively be it's own upper stage?

Then you'd have to take your upper stage tanking through re-entry and landing.  There's no room in the current DC design for that.

[Robotbeat]

For air-launch, hydrogen is pretty dang optimal. The Isp advantage makes a big difference since you're limited strongly by lift-off-mass (or, more pedantically, slung mass), more so than a ground launch where you can compensate with more thrust or a couple more solids without much trouble.

[Oli]

ISP:  419 seconds

Shouldn't that be more?

Is it even possible to carry and launch a cryogenic vehicle from this platform?

That's a good question.  From what I understand, there are 2 issues which must be considered.  First, are the engines able to be lit while in flight?  How much boil-off of the cryogenic fuel will take place before LV ignition?

For the LV proposal, I selected the J2-X engine.  The J2-X can be lit while in flight, aka "Air-start"; it also has a higher ISP than the RS-68 engine or RS-68A engine as used on the Delta IV.  The SSME would have an even higher IPS, but the SSME is not an "Air-start" engine.

As for Boil-off, the rate depends on the insulation, tank diameter, and time.  It is an issue which would need careful analysis.  I can think of a few ways to mitigate boil-off, but at this time I am really interested in knowing whether the concept is feasible.

Those J2-X engines are pretty expensive to throw away, it would make more sense if this were a flyback stage...

With air-launch, it's a lot easier to get booster RTLS without having to pay a big penalty: http://selenianboondocks.com/2008/09/orbital-access-methodologies-part-vi-air-launched-glideforward-tsto/

Booster flyback doesn't need much fuel though.

However for point-to-point transport air-launch might be very interesting...

[Patchouli]

Replace the three RL-10s with a single BE-3 the ISP is lower but it's thrust is much higher so you only need one and it's a lot cheaper.
Though on the cost of the J-2X 25 million was with a low production for Ares would have needed.
If it was used in a rocket that flew more often it might be cheaper.
I wonder could the hybrids in DC be replaced with two Chase-10 or Xcor methane engines allowing it to effectively be it's own upper stage?

Then you'd have to take your upper stage tanking through re-entry and landing.  There's no room in the current DC design for that.

I was mostly thinking in acting as a third stage with DC it's self supplying  about 1000M/sec  of the delta V vs replacing the second stage to compensate for the lower ISP of the BE-3 vs RL-10.
Though advanced hybrids or even really good hypergolics may have enough performance to make this unnecessary.

To eliminate the second stage you could make use of drop tanks like the Soviet MAKS concept.

In that case the drop tank would be much smaller maybe around 1/3rd the size of the one on MAKS esp if the Chase-10 is used instead of the RL-10 for the integrated engine since methane is more dense.

It's kinda backwards of what would be ideal as the lower ISP propellant should be in the first stage.
Though this arrangement is not unlike the Antares rocket and the Delta II which have a higher performance first stage then their second stages.

[TomH]

Is it even possible to carry and launch a cryogenic vehicle from this platform?

That's a good question.  From what I understand, there are 2 issues which must be considered.

Carry is the optimal word to me. In my mind's eye, this is what I see. You suspend that thing horizontally from the single suspension point. You then begin loading liquid prop. The fragile fuselage begins to sag at both ends from the weight. Well before the tanks are full, the thing just rips apart. Even if you manage to avoid that, you take off and the vibrational harmonics finish the job.

[Todd Martin]

Just a couple points:

I estimated 419 ISP for the J2X rather than the 448 ISP listed at vacuum since the engine will light around 30,000 feet.  At that altitude, air pressure is about 25% sea level.

The Stratolaunch carrier aircraft is designed to support about 500,000 lbs.  In the photos, the attachment is along the central wing which is ~ 26 feet wide.  If the moment arm becomes un-manageable due to LV length, then a 2nd attachment point further toward the tail could be made but the aircraft designers have not felt that to be necessary so far.

[TomH]

Just a couple points:

I estimated 419 ISP for the J2X rather than the 448 ISP listed at vacuum since the engine will light around 30,000 feet.  At that altitude, air pressure is about 25% sea level.

