Micro-Rocketry to Orbit?

[braddock]

I just wanted to pose this as a thought experiment to help me better understand the limits of rocketry.

What is the smallest theoretical rocket that can make orbit?  What factor causes the limit?

As far as I can tell, the Tsiolkovsky rocket equation doesn't seem to address the problem of scale.  If I wanted to launch a paper clip into orbit, and I had a 350 Isp micro-machined thimble engine, then I just plug that in and get the same mass fraction that the big boys use.  But this seems intuitively wrong.

I can imagine that atmospheric drag will scale non-linearly with size, but say I hitch a balloon ride up the first 100k feet to avoid most of that...

So is the problem just building a teeny tiny engine of some exotic type?  Or possibly carrying a very high muzzle-velocity gun up on a balloon to shoot my pea into orbit?

[mlorrey]

Aerodynamic losses are significantly greater, they do scale not just because of the scale of the vehicle size to the depth of the atmosphere, but that air is "thicker" for small things than for large things (a reason why small objects have lower terminal velocities versus large objects of equal density.)

Simultaneously, gravity losses are greater because a useful orbit is also "higher" if your unit measure is the size of the launcher. At the same time and while structural requirements due to gravitational stresses are less, structural requirements based on pressure (tank pressure, pump pressure, combustion chamber pressure) may not scale that well.

[Propforce]

braddock - 9/6/2006  1:11 PM

What is the smallest theoretical rocket that can make orbit?  What factor causes the limit?

Certain things are scalable, certain things are not.  Boundary layer is not scalable.  Surface area accounting for aerodynamic drags, however; is scalable (with the suface area!).  Generally speaking, smaller launcher has a higher surface area relative to its size, hence less efficient than a big rocket.  

As far as I can tell, the Tsiolkovsky rocket equation doesn't seem to address the problem of scale.  If I wanted to launch a paper clip into orbit, and I had a 350 Isp micro-machined thimble engine, then I just plug that in and get the same mass fraction that the big boys use.  But this seems intuitively wrong.

Are you launching it as a single stage? two stages?  air launched? balloon launched?  The rocket equation alone does not tell you what the losses are that adds up to the total "ideal" delta-vee.  You'll have to run a 3-DOF trajectory optimization program (e.g., POST, OTIS) with the correct vehicle aero data in order to get a credible design.  In addition, you'll need to get a good idea of your propellant mass fraction that you can design to.  Again, this is where small launch vehicle suffers, as the smaller the vehicle the lower prop mass fraction one is able to achieve because the "fixed" inert weight becomes a larger % of the total weight.  Then you'll need credible "load-path" and structural strength sizing for your rocket.  Is your rocket going to flex like a spaghetti during flight, or is it going to be rigid like a pencil?  Do you model your guidance, control & flight stability adequately? Too many people fail to appreciate how heavy all those avionic boxes weight on a vehicle, and how heavy the wire harnesses are !!

So is the problem just building a teeny tiny engine of some exotic type?  Or possibly carrying a very high muzzle-velocity gun up on a balloon to shoot my pea into orbit?  

Does your "Pea" need to stay intack once to get to orbit?  Does it burn up on the way up due to high aerodynamic heating?  Can your pea handle the high-G load and the high dynamic preesure?  

Put it this way, "mother nature" will teach you where the "limits" are !

[braddock]

Okay, so far it seems like the show stoppers are:

1) Volume doesn't scale linearly with Surface Area (even if 1/10th the mass may need only 1/10th the fuel, but the scaled rocket will only have ~1/5th the surface area).  Larger relative surface area = larger relative air friction resistance.  

2) Also larger relative surface area MAY require larger proportion of weight to be structural (although lower pressures and smaller stresses could mitigate that)

3) Larger relative air viscosity (air is "thicker").  This is what the Micro UAV folks have run into as well (often to their advantage, although probably not to ours)

I'm still rather interested in the case of a balloon or otherwise boosted start at 100k ft where the atmosphere density is a fraction of ground level.  There must be some reason amateur rocketeers haven't taken a shot at orbit from a balloon (the likes of JP Aerospace and others have done balloon launches)?

