Reusable LOX/Kerosene SSTO with drop tanks

elderflower · 2016-10-11 03:45:29

Carbon/epoxy is unforgiving stuff. There is almost no energy dissipation prior to failure. It stores a lot of energy in elastic deflection and on failure this is released explosively. The utmost care is needed in laying it up and still we get failures from time to time.

RGClark · 2016-10-11 07:32:07

GW Johnson wrote:

Bob Clark:
The very best you can get is T/Tplain = 1.4. It is seemingly not possible to exceed the 1.4, but it is very possible to do worse. I saw this in a report about unsteady augmentors associated with a pulsejet from long, long ago.
For the unsteady case, T/Tplain was nearer 2. At least at the time of the reports (1959-1964), no one had ever, ever shown T/Tplain > 1.4 in steady flow. Factor 1.1-1.2 is reasonably likely, though.
Sorry.
GW

I only heard of "unsteady flows" in the context of detonations, whereas steady flows would be for example what you get from rocket nozzles. How was unsteady flow achieved in the context of a ducted rocket?

Bob Clark

Last edited by RGClark (2016-10-11 14:42:22)

GW Johnson · 2016-10-11 12:22:27

Hi Bob:

It was a valveless pulsejet whose thrust was improved by fitting ring-shroud augmenters to the inlet and the exit. Very unsteady pulsed operation, although technically deflagrations, not detonations. The hot gas pulses acted like free-flying pistons, pushing slugs of cold air through. The cold air fills the ring during the inflow phase of the pulsejet tube.

This stuff was done at Hiller Aircraft, run by a fellow name of Lockwood. Hiller got with Snecma in France, which already had a good working example of the plain valveless pulsejet. Time was 1959-1964, mostly US Army funding. They fitted the augmentors and called it a "pulse reactor". It had a U-tube shape to it. Lift propulsion for vertical takeoff was the most-proposed application. Enormous thrust/weight ratio (>>50) and TSFC as low as 0.7, in a 500 lb thruster you could pick up with one hand. TSFC usually nearer 1.0, similar to low bypass turbine performance.

GW

RGClark · 2016-10-11 15:02:57

GW Johnson wrote:

Hi Bob:
...
This stuff was done at Hiller Aircraft, run by a fellow name of Lockwood. Hiller got with Snecma in France, which already had a good working example of the plain valveless pulsejet. Time was 1959-1964, mostly US Army funding. They fitted the augmentors and called it a "pulse reactor". It had a U-tube shape to it. Lift propulsion for vertical takeoff was the most-proposed application. Enormous thrust/weight ratio (>>50) and TSFC as low as 0.7, in a 500 lb thruster you could pick up with one hand. TSFC usually nearer 1.0, similar to low bypass turbine performance.
GW

Those performance numbers sound great for a jet-augmented rocket launcher. The thrust/weight ratio is closer to that of a rocket rather than a jet, where it's typically in the range of 10. The main problem with using jets for low speed, low altitude portion of the flight to orbit was the heavy weight of jets compared to the thrust they put out. But 50 to 1 would be quite good. And a TSFC, the analog for jets to Isp for rockets, that is close to 1.0, similar to that of turbojets, is also excellent. This is typical for jets but far superior to rockets.

Why wasn't it used on rockets?

Bob Clark

GW Johnson · 2016-10-11 17:01:28

Bob:

Like all airbreathers, it's low frontal thrust density compared to rockets of almost any kind. The cycle pressure is low, so the frontal thrust density is low. These things would pick up only light loads, not heavy, like a launch rocket.

All rockets have high pressures, and thus high frontal thrust density, by a couple of orders of magnitude, at least. It depends upon how the chamber and nozzle size out relative to the body size. But for solids whose chamber is the body, frontal thrust density is the very highest, in spite of lower Isp. That's why launch booster strap-ons are all solids. Their Isp is why they stage off before the liquid core first stage burns out.