The Stratolaunch carrier aircraft is designed to support about 500,000 lbs.  In the photos, the attachment is along the central wing which is ~ 26 feet wide.  If the moment arm becomes un-manageable due to LV length, then a 2nd attachment point further toward the tail could be made but the aircraft designers have not felt that to be necessary so far.

It's not the weight the attachment point can carry, it's whether or not a super thin skinned fuselage can support the leveraged weight of the prop. I did not say the attachment point would fail. I am suggesting the fuselage itself would rip apart when suspended from the center horizontally.

[ChrisWilson68]

Replace the three RL-10s with a single BE-3 the ISP is lower but it's thrust is much higher so you only need one and it's a lot cheaper.
Though on the cost of the J-2X 25 million was with a low production for Ares would have needed.
If it was used in a rocket that flew more often it might be cheaper.
I wonder could the hybrids in DC be replaced with two Chase-10 or Xcor methane engines allowing it to effectively be it's own upper stage?

Then you'd have to take your upper stage tanking through re-entry and landing.  There's no room in the current DC design for that.

I was mostly thinking in acting as a third stage with DC it's self supplying  about 1000M/sec  of the delta V vs replacing the second stage to compensate for the lower ISP of the BE-3 vs RL-10.
Though advanced hybrids or even really good hypergolics may have enough performance to make this unnecessary.

Ah, thanks for the clarification.  I thought you meant it would be the second stage.

To eliminate the second stage you could make use of drop tanks like the Soviet MAKS concept.

In that case the drop tank would be much smaller maybe around 1/3rd the size of the one on MAKS esp if the Chase-10 is used instead of the RL-10 for the integrated engine since methane is more dense.

It's kinda backwards of what would be ideal as the lower ISP propellant should be in the first stage.
Though this arrangement is not unlike the Antares rocket and the Delta II which have a higher performance first stage then their second stages.

That's not entirely unlike what the shuttle did, so it certainly can be done.  There is some appeal to dropping just the tank and bringing back the engines.

[Todd Martin]

Regarding the likelihood of an LH2 tank rupturing at the attachment point to the Stratolaunch aircraft, certainly a whole new LH2 tank would need to be designed for this application.  I did a little research and found some of the puzzle pieces to see the size of the issue.

The contact area of a 5 meter fairing (197 inches) to a 26 foot wide wing (312 inches) is 61,464 square inches.  The LV weight of 500,000 lbs, so the pressure applied over this area from the weight would be 8.13 psi.  Not a ridiculous number.

The Shuttle ET was pressurized around 32 psi, a 5 meter tank should be able to go higher, perhaps as high as 50 psi.  The pressure inside the tank helps provide stiffness to the pressure body.

Buckling pressure can be predicted by measuring tank stiffness.  There is a linear relationship between measured stiffness and measured buckling pressure.  Tank stiffness increases with the square of the wall thickness.  So, if all else is equal and you need to double the tank stiffness, then the tank wall thickness would increase by the square root of 2.

Current tanks are primarily aluminum-copper or aluminum-lithium.  Since Stratolaunch has to develop a whole new tank anyway, they could choose to go with a composite tank.  Boeing has demonstrated some significant advances in composite tank technology for LH2:  http://www.parabolicarc.com/2014/04/05/composite-tank/ and they claim a 30% weight savings & a 25% cost savings to a traditional aluminum tank.

[RanulfC]

Is it even possible to carry and launch a cryogenic vehicle from this platform?

That's a good question.  From what I understand, there are 2 issues which must be considered.

Carry is the optimal word to me. In my mind's eye, this is what I see. You suspend that thing horizontally from the single suspension point. You then begin loading liquid prop. The fragile fuselage begins to sag at both ends from the weight. Well before the tanks are full, the thing just rips apart. Even if you manage to avoid that, you take off and the vibrational harmonics finish the job.