What _IS_ the smallest rocket to have ever made orbit?

[mlorrey]

braddock - 9/6/2006  7:48 PM

Okay, so far it seems like the show stoppers are:

1) Volume doesn't scale linearly with Surface Area (even if 1/10th the mass may need only 1/10th the fuel, but the scaled rocket will only have ~1/5th the surface area).  Larger relative surface area = larger relative air friction resistance.  

2) Also larger relative surface area MAY require larger proportion of weight to be structural (although lower pressures and smaller stresses could mitigate that)

3) Larger relative air viscosity (air is "thicker").  This is what the Micro UAV folks have run into as well (often to their advantage, although probably not to ours)

I'm still rather interested in the case of a balloon or otherwise boosted start at 100k ft where the atmosphere density is a fraction of ground level.  There must be some reason amateur rocketeers haven't taken a shot at orbit from a balloon (the likes of JP Aerospace and others have done balloon launches)?

What _IS_ the smallest rocket to have ever made orbit?

Ah, good question. The US NAVY's NOTS program,  which in 1960-62 launched Caleb rockets from underwing an F4D-1 Skyray (not the later F-4 Phantom, different plane model), made 7 launches, with three failed launches. One of the successful launches, according to range radar, had sufficient downrange velocity to attain orbit, however payload telemetry failed, so achievement of orbit could not be confirmed. Project staff assert they made orbit.

There was also the earlier Project Pilot, also put on by the NOTS program, which made 10 attempts to loft a 1 kg payload. All failed, though one was also claimed to have reached orbit, though telemetry also failed in this instance.

The Caleb was a four stage solid fuelled rocket. The goal of the project was to achieve an air launched ASAT weapon. Total Mass: 1,350 kg (2,970 lb). Core Diameter: 0.60 m (1.96 ft). Total Length: 4.90 m (16.00 ft). Payload: 7.00 kg. Empty weight: 224 kg. Fuel fraction:  0.829. Isp (sl): 204 sec. Isp (vac) 250 sec.

Theoretical:
The smallest launcher that never flew, but could have reached orbit, would have been Dr. Gerald Bull's GLO-1B cannon launched missile, http://www.astronautix.com/lvs/glo1b.htm
which he had developed to be launched from his 16 inch gun system at his Carribean island test site. Bull's larger Babylon Gun, to be built for Saddam Hussein had Bull not been assasinated by the Mossad prior to the Gulf War, would have theoretically lofted 1,000 lb into LEO via a missle launched through its 36 inch bore.

Actual confirmed: The Russian Shtil launcher seems to me to be the smallest launcher that has put payloads reliably into orbit. This vehicle is a Sea Launched Ballistic Missile essentially of similar size as SpaceX's Falcon 1 launcher.

[braddock]

Very interesting.  I found an article on Project Pilot that suggests that it actually achieved orbit twice, although the telemetry reports were ify.  Only source of this article seems to be in an old mailing list archive: http://www.uoregon.edu/~stevev/sd-archive/raw/sd1998-3.txt (search for NOTSNIK)

So a 16' x 2' air-launched solid (Caleb) can at least theoretically get a 15 lb payload to orbit, with a payload-to-mass ratio of .5%.  Can't find altitude info on Caleb, but Project Pilot was launched at 41k feet.  

http://www.astronautix.com/graphics/n/nots.jpg">

Interesting lower limit example.  

And I spent a couple hours last night reading about Dr. Bull's Project HARP gun-launched rockets, which was also fascinating.  Astronautix has an excellent series of articles on that.  http://www.astronautix.com/articles/abroject.htm

I still don't have a good idea of what defines the lower limit for an air-launched orbital launch vehicle, however.  Caleb seems to have a reasonable payload mass fraction for a solid (although by my figures Pegasus is closer to 2%, but is also much more refined).

If 2'x16' (=~50 cu ft of propellent if cylindrical) lofts 15 lbs, then I could naively expect that an 8ft x 6 inch rocket could loft a half pound payload (8ft x 6in =~ 1.5 cu ft of propellent =~ (50 cu ft)/30).  