GW

RGClark · 2016-10-30 10:42:28

GW Johnson wrote:

Hi Bob:
It was a valveless pulsejet whose thrust was improved by fitting ring-shroud augmenters to the inlet and the exit. Very unsteady pulsed operation, although technically deflagrations, not detonations. The hot gas pulses acted like free-flying pistons, pushing slugs of cold air through. The cold air fills the ring during the inflow phase of the pulsejet tube.
This stuff was done at Hiller Aircraft, run by a fellow name of Lockwood. Hiller got with Snecma in France, which already had a good working example of the plain valveless pulsejet. Time was 1959-1964, mostly US Army funding. They fitted the augmentors and called it a "pulse reactor". It had a U-tube shape to it. Lift propulsion for vertical takeoff was the most-proposed application. Enormous thrust/weight ratio (>>50) and TSFC as low as 0.7, in a 500 lb thruster you could pick up with one hand. TSFC usually nearer 1.0, similar to low bypass turbine performance.
GW

Another proposal for an altitude compensating engine, GW:

Altitude compensation attachments for standard rocket engines, and applications, Page 4: the double aerospike.
Nozzle%2B-%2Bdouble%2Bspike%2B2.jpg
http://exoscientist.blogspot.com/2016/1 … s-for.html

A question I have about it though is whether it is correct that only requiring an aerospike to expand from sea level pressures, instead of full combustion chamber pressures, to vacuum results in a shorter, slimmer and therefore lighter aerospike.

Bob Clark

Last edited by RGClark (2016-10-30 10:43:16)

GW Johnson · 2016-10-31 08:59:44

Hi Bob:

I don't know why in theory your partial aerospike idea wouldn't work. In practical terms, hanging onto a piece of hardware in the middle of a hot, very supersonic stream will be, shall we say, difficult.

If you go look at "Ramjet Cycle Analysis" posted in Dec 2012 over at "exrocketman", Figure 3 is a sketch of a somewhat-similar idea. It was a rotatable paddle located in a ramjet throat, which changed both throat area and expansion ratio. We tested this, including an 800 sec burn in full size in a direct-connect ramjet ground test.

With the paddle crossways to reduce throat area, the expansion was a semi-free expansion against the wake zone of the paddle. It was also still expanding against the bell. We learned the hard way in cold-flow tests that good thrust efficiency required that the paddle wake must close before it reached the physical bell exit.

The other thing we learned was this thing would only survive at ramjet pressures. None of the materials would survive at rocket chamber pressures. They were just too hot and weak.

Try turning your geometry inside-out. The difficulty is the hand-of-God required to hold the thing in position. Don't do that. How about some pigtail surfaces hanging from the trailing edge of the bell? Or discrete chambers and partial bells on either side of an aerospike surface supported from not within a supersonic stream?

GW

Last edited by GW Johnson (2016-10-31 09:00:51)

RGClark · 2017-02-04 22:52:33

GW Johnson wrote:

Putting a ring shroud around a rocket engine can induce airflow through the shroud that increases thrust, yes. At very subsonic to static speeds. If the friction of the airflow past the rocket engine can be reduced, you might get factor 1.4 more thrust statically, less as your speed increases. Smooth shapes and proper venturi internal profiles are required to make this work at all. The ducted propellor is another example of the same device with a different prime mover. Same sort of airspeed restrictions apply. Static to about 200 mph.
The air ejector pump takes this to extremes by going to a geometry that is a whole lot more complicated (and heavy) than a simple venturi shroud ring. It only works statically, and nobody is interested in its thrust, only the pressure increase it can supply without any moving parts. As a pump, its efficiency is very low. There's a whole lot more massflow coming from very high pressure in the driving jet, than any low pressure airflow it can induce. The thing essentially works by fluid friction between the two streams, which is inherently wasteful. Applications I am familiar with include taking a rocket propellant mix bowl to hard vacuum conditions for mixing without air entrainment.
I'm not at all sure this thing would ever be worth the extra weight to add it to a launch rocket. Using the skirt extension for your shroud ring has the incorrect venturi geometry, you will not get much of the static thrust multiplier of 1.4.
GW

Another method to increase thrust at sea level is the "thrust augmentation nozzle", TAN. It's sort of like an afterburner for rocket engines. What it does is inject propellant into the nozzle and ignites it to generate additional thrust. See discussion here:

Thrust Augmented Nozzles
Posted on November 12, 2007 by Jonathan Goff

http://selenianboondocks.com/2007/11/th … d-nozzles/

In experiments the researchers were able to increase thrust by up t0 70%.