Boeing did some study of air-launching a Delta-IV by pretty much wrapping the "aircraft" airframe 2/3rds of the way around the LV so as not to have to significantly redesign the Delta-IV. (Keep in mind though the "real" reason for the exercise was to pitch replacing the Delta-IV with a cargo pod as a means to try and sell a new heavy lifter aircraft design )

In the real world though what you'd have to do is create a stong enough LV "frame" and integrated aircraft "strong-back" to carry the load which of course is why the folks doing the actual design have tossed out liquid propellant entirely in favor of solid since they are structurally far more robust than any liquid tankage can be.

The Shuttle ET, strong-points and SRB connections would be the place to start for such an arrangement I'd think.

Randy

[Todd Martin]

Boeing did some study of air-launching a Delta-IV by pretty much wrapping the "aircraft" airframe 2/3rds of the way around the LV so as not to have to significantly redesign the Delta-IV. (Keep in mind though the "real" reason for the exercise was to pitch replacing the Delta-IV with a cargo pod as a means to try and sell a new heavy lifter aircraft design )

In the real world though what you'd have to do is create a stong enough LV "frame" and integrated aircraft "strong-back" to carry the load which of course is why the folks doing the actual design have tossed out liquid propellant entirely in favor of solid since they are structurally far more robust than any liquid tankage can be.

The Shuttle ET, strong-points and SRB connections would be the place to start for such an arrangement I'd think.

Randy
[/quote]
Randy, that's a great point that it would be preferable to have the aircraft provide as much of the structural support needed as possible, like a strong-back.  A stratolaunch aircraft designed from the start to support a hydrolox stage may differ significantly from what they built.  The difference is a performance margin for the 1st stage.

As for using a drop-tank and DC as the second stage, it would remove the expense of 2 expendable RL10B 2nd stage engines, but I am unclear how it would be arranged.  You can't put the drop tank behind DC because the DC engine nozzles would impinge on the drop tank.  You can't put the drop tank in front of DC and be "pushed" due to plumbing issues.  That leaves a side-mount design, like the Shuttle.  The ET of the shuttle could be considered a drop tank and the MAKS concept shown above is a side-mount.  In this case, the 1st stage booster is limited to 5 meters due so that the aircraft landing gear can reach the ground.  A sidemount drop tank & DC would need to fit on top of a 5 meter booster without a fairing.  Could that be acceptable from an aerodynamic point of view?   

[jongoff]

Is it even possible to carry and launch a cryogenic vehicle from this platform?

That's a good question.  From what I understand, there are 2 issues which must be considered.

Carry is the optimal word to me. In my mind's eye, this is what I see. You suspend that thing horizontally from the single suspension point. You then begin loading liquid prop. The fragile fuselage begins to sag at both ends from the weight. Well before the tanks are full, the thing just rips apart. Even if you manage to avoid that, you take off and the vibrational harmonics finish the job.

Seriously? Better tell the Launcher One guys to call everything off. There's obviously no way to design a liquid fueled rocket that can be air-launched without snapping in two. </sarcasm>

~Jon

[NovaSilisko]

Boeing did some study of air-launching a Delta-IV by pretty much wrapping the "aircraft" airframe 2/3rds of the way around the LV so as not to have to significantly redesign the Delta-IV. (Keep in mind though the "real" reason for the exercise was to pitch replacing the Delta-IV with a cargo pod as a means to try and sell a new heavy lifter aircraft design )

I... would be very much interested in learning more about this study. I have never heard of this before.

[TomH]

Is it even possible to carry and launch a cryogenic vehicle from this platform?

That's a good question.  From what I understand, there are 2 issues which must be considered.

Carry is the optimal word to me. In my mind's eye, this is what I see. You suspend that thing horizontally from the single suspension point. You then begin loading liquid prop. The fragile fuselage begins to sag at both ends from the weight. Well before the tanks are full, the thing just rips apart. Even if you manage to avoid that, you take off and the vibrational harmonics finish the job.