This would seem to quickly cut things down to the capabilities of amateur rocket folks?  But yet I don't see them orbiting small payloads, so I must be missing something still?

[Jim]

Cost of the aircraft

[mlorrey]

Jim - 10/6/2006  9:01 AM

Cost of the aircraft

True. For those "amateurs" with big bank accounts, I know where to get flight capable Mig-21 aircraft, delivered to your location, for under $100,000.00. The Mig-21 has higher speed, altitude, and payload capacity than the Skyray, so one ought to be able to replicate Project Pilot and/or Caleb relatively easily.

So, any takers?

[mlorrey]

braddock - 10/6/2006  8:04 AM

Very interesting.  I found an article on Project Pilot that suggests that it actually achieved orbit twice, although the telemetry reports were ify.  Only source of this article seems to be in an old mailing list archive: http://www.uoregon.edu/~stevev/sd-archive/raw/sd1998-3.txt (search for NOTSNIK)

So a 16' x 2' air-launched solid (Caleb) can at least theoretically get a 15 lb payload to orbit, with a payload-to-mass ratio of .5%.  Can't find altitude info on Caleb, but Project Pilot was launched at 41k feet.  

http://www.astronautix.com/graphics/n/nots.jpg">

Interesting lower limit example.  

And I spent a couple hours last night reading about Dr. Bull's Project HARP gun-launched rockets, which was also fascinating.  Astronautix has an excellent series of articles on that.  http://www.astronautix.com/articles/abroject.htm

I still don't have a good idea of what defines the lower limit for an air-launched orbital launch vehicle, however.  Caleb seems to have a reasonable payload mass fraction for a solid (although by my figures Pegasus is closer to 2%, but is also much more refined).

Well, you've got to decide what your fuel choice is (ergo, your Isp), how many stages (less stages is less complex, but more stages is easier to get payload to orbit), as well as your air launch altitude and speed.

The NF-104 program is a good example of a rocket-equipped jet aircraft doing zoom maneuvers on a consistent, even daily, basis, and is the basis of the various RASCAL proposals. I've read that one attempt was made to launch an upper stage from an NF-104 into orbit, but it failed. The USAF tended to look at the NF-104 with a bit of distain given its concurrence with the X-15 program, and despite the excellent work done by Capt Bob Smith on the program, but it was sabotaged by Chuck Yeager's quest for more glory in his waning years, when he crashed one of the aircraft because he refused to follow Smiths instructions in the flight profile.

I think a reasonable case could be made for a kerosene/lox or kerosene/H2O2 fuelled two stage vehicle to be dropped from a Mig-21 at 40kft and Mach 2. The standard payload of the 21 is 2,000 kg of missiles plus a 23 mm cannon and ammo. If the cannon, radarset, and obsolete avionics were stripped out, you could probably get the max upper stage GTOW up to 4,000-5,000 kg. That should allow an upper stage payload in orbit of about 50-75 kg.

Let the Mig-21 do the climb before release, so a Pegasus-style wing would be unnecessary.

[Zachstar]

Might find this interesting http://www.microlaunchers.com/home.htm

[mlorrey]

Zachstar - 10/6/2006  4:40 PM

Might find this interesting http://www.microlaunchers.com/home.htm

It is interesting what micromachinery is being made. Just the other day I was chatting with a fellow at a company that produces MEMS chip liquid cooling devices, made by diffusion bonding hundreds of leaves of 1-3 mil metal, each with their own micro cut patterns.

[Jim]

mlorrey - 11/6/2006  3:59 PM

Zachstar - 10/6/2006  4:40 PM

Might find this interesting http://www.microlaunchers.com/home.htm

It is interesting what micromachinery is being made. Just the other day I was chatting with a fellow at a company that produces MEMS chip liquid cooling devices, made by diffusion bonding hundreds of leaves of 1-3 mil metal, each with their own micro cut patterns.

The problem is micromachinery doesn't work very well with fluids.  The same reasons that there are no small jst engines, would be the same for turbo pumps.  Reynolds numbers don't scale.