The researchers also studied theoretically the case of using RP-1 to inject into the nozzle of a hydrolox engine. They were studying how payload could be increased for a SSTO. Remarkably the payload went up by nearly 3 times, from 25,000 lbs to 70,000 lbs using TAN. See Table 1 here:

THRUST AUGMENTED NOZZLE (TAN) the New Paradigm for Booster Rockets (Preprint).
http://www.dtic.mil/cgi-bin/GetTRDoc?AD … tTRDoc.pdf

I'm wondering if we can get even better performance if we bring outside atmospheric air into the nozzle to burn with the fuel rather than the oxidizer. This would increase Isp if it would work.

I'm thinking of a scenario like this:

Rocket motor thrust nozzle with means to direct atmospheric air into the interior of the nozzle.
US 3469787 A.
shuttered%2Bnozzle%2B2.jpg
https://www.google.com/patents/US3469787

We would open up vents on the nozzle to allow air to flow in, then burn it with the fuel. We would have to insure the vents we opened were further down on the nozzle so that the reduced pressure of the exhaust flow further down would allow the atmospheric air to enter in. We also don't want after we ignite the fuel with the air for this exhaust to exit back out the vents, further reason for making the opened vents to be further down the sides of the nozzle.

Bob Clark

Last edited by RGClark (2017-02-05 07:20:05)

GW Johnson · 2017-02-07 10:54:51

Hmmmmm. I dunno, Bob.

Vents in the bell of an engine mounted in the wake of a stage would see essentially whatever is the local atmospheric pressure as the source for the air. For air to flow in, the pressure inside the bell at the vent location must be less, or else hot gas flows out through the vent instead.

For a curved bell, we're talking wall local inside pressure less than atmospheric, and it will in turn be higher than bulk average in the bell at that location, because of the local recompression influence of the recurving bell shape. That tells me the vent idea might actually work better on a straight conical bell instead of a curved bell.

That pressure difference effect means you can only have open vents when the nozzle is overerexpanded, which is otherwise a backpressure-on-the-exit plane thrust reduction anyway. Running too severely overexpanded risks shock separation in the bell. Not easy to predict accurately, but I do have a crude correlation if you want it. Separation limits just how overexpanded you can be.

Once off the deck, the risk with most of these engine designs (liquid or solid) is underexpansion, not overexpansion, assuming no changes to chamber pressure. So the bell vent thing sort of resembles a solution looking for a problem to solve, to me. (A lot of patents are exactly that sort of thing.)

As critical as max thrust is to a vertical launch rocket at liftoff, most first stage engines should just be designed for perfect expansion to sea level backpressure, so that CF (and so thrust for a given Pc) is maximum at liftoff. Overexpansion reduces thrust at that condition. Underexpansion doesn't reduce thrust, but does cost you Isp.

As you climb, the bell is increasingly underexpanded, and so not as efficient as it could be (less Isp than potentially possible), but at least thrust increases because (Pe-Pback)Ae is always positive.

In effect, once you have selected a Pc, the ratio of that to sea level ambient sets your expansion ratio with the Pe = Pamb perfect expansion criterion. That's essentially the very best you can do with a fixed bell at sea level.

GW

GW Johnson · 2017-02-07 11:00:56

Here's an idea for you. Arrange 4+ engines in a circle around an extendible centerbody, somewhat along the lines of Spacex's "octaweb". The centerbody is your free expansion spike for perfect expansion effects once at altitude. I'd make it extendible so it doesn't project past the nozzles on the pad before launch. Light off, lift off, and then extend the spike as you start to rise. Its tip is where you calculate your free expansion thrust performance.

GW

elderflower · 2017-02-09 03:50:00

How would that work with gimballed engines for thrust direction control, GW? Maybe we would need a different method?

GW Johnson · 2017-02-09 11:01:13

Gimbal the spike, not the engines. Or maybe just differential throttling would provide enough effect. I dunno.

The other way to control attitude is fixed engines but use attitude thrusters. That's how the flight control on the old "Scout" satellite launcher worked. It was a 4 stage solid, with inexpensive, lightweight fixed nozzles on those solids.