Seriously? Better tell the Launcher One guys to call everything off. There's obviously no way to design a liquid fueled rocket that can be air-launched without snapping in two. </sarcasm>

~Jon

Thank you for such a pleasant reply Jon. Please look back in my post and see if you see where I wrote,  There's obviously no way to design a liquid fueled rocket that can be air-launched without snapping in two. Obviously a liquid fueled rocket can be designed for air launch, but not all rockets are the same. In this case, the OP hypothesizes about the use of Stratolaunch and a launcher large enough to place a full sized DC into LEO. To do that, you need a fusilage that is much stronger than a vertically launched rocket, particularly with regard to a rocket of this length given the leverage such a length will generate. So you end up adding mass to your LV. Building a strongback to support your LV adds mass and drag to your aircraft platform. All of this cuts into performance. I have not done any strength calculations, however anyone with basic understanding of physics should realize that horizontally air launching a liquid fueled rocket of this size from a single suspension point starts to become problematic.

Here's bidding you a good day.

[jongoff]

Thank you for such a pleasant reply Jon.

Apologies for the snark. It just seemed like such a silly issue.

Please look back in my post and see if you see where I wrote,  There's obviously no way to design a liquid fueled rocket that can be air-launched without snapping in two. Obviously a liquid fueled rocket can be designed for air launch, but not all rockets are the same. In this case, the OP hypothesizes about the use of Stratolaunch and a launcher large enough to place a full sized DC into LEO. To do that, you need a fusilage that is much stronger than a vertically launched rocket, particularly with regard to a rocket of this length given the leverage such a length will generate. So you end up adding mass to your LV. Building a strongback to support your LV adds mass and drag to your aircraft platform. All of this cuts into performance. I have not done any strength calculations, however anyone with basic understanding of physics should realize that horizontally air launching a liquid fueled rocket of this size from a single suspension point starts to become problematic.

Here's bidding you a good day.

The thing to remember is that pressurized structures tend to be extremely stiff in bending. But yeah, you'll need to make sure you design the launch supports and rocket internal structure to react the predicted loads (plus margin) reasonably well. If you do it poorly it won't work, but I don't think anyone who has ever looked at air-launched rockets has thrown in the towel for this reason. Rockets already experience rather high bending loads from aerodynamic forces during flight. I haven't run any specific numbers either, but I'd be really surprised if this were the show-stopper you think it is. But the loads probably will drive you toward a more limited set of design options than a vertically ground-launched design.

Apologies again for being rude before. I need to keep the flame without sharing it quite so often.

~Jon

[RanulfC]

Boeing did some study of air-launching a Delta-IV by pretty much wrapping the "aircraft" airframe 2/3rds of the way around the LV so as not to have to significantly redesign the Delta-IV. (Keep in mind though the "real" reason for the exercise was to pitch replacing the Delta-IV with a cargo pod as a means to try and sell a new heavy lifter aircraft design )

I... would be very much interested in learning more about this study. I have never heard of this before.

It wasn't a generally available "concept" that I've found. I found it initially in a book/DVD that my wife bought me that discussed "future" launch vehicles that was published around 2010 IIRC. Illustration mentioned/showed the AC and LV in flight and launching and then a smaller illustration of the LV replaced with a cargo/passenger pod. At the time I found a couple of mentions on-line about it being pitched to the AF as a "responsive space" LV concept and a couple of dozen more that mentioned it as being pitched to the AF as a heavy lift cargo aircraft concept. Then nothing. Short search today has turned up nothing but I expected that. Once I get home I'll see if I can get some links to at least the illustrations

Randy, that's a great point that it would be preferable to have the aircraft provide as much of the structural support needed as possible, like a strong-back.  A stratolaunch aircraft designed from the start to support a hydrolox stage may differ significantly from what they built.  The difference is a performance margin for the 1st stage.

Well in all honesty the carrier aircraft DOES in fact provide some of that as the "stong-back" or interface provides an extensive support structure to transfer carry-loads between AC and LV. TomH's point as I understood it (and as he confirms later) is that given a need for a certain size of tankage to achieve the goal, and especially using LH2, the overall size and length of the tankage is going to require quite a bit of structure ON THE LV to achieve which is very likely going to have a large adverse effect on the payload capability of the LV. And it becomes a spiral as you need to add more propellant and therefore bigger tanks and therefore more mass for support...