[mlorrey]

Jim - 11/6/2006  6:20 PM

mlorrey - 11/6/2006  3:59 PM

Zachstar - 10/6/2006  4:40 PM

Might find this interesting http://www.microlaunchers.com/home.htm

It is interesting what micromachinery is being made. Just the other day I was chatting with a fellow at a company that produces MEMS chip liquid cooling devices, made by diffusion bonding hundreds of leaves of 1-3 mil metal, each with their own micro cut patterns.

The problem is micromachinery doesn't work very well with fluids.  The same reasons that there are no small jst engines, would be the same for turbo pumps.  Reynolds numbers don't scale.

What makes you think there are no small jet engines? I know a number of modellers who fly jet engines in the 10-100 lb thrust range, and a number of institutions, including MIT, are developing MEMS turbine generators for powering remote devices, many of which are being developed for Army infantry use:

http://www.enme.umd.edu/SSSC/pdf/update/Transducers%2097%20-%20Paper.pdf
http://www.enme.umd.edu/SSSC/pdf/update/HH2000-Turbine.pdf
http://www.enme.umd.edu/SSSC/pdf/update/MEMS%2099%20-%20Paper.pdf

In fact, turbines reach adiabatic efficiencies of 70% and are able to maintain supersonic turbine speeds while maintaining laminar flow, with Reynolds numbers up around 20,000. The one MIT is working on is an 80 watt turbogenerator, with a turbine disc diameter of about 3-4 mm. The full engine assembly is 1 cm dia.

NASA is also working on a "steam" turbine, using Xenon as the working fluid, which is heated by various processes: fuel cells, solar concentrators, radiothermal devices, etc.
http://www.nasatech.com/Briefs/Jan03/NPO20933.html

The reason for this is because at the microscale, gasses flow with the viscosity of liquids at macro scales.

[Jim]

Efficent small jet engines

[cpooley]

Braddock emailed me asking about Microlaunchers.  It is a plan to devekop commercial space access via a launcher to place 1 pound spacecraft to escape.  That was thought do-able 10 years ago.  Now more easily (eg 1-gram RC plane linket to in my site).
1.  starting small because starting "big" has always failed--too expensive, risky
2.  to escape-- LEO too crowded, not enpugh new things to do with that--there's a solar system to check out.  (A ML could orbit maybe 3 Cubesats)
------------------------------
Most of the problems related to small scale have to do with aerodynamisc.  The 1st stage will be inefficient.  I'm assuming 10% GLOW for upper stages.  The upper stages, starting in vacuum, can operate more like "normal" size rockets.
---------------------------
The engineering of this looks very attainable.  Currently raising the funding for the "chapter written in fire and metal".

Charles Pooley  http://www.microlaunchers.com

[braddock]