GW

Last edited by GW Johnson (2017-02-09 11:04:22)

elderflower · 2017-02-09 11:55:37

Attitude thrusters is probably the way. Steering either the spike or the main nozzles is going to give huge imbalanced loads on the spike, overexpanding one side and under expanding the other and with a sideways component of gas velocity past it. Also flow detachment and reattachment might cause massive vibration.
I do like the idea, though. It looks like a good way to vary expansion without complicated variable nozzles.

RGClark · 2017-02-11 00:17:49

GW Johnson wrote:

Here's an idea for you. Arrange 4+ engines in a circle around an extendible centerbody, somewhat along the lines of Spacex's "octaweb". The centerbody is your free expansion spike for perfect expansion effects once at altitude. I'd make it extendible so it doesn't project past the nozzles on the pad before launch. Light off, lift off, and then extend the spike as you start to rise. Its tip is where you calculate your free expansion thrust performance.
GW

I like this and think it might work.

Bob Clark

GW Johnson · 2017-02-11 16:30:17

Hi Bob, long time no talk:

Glad you like the idea. If you have a deep flame pit, you can use an even simpler and lighter fixed spike. Elderflower is right: just go with attitude thrusters. KISS is beautiful, ain't it?

GW

RGClark · 2017-02-15 12:00:00

GW, that DARPA proposal I wrote was not successful. It was the one we were discussing about new methods of thermal protection in 2013 here:

Index » Interplanetary transportation » Reusable Rockets to Orbit.
2013-10-10 16:47:00
http://newmars.com/forums/viewtopic.php … 19#p117119

This discussion was from 2013. I finally got around to writing the proposal in 2015.

However, what DARPA really wants now in regards to launch access is low cost flights for small payloads, possibly using a reusable booster. The thermal protection issue was not key to that. But what is still a key question and what doomed the X-33 was the inability to get lightweight conformal tanks.

I took a look again at your blog post:

Sunday, October 6, 2013
Building Conformal Propellant Tanks, Etc.

Done successfully, you have a tank only a few percent heavier than a cylinder of the same volume, but not heavier by factors. It will be at least a little bit heavier, that is inevitable. That’s simply the price you must pay for the shape you want. Update 10-7-13: for the same panel thicknesses and weights as cylindrical construction, a lower-bound estimate of the weight growth factor is the perimeter length ratio, computed from cross-section views.

http://exrocketman.blogspot.com/2013/10 … s-etc.html

I'm still struck by your statement that you can get close to the same weight efficiency for lobed tanks as for cylindrical ones using metal tanks. You mentioned the figure of merit that determines the weight growth is perimeter to length. I had suggested that DARPA look again at the X-33 to fill the role of their desired reusable first stage booster:

Saturday, October 5, 2013
DARPA's Spaceplane: an X-33 version.
http://exoscientist.blogspot.com/2013/1 … rsion.html

Looking at pictures of the X-33's lobed tanks, it does look like the perimeter to length ratio would be low, so you could get close to the usual cylindrical tank weight efficiency:

LOCKHEED MARTIN X-33 COMPOSITE TANK TESTING, HYDROGEN TANK MULTI-LOBED AT K SITE, NASA PLUM BROOK STATION.

If you could provide data that support the validity of your perimeter to length ratio figure of merit or even prove it on a small test tank, you might get funding for this.
Note it's not just DARPA that would need this. The Air Force would also need such weight efficient tanks for their hypersonic vehicles which also have a flat rather than round cross-section:

The HTV-3X Hypersonic Test Vehicle.

Bob Clark

Last edited by RGClark (2017-02-15 21:51:24)

elderflower · 2017-02-15 12:31:05

There are drawbacks here for cryogenic contents. Departure from spherical increases surface area and therefore increases heat gain and insulation requirements. Insulation is not weightless, especially if it gets invaded by water.

GW Johnson · 2017-02-15 13:12:58

Bob:

The lobed tank posting had to do with segments of spherical and cylindrical geometries, not generalized or arbitrary 3-D shapes. That's the fundamental limitation of what I did so long ago. You build the thing just like an air mattress: membranes of constant radius joined at corners which are in turn connected across the pressure zone by straight tension walls. Most of the time you will want vent holes in those walls.