LauncherOne is small and starte OUT more compact and structurally "solid" and of course the current StratoLaunch vehicle is "very" solid

Anyone who's done any air-travel will recall "bumping" down the taxi or runway at some point. You may not feel it in the bigget jets but those "bump" transits are still there and they would be transfered to the LV  and most LVs are simply not designed to handle those kind of pressure loads coming from point-sources like those of a carry-and-release mechanism. So your LV ends up being built more like an aircraft and less like a rocket to handle those loads. Jon is right though that pressure stabilization can do a lot, look at the AirLaunch LV which was pretty much "dumped" off the carrier and ramp But during most of the flight and transport the MAJORITY of the load bearing was the carrier platform NOT the LV which is not how the StratoLaunch aircraft works where the majority of the loading is on the LV all the time.

It CAN be done, I mean look at the bending loads the Shuttle ET was subjected to, but its not going to be a "simple" set of tankage and vehicle to accomplish

As for using a drop-tank and DC as the second stage, it would remove the expense of 2 expendable RL10B 2nd stage engines, but I am unclear how it would be arranged.  You can't put the drop tank behind DC because the DC engine nozzles would impinge on the drop tank.  You can't put the drop tank in front of DC and be "pushed" due to plumbing issues.  That leaves a side-mount design, like the Shuttle.  The ET of the shuttle could be considered a drop tank and the MAKS concept shown above is a side-mount.  In this case, the 1st stage booster is limited to 5 meters due so that the aircraft landing gear can reach the ground.  A sidemount drop tank & DC would need to fit on top of a 5 meter booster without a fairing.  Could that be acceptable from an aerodynamic point of view?   

Well I wasn't actually suggesting using a drop-tank/ET configuration as much as I was stating that a good starting point for the TANKAGE design would be the Shuttle-ET as it had to handle ALL its forces in an in-direct manner
But aerodynamically it probably COULD be done... Sort of...
http://www.astronautix.com/lvs/shuls200.htm
http://www.astronautix.com/lvs/staipper.htm
http://www.pmview.com/spaceodysseytwo/spacelvs/sld019.htm

Which would require a HUGE redesign of the DC I'd think to accomplish, but could be done. Yes you're "pushing" the tanks but the plumbing really is pretty straight-forward if you do it that way and one thing I NEED to point out to keep in mind here:

Your "configuration" HAS to be "aerodynamic" because a requirement in the design (StratoLaunch) is that the LV has to be capable of doing its OWN "pull-up" from level to near vertical flight.
(Actually unlike most "top-launch" plans where the carrier AC does a pull-up and "push-over" for release, bottom launch tends to have a begining "negative" AoA due to the release of the LV and it dropping away from the carrier AC. As far as I know only the AirLaunch "Lanyard-and-trapieze" system which orientated the LV before full release and the Crossbow concept where the LV engine was ignited prior to release to allow the carrier AC to bear the full aerodynamic loading didn't have this "problem" )

A very robust airframe structure is required to perform this and the LV has to have enough aerodynamic stability and structural margin to "glide" then take full engine thrust followed by a pull-up manuever to the proper AoA. Typical LV design tends to VERY much "skimp" on structure that is not required and typically no LV is going to experiance ANY loading of a similar nature so the "margin" simply isn't there. "Hanging" from an interface-structure or stong-back is going to also move most of your "loading" for transit loads to being spread among a number of "hold-and-release" mechanical supports (more is better) but those are still going to be "point" load sources that the LV structure itself is going to have to spread and distribute throughout the vehicle.

I highly suspect (HMX "might" chime in here ) that was the reason AirLaunch advocated a pressure-fed (and therefore stabilized) LV for their concepts despite the lower performance versus a turbopump design. The pressurized design probably required much less mass for structure than the non-pressure-fed design.

Randy

[RanulfC]

Please look back in my post and see if you see where I wrote,  There's obviously no way to design a liquid fueled rocket that can be air-launched without snapping in two. Obviously a liquid fueled rocket can be designed for air launch, but not all rockets are the same. In this case, the OP hypothesizes about the use of Stratolaunch and a launcher large enough to place a full sized DC into LEO. To do that, you need a fusilage that is much stronger than a vertically launched rocket, particularly with regard to a rocket of this length given the leverage such a length will generate. So you end up adding mass to your LV. Building a strongback to support your LV adds mass and drag to your aircraft platform. All of this cuts into performance. I have not done any strength calculations, however anyone with basic understanding of physics should realize that horizontally air launching a liquid fueled rocket of this size from a single suspension point starts to become problematic.