Super Loki to LEO?
~~~~~~~~~~~~

Okay, as a kind of culmination of what I've leared here, let me see if I can launch a large glass marble (18g) into LEO.

I'd appreciate it if the folks here could point out any flaws in my logic or math here, since this is my first attempt at using the rocket equation.

Fantasizing, I figure I could fit an integrated microchip die with on-board solar cells and an imager and simple lens, and LED for optical communications, and guidance into something the size and weight of a glass marble (18g) pretty easilly with a custom VLSI design.

I take an existing Super Loki sounding rocket, and replace the usual 6 kg instrumented "dart" payload with a second stage small rocket to get me into LEO.  Take off weight is only about 75 lbs.

I assume the Loki will get my second stage above around 200,000 feet; above enough air resistance that I can ignore drag and use the straight rocket equation.

I'll assume an Isp of my solid second stage of only 200 secs.  
This would make my exhaust velocity:

Ve = 200*g0 = 200*(9.8 m/s/s) = 1960 m/s

I'll assume ignition of my second stage at apogee of 200k feet with an initial velocity of 0 m/s.
I'll therefore use a delta V from standstill to LEO of
dV = 9.5 km/s

Mass Fraction: http://upload.wikimedia.org/math/4/d/b/4db710073244697d25c8520051a7bd03.png">

Plugging in the right-hand-side, I get a mass fraction of .992
Ie, 99.2% of my second stage mass has to be propellant for this work.

With a second stage mass of only 6 kg, this means that my stage structure and payload must have a mass of only 47 grams.  18g of that is my marble, leaving 29g for structure.

This is clearly not much margin, but it would seem plausible that not much more than a Super Loki and a solid second stage can get a very small payload to orbit.

Comments and corrections?

[meiza]

You probably don't need 9.5 km/s when starting at high altitude? You're talking about essentially SSTO performance. 8 km/s would give mass fraction of 0.98 and 120 g of empty mass. It would be wiser anyway to get some horizontal velocity from the first stage too and stage at lower altitude while you still have vertical and horizontal speed. (Lessens gravity losses.) So perhaps you can calculate both stages as giving about 5 km/s deltav. The payload of the first stage is the fully fueled second stage, it would prolly be smart to make it heavier than 6 kg. And use three stages.
And how are you going to have maneuvering and guidance in a 29 g or even 120 g empty weight rocket? You have to maneuver to lift the apogee (you can't reach the orbit with a pure gun from ground since the perigee will always be inside earth). Maybe if you shot the Loki straight up and then the upper stage was positioned sideways and then spun up...

[Monroe]

Old topic but maybe we can give it a kick.

Team Prometheus is attempting to orbit a 20 gram payload launched from 100,000ft.

  Our attempt is based on project "Pilot" The exception being balloon launched from a Zero Pressure Balloon. The rocket weigh's about 150lbs, is 5 stages. Is unguided by mechanical means just like Pilot. The third stage fires by horizon location it's spin stabilized. The final stage nozzle is under the noze cone due to the spin when the rocket reaches the other side of the world it will be facing the right direction. The last stage circularizes the orbit.

Monroe

[Tcommon]

As far as I can tell, the Tsiolkovsky rocket equation doesn't seem to address the problem of scale. 

you can scale a rocket, but you can't scale the planet.

Reynolds numbers don't scale.

Reynolds numbers are dimensionless 

similitude can be hard to visualize.

[JohnFornaro]

The SR-71 is my favorite plane, but the F-104C is a close second.  If Chuck Yeager was trying to "sabotage" that program up there, he certainly chose a risky way to do it!

But anyhow, this is a great discussion.  Here is my "affordable" miniature idea.  I'd like to land the smallest possible lunar lander at one of the Apollo sites, photograph it a half dozen times, and successfully transmit the images back to Earth on an open channel.

I lost the envelope, but it seemed at the time that an F1 would do it.

[RanulfC]

As a "note" IF (very large "IF here "I" got lucky at one point) you can find it LLNL did a (Jordin Kare author) presentation on an SSTO-demonstrator sized to put "10kg" into LEO called the "Mockingbird" (AKA: "Bricklifter,")

It used high-performance H2O2/JP8 rockets (8 low-expansion, 1 high-expansion) and was around 6 feet tall.

I can't for the life of me figure out how I found and downloaded the presentation as I've never been able to find it again but hey with more folks trying maybe someone will get lucky again

Randy
Edit: "MARD" was another "key-word" Multi-Application Rocket Drone