We never built a physical object for test, no one wanted to fund it, even as an IR&D project. 25-30 years ago, the only inquiries about conformal tanks or solid motor cases came in via marketing, and there weren't enough marketing $ available to fund an experiment. So what we did was verify that my by-hand approximation calculations were correct with a 3-D finite-element thermo-structural model. Memory fails, but I believe we verified it with NASTRAN, or else something similar.

The blend radius and dimensions of the corner where membranes meet were nothing but my first guess to try: min section dimension twice the membrane thickness. The finite element model indicated localized yielding right in the crotches of the corner radii, but nothing that would risk failure. We assumed isotropic metal, a steel, I think, but the same basic design approach would work in composites, if you pay very careful attention to fiber directions. Use my recommendations to get started, verify that with FE modeling, then go build with confidence. But always verify finished items by test! Lot acceptance works in metal for that, but with composites you must still verify every single produced item. There's too much variation, still.

Again, the key to this is uniform membrane tension only! There can be no shear or bending, or else you have to thicken it up, which is fatal to flight designs. I first did this as "the only way out" when asked to size an elliptical section rocket motor case for 2000 psig. The damn thing was so thick it was almost as heavy as a solid bar the same size. That flawed "elliptical-is-almost-as-good-as-circular thinking" was the conventional wisdom of the time. Wisdom it was not then, and still is not today.

So I sized a 3-tube air mattress case, which astonished the marketeer who asked me to size the elliptical case. He paid for the FE model, and my air mattress idea was sent to the inquiring customer. We even proposed a 6-tube air mattress flat case for another customer. This included how to fit a nozzle, how to insulate, what shape propellant grains, and how to cast them.

Back in those times, the industry was not yet serious about conformal vessels. Further, the industry has always distrusted radical ideas, even when they are shown to be good. Plus, there is a very serious "not invented here" problem dealing with government labs, whenever something new or radical is involved. (Ask me about the ship-launched radar decoy tale sometime.)

GW

Last edited by GW Johnson (2017-02-15 13:14:11)

Quaoar · 2017-02-16 12:18:11

GW Johnson wrote:

Hi Bob, long time no talk:
Glad you like the idea. If you have a deep flame pit, you can use an even simpler and lighter fixed spike. Elderflower is right: just go with attitude thrusters. KISS is beautiful, ain't it?
GW

Hi GW.

Japanese project Kankou-maru used a fixed spike, like yours:

"Thrust for takeoff is supplied by 12 Mitsubishi LE-9 engines, burning liquid oxygen and liquid hydrogen. 4 of the engines are LE-9B-3 "booster" engines, optimized for low altitude operation. The other 8 engines are LE-9S-3 "Sustainer" engines, optimized for vacuum operation. The vehicle afterbody is designed to use the vehicle exhaust and the atmosphere as an "aerospike" nozzle to increase efficiency at all altitudes.

http://www.spacefuture.com/archive/kank … nual.shtml

Last edited by Quaoar (2017-02-16 12:18:43)

GW Johnson · 2017-02-16 13:33:41

As I said in the posting about composites, the industry does not like new and radical ideas. "Old space" here in the US was going to try a linear aerospike propulsion scheme on the X-30 that got cancelled. I notice they have not proposed anything like that since. Bad taste in their mouth I guess.

That program died more from politics and poor (greedy) management than it did from anything technical, and they did make some technical mis-steps, too.

Although, I myself do not believe single stage chemical is a viable way to reduce launch costs. The energy simply isn't there to do it at a high payload fraction. And you have to squeeze the inert fraction down to the point where it might not even fly once, much less reliably.