Uhm point Tom? No one has proposed or suggested that a "rocket of this size" be suspended from a single point and no concept I've ever seen has suggested such either which is why (I'm sure) Jon snarked. Your argument misses that a larger strong-back/carry assembly is the BASELINE for the StratoLaunch design. Yes this effects carrier AC performance but that's already taken into account. (What I didn't see taken into account in the original design was how much stress this was going to put on the third, liquid, stage as all its loading was going to be carried through the second stage mounting AND internal structure which was going to be pretty large even just aerodynamically due to the size difference in the stages) I suspect that a larger one would be needed for a full LH2 design but what you lose in performance for your carrier AC is probably offset by the increased overall performance of the LV itself.

I believe I understand what you're trying to point out (see above) but its not quite as bad as you believe it would be due to the nature of carried payloads. As I noted on the other hand though, by their nature 'carried payloads' that are not fully supported by the carrier aircraft (top or internal carry as opposed to bottom carry) have to provide a good margin of self support and load bearing but its not hopeless from the start or their wouldn't be any suggestions of this type in the first place

Randy

[Patchouli]

Well I wasn't actually suggesting using a drop-tank/ET configuration as much as I was stating that a good starting point for the TANKAGE design would be the Shuttle-ET as it had to handle ALL its forces in an in-direct manner
But aerodynamically it probably COULD be done... Sort of...
http://www.astronautix.com/lvs/shuls200.htm
http://www.astronautix.com/lvs/staipper.htm
http://www.pmview.com/spaceodysseytwo/spacelvs/sld019.htm

Going with hypergolic ,kerosene,or methane engines in the vehicle would reduce the physical size of the tanks.

[Todd Martin]

Let's apply the Rocket Equation to my proposal to see how it much margin there is left over to play with (LV figures are at the beginning of the thread).

Mass Ratio
  MR = (Mpt/Me)+1
  where,
     Mpt = Mass of the Propellant
     Me = Mass of rocket empty of propellant

The mass ratio of the 1st stage is,
  MR1 = (169883kg/22900kg+2780kg+20410kg+11300kg)+1
  MR1 = 3.96015

The mass ratio of the 2nd stage is,
  MR2 = (20410kg/(2780kg+11300kg)+1
  MR2 = 2.44957

As stated earlier, the aircraft is traveling at 250 m/s

Ve = Exhaust Velocity = ISP * 9.806 m/s

The Exhaust Velocity of the 1st stage is:
  Ve1 = 419 * 9.8066 m/s
  Ve1 = 4108.9654 m/s

The Exhaust Velocity of the 2nd stage is:
  Ve2 = 470 * 9.8066 m/s
  Ve2 = 4609.102 m/s
 
Delta-V = Ve1  * ln(MR1) + Ve2 * ln(MR2) + DeltaVstrato
Delta-V = 4108.9654 m/s * ln(3.96015) + 4609.102 m/s * ln(2.44957) + 250m/s
Delta-V = 10034.45 m/s

There are drag & gravity losses to reach Low Earth Orbit (LEO).  Normally, this is 1500 to 2000 m/s.  In this case we are beginning launch at 30,000 feet so these losses would be reduced by an estimated 600 m/s.  An average total drag & gravity loss for Stratolaunch would be (1500 m/s + 2000 m/s) / 2 - 600 m/s = 1150 m/s 

Net Delta-V to LEO = 10034.45 m/s - 1150 m/s = 8884.45 m/s

The International Space Station (ISS) is orbiting at 7660 m/s.

Delta-V margin = 8884.45 m/s - 7660 m/s = 1224.45 m/s

[Patchouli]

DC also has a fair amount of delta-V it's self not sure how much but I think it's much more then Soyuz and Dragon but less then Orion so it's effectively a third stage.