[RanulfC]

"Think" I found it, I can't access the site from this computer though:
http://www.quantumg.net/mockingbird.pdf

Randy
Edit: ARrrggghhhh! I "just" noticed it's QuantumG's page.... Ok fess up fella what OTHER information are you holding we should know about!

[RanulfC]

But anyhow, this is a great discussion.  Here is my "affordable" miniature idea.  I'd like to land the smallest possible lunar lander at one of the Apollo sites, photograph it a half dozen times, and successfully transmit the images back to Earth on an open channel.

I lost the envelope, but it seemed at the time that an F1 would do it.

Well...
http://www.spacex.com/FalconLunarCapabilityGuide.pdf

And:
http://forum.nasaspaceflight.com/index.php?topic=13804.0

Are good places to start at any rate

Randy

[RanulfC]

Re-reading the MARD presentation I wonder if something like this isn't the "perfect" experimental Nano-Sat vehicle?

SSTO is hard of course but the small size and simplicty of the vehicle would be a large cost saver. And then think of all the experimental "improvments" that can tried incrimentally on such a vehicle, launch-assist, air-breathing, staging... Heck it's not that much more complicated than many of the high-end, high-power rocketry vehicles...

Randy

[QuantumG]

"Think" I found it, I can't access the site from this computer though:
http://www.quantumg.net/mockingbird.pdf

Randy
Edit: ARrrggghhhh! I "just" noticed it's QuantumG's page.... Ok fess up fella what OTHER information are you holding we should know about!

Great minds.. I was just going to post it.  The pdf originally had some horrible long name and was publicly available on the LLNL website.. it might still be, I don't know.

There's been a lot of technology development since '94.  Aerogel is now everywhere and sheets can be commercially procured.  XCOR has pumpfed engines available off-the-shelf with the performance capabilities that are described.  Nitrous mono-props like Firestar's NOFBX offer greater performance for less complexity. 

The biggest objection to the concept remains the aerodynamics.  They were very optimistic in their numbers and probably need to add another 1 km/s of delta-v, and all the knock-on effects that has.

But there's something incredibly sexy about having a 75 kg rocket which you can throw in the back of a pickup truck and go launch a nanosat to orbit from anywhere, with a fuel truck as the only support vehicle required.

[mmeijeri]

Nitrous mono-props like Firestar's NOFBX offer greater performance for less complexity. 

Greater performance than JP5/H2O2?

[QuantumG]

See Appendix A.  They assume 301s (vac) for the 8 low expansion engines and 327s (vac) for the 1 high expansion engine.  Firestar are claiming 320s (vac) for pressure-fed NOFBX.  So, if you were to pump it, you'd get higher performance. 

If you're wondering how you get 9 km/s or more with that sort of performance and a 1500/75 mass ratio, join the club.

Edit: One possible answer is that the Preliminary Mass Budget has a nominal mass of 70.8 kg.. that goes a long way.  Also funny: the "min" mass is 45.2 kg.

[RanulfC]

Nitrous mono-props like Firestar's NOFBX offer greater performance for less complexity. 

Greater performance than JP5/H2O2?

You can get some idea(s) from the NOFBX monopropellant thread here:
http://forum.nasaspaceflight.com/index.php?topic=24352.0;all

The thread also gets into some small-LV design questions and issues that would be good to review/answer here also...

Presentation is here:
http://www.aiaa.org/pdf/industry/presentations/Greg_Mungas.pdf

Firestar Technologies:
http://www.firestar-engineering.com/NOFB-MP.html
http://www.firestar-engineering.com/NOFBX-MP.html

And a couple of blog posts on the talk Max Vozoff gave at SA-11:
http://www.hobbyspace.com/nucleus/index.php?itemid=28537
http://www.transterrestrial.com/?p=33007

Oh the title of the talk? "Hydrazine-sucks!"

SBIR proposal for SSTO Mars sample return:
http://sbir.gsfc.nasa.gov/SBIR/abstracts/09/sbir/phase2/SBIR-09-2-S3.08-8305.html?solicitationId=SBIR_09_P2

Patent Application for NOFB:
http://www.freepatentsonline.com/y2009/0133788.pdf

And just for the heck of it, FireStar also has a patent for a regenativly cooled engine:
http://www.freepatentsonline.com/y2010/0205933.pdf

Randy

[mmeijeri]

See Appendix A.  They assume 301s (vac) for the 8 low expansion engines and 327s (vac) for the 1 high expansion engine.  Firestar are claiming 320s (vac) for pressure-fed NOFBX.  So, if you were to pump it, you'd get higher performance. 

The JP-5/H2O2 numbers may be low because of the thrust requirements. I doubt NOFBX would still offer higher Isp if it had to deliver the same thrust. Hydrogen peroxide is a more powerful oxidiser than nitrous oxide after all.