GW

RGClark · 2017-02-17 12:08:11

GW Johnson wrote:

Hi Bob:
It was a valveless pulsejet whose thrust was improved by fitting ring-shroud augmenters to the inlet and the exit. Very unsteady pulsed operation, although technically deflagrations, not detonations. The hot gas pulses acted like free-flying pistons, pushing slugs of cold air through. The cold air fills the ring during the inflow phase of the pulsejet tube.
This stuff was done at Hiller Aircraft, run by a fellow name of Lockwood. Hiller got with Snecma in France, which already had a good working example of the plain valveless pulsejet. Time was 1959-1964, mostly US Army funding. They fitted the augmentors and called it a "pulse reactor". It had a U-tube shape to it. Lift propulsion for vertical takeoff was the most-proposed application. Enormous thrust/weight ratio (>>50) and TSFC as low as 0.7, in a 500 lb thruster you could pick up with one hand. TSFC usually nearer 1.0, similar to low bypass turbine performance.
GW

Further about the allowing air into the nozzle to burn with fuel idea: suppose our nozzle is an aerospike. Since it is already open to the air there would be no need to have opening vents on the nozzle.
But by combusting with fuel, we would get a high pressure increase so there is a question if the exhaust gases would still follow the slope of the spike or balloon out. I'm thinking the exhaust gases from the regular combustion chamber would be flowing down at high speed so they would have an effective low pressure. Then if we only added enough fuel to increase to total pressure after the external combustion to sea level level ambient, then the ambient air could still constrain the combustion.

Bob Clark

Last edited by RGClark (2017-02-19 09:09:26)

GW Johnson · 2017-02-18 10:15:49

Hi Bob:

I think if you inject fuel into a free-expansion nozzle plume and burn successfully with the air mixing in on the boundary, it just makes the plume balloon outward more. Whether that actually speeds up the flow for more thrust, well, who knows?

The practical problem is locating the injection hardware. There’s no physical structures anywhere near the mixing layer, except the throat or exit lips upstream. The flammable ratios are reached well downstream, even if you inject hydrogen. And you are relying on spontaneous (and extremely rapid) autoignition.

But this sort of thing is way outside what I had real experience with, so what do I really know? I do know that the combustion models in a lot of CFD codes still leave a great deal to be desired, especially as regards what could burn and what could not.

GW

SpaceNut · 2017-02-18 10:41:12

I see that we are still trying to get something from nothing with the engine, fuel, and nozzles combinations...
The Space x Falcon 9 Heavy and even the Delta IV heavy are as close to a drop tank design that we can manage at this time....

GW Johnson · 2017-02-18 14:16:09

Spacenut:

I'm pretty sure some sort of free expansion nozzle design will provide a significant smidge of extra performance to a launch vehicle. I'm not at all sure that adding other whistles and bells to it will be useful at all. That's what I tried to indicate in my posts above.

Free expansion nozzles are commonly thought of in an axisymmetric geometry, although that is very hard to achieve from a thermo-structural/materials standpoint. Far easier would be a cooled center spike surrounded by a circumferentially-symmetric pattern of separate conventional engines with very short bells.

The bells size for ideal expansion at sea level back pressure. It is as you climb to higher altitudes with lower backpressures that the spike comes into play. You get just a little more thrust out of those plumes expanding against the spike than you do with just conventional underexpanded bells. But it's just a smidge, the conventional underexpanded bell really isn't too bad, just as we do it now.

What you do NOT want is an overexpanded bell. That is a thrust loss right up to the backpressure-induced flow separation point, where it becomes a MAJOR thrust loss. That's just the simple compressible fluid mechanics of supersonic nozzles.

Predicting the critical flow separation backpressure point is inexact at best, but I do have a rough-and-ready correlation for it.

GW

Last edited by GW Johnson (2017-02-18 14:18:20)

SpaceNut · 2017-02-18 14:57:23

http://stem-az.org/resources/4-Rocketry.pptx
Slide 12 talks about the rocket nozzle chamber and on slide 18 it talks about the exhaust plume....
In order to fix the force for changes as we rise from launch to orbit requires more than just a lengthening of the nozzle but also an internal shape change as well. Think parabolic for how a light beam is focused versus unfocussed for the chamber nozzle area and thats what is happening to the exhaust plume.

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Re: Reusable LOX/Kerosene SSTO with drop tanks

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Re: Reusable LOX/Kerosene SSTO with drop tanks

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Re: Reusable LOX/Kerosene SSTO with drop tanks

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Re: Reusable LOX/Kerosene SSTO with drop tanks

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Re: Reusable LOX/Kerosene SSTO with drop tanks

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Re: Reusable LOX/Kerosene SSTO with drop tanks

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Re: Reusable LOX/Kerosene SSTO with drop tanks

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