[RanulfC]

"Think" I found it, I can't access the site from this computer though:
http://www.quantumg.net/mockingbird.pdf

Randy
Edit: ARrrggghhhh! I "just" noticed it's QuantumG's page.... Ok fess up fella what OTHER information are you holding we should know about!

Great minds.. I was just going to post it.  The pdf originally had some horrible long name and was publicly available on the LLNL website.. it might still be, I don't know.

Probably is THANK YOU for making it SO much easier to find

As I recall I had to put in so many keywords, (including Jodin Kare's name) the FIRST time I found it I had to wade through around 10 pages of Misc-gunk before I managed to stumble across it. This time I managed to find it with ONLY "MARD" and it came right up!

There's been a lot of technology development since '94.  Aerogel is now everywhere and sheets can be commercially procured.  XCOR has pumpfed engines available off-the-shelf with the performance capabilities that are described.  Nitrous mono-props like Firestar's NOFBX offer greater performance for less complexity. 

The biggest objection to the concept remains the aerodynamics.  They were very optimistic in their numbers and probably need to add another 1 km/s of delta-v, and all the knock-on effects that has.

Hmmm, there is always shoving it out the back of a C-130 freighter

But there's something incredibly sexy about having a 75 kg rocket which you can throw in the back of a pickup truck and go launch a nanosat to orbit from anywhere, with a fuel truck as the only support vehicle required.

I LOVE that idea, though to be honest I keep seeing this image of both "trucks" pulling up to some red-neck, middle-o-nowhere gas-station ala-Storm-Hunters with everybody scurring to get fuel while they 'track' an orbital window

Speaking, (somewhat anyway) of the orbital nano-sat market:
http://spacebusinessblog.blogspot.com/2011/05/business-case-for-cubesat-based-earth.html

Is an article by Colin Doughan looking at the possible market for LEO-Nano-Sat imaging constellations...

Randy

[braddock]

Old topic but maybe we can give it a kick.

Team Prometheus is attempting to orbit a 20 gram payload launched from 100,000ft.

I'd love to hear more about your plans with Team Prometheus.

Newer than the origin of this thread is the N-Prize to launch a 10 to 20 gram satellite to orbit for under 1,000 british pounds.  http://www.n-prize.com

Their "press" page has a lot of promising links to a couple Space Shows I look forward to listening to.

Also newer than this thread is a 2009 Google TechTalk with Charles Pooley of Microlaunchers, a frequent poster to our forum. 

[Patchouli]

"Think" I found it, I can't access the site from this computer though:
http://www.quantumg.net/mockingbird.pdf

Randy
Edit: ARrrggghhhh! I "just" noticed it's QuantumG's page.... Ok fess up fella what OTHER information are you holding we should know about!

Great minds.. I was just going to post it.  The pdf originally had some horrible long name and was publicly available on the LLNL website.. it might still be, I don't know.

There's been a lot of technology development since '94.  Aerogel is now everywhere and sheets can be commercially procured.  XCOR has pumpfed engines available off-the-shelf with the performance capabilities that are described.  Nitrous mono-props like Firestar's NOFBX offer greater performance for less complexity. 

The biggest objection to the concept remains the aerodynamics.  They were very optimistic in their numbers and probably need to add another 1 km/s of delta-v, and all the knock-on effects that has.

But there's something incredibly sexy about having a 75 kg rocket which you can throw in the back of a pickup truck and go launch a nanosat to orbit from anywhere, with a fuel truck as the only support vehicle required.

I wonder what the payload would be if the the payload it's self could do part of the orbital insertion burn acting as it's own second stage?

Looking at some SSTO concepts you might be able to double or triple the payload that way.

[sanman]

Regarding scaling factors, etc - what if you don't assume launch at sea-level? What if you assume launch at a higher altitude, even using a propellant with lighter molecular weight?

[RanulfC]

Regarding scaling factors, etc - what if you don't assume launch at sea-level? What if you assume launch at a higher altitude, even using a propellant with lighter molecular weight?

Well, ANY "launch-assist" is going to help in some way. (Hence my comment about dropping it out the back of a C-130 )

Various different fuels could be tried, though using "lower" molecular weight fuel will probably increase the vehicle size which was one reason for they were using the higher-density propellants in the first place.

As usual this reference:
http://www.dunnspace.com/alternate_ssto_propellants.htm

The ubiquitous "Aternate SSTO Propellants" chart shows that using H2O2 as an Oxydizer, you can get payload increases (all things being rather "equal" of course) using several alternate fuels.

H2O2/methylacetylene (usually closer to commercial MAPP gas actually) gives about 16% payload than RP-1. (Which in itself would give a small increase over the suggested JP-8)

H2O2/propylene would yeild around 15% more payload than RP-1 and from an access-and-handling standpoint might be the best choice of all. However there is no "history" on this combination so work would have to be done on engine design and propellant handling.

Probably the "best" from an overall performance standpoint is H2O2/propargyl alcohol, (actually chemically "methyl acetylene alcohol" but that's a nit ) since it would theoretically yeild around a 40% increase in payload over RP-1. However it is toxic and a suspected carcinogen, so you're handling and storage costs go up, and again there is no "history" with a propulsion system designed around these propellants so you're going to be starting from scratch.

Still that gives an idea of some of the possibilites in JUST changing propellants. Other Launch-Assist can't really do anything but help it just depends on how MUCH you want to "help" and how much intergration of that help ends up raising your operations costs.

Randy

[RanulfC]

I wonder what the payload would be if the the payload it's self could do part of the orbital insertion burn acting as it's own second stage?

Looking at some SSTO concepts you might be able to double or triple the payload that way.

"Staging" always helps the payload in some way it's just a question of "if" the added structure, complexity, and other operational issues and costs are "worth" it.

It gets "dicey" to say the least when you start messing with a "proposed-SSTO" since you COULD get lynched for simply suggesting it! (SSTO-Advocates can be SO touchy ) And as Quantum-G noted there is something that latches onto the imagination with a "Pick-Up truck launched SSTO" image.

But staging is always possible the Garvey Nano-Sat Launch Vehicle:
http://www.garvspace.com/NLV.htm

Is to be a TSTO and there doesn't seem to be any technical reason you couldn't add a second stage to the MARD concept-vehicle with some design changes for the "booster" MARD.

Another "idea" concept is the "W" RBCC test vehicle suggested in this report:
http://smartech.gatech.edu/bitstream/handle/1853/8439/aiaa_96-2688.pdf?sequence=1

Where a "booster" stage incorperates H2O2/RP-1 Ejector-Ramjets using H2O2-primary ejector rockets with RP-1 ramjets and a fully rocket "second" stage using a plug-cluster engine.

Or you could go with the original "MARD/Mockingbird" idea and used "clustered" modules to increase your payload fraction

There are a LOT of options to be discussed....

Randy

[RanulfC]

Hmm, looking recently again at the MARD document Quant posted and a certain blog I frequent led to an idea:

According to research the Rocket Engine Nozzle Ejector (RENE) produced 55% more thrust and ISP than a standard rocket cluster during testing. Adding a thrust-shroud to the "standard" Mocking Bird would enable the "RENE-Bird" to perform quite a bit better.
(Probably still have to stage the lower shroud and nose fairing and enter "forward" end first but I'm not sure that's a "bad" thing )

Randy

[NotGncDude]

Old topic but maybe we can give it a kick.

Team Prometheus is attempting to orbit a 20 gram payload launched from 100,000ft.

I'd love to hear more about your plans with Team Prometheus.

Newer than the origin of this thread is the N-Prize to launch a 10 to 20 gram satellite to orbit for under 1,000 british pounds.  http://www.n-prize.com

Their "press" page has a lot of promising links to a couple Space Shows I look forward to listening to.

Also newer than this thread is a 2009 Google TechTalk with Charles Pooley of Microlaunchers, a frequent poster to our forum. 

w00t! Charles Pooley is da man

[NotGncDude]

Jim - 11/6/2006  6:20 PM

mlorrey - 11/6/2006  3:59 PM

Zachstar - 10/6/2006  4:40 PM

Might find this interesting http://www.microlaunchers.com/home.htm

It is interesting what micromachinery is being made. Just the other day I was chatting with a fellow at a company that produces MEMS chip liquid cooling devices, made by diffusion bonding hundreds of leaves of 1-3 mil metal, each with their own micro cut patterns.

The problem is micromachinery doesn't work very well with fluids.  The same reasons that there are no small jst engines, would be the same for turbo pumps.  Reynolds numbers don't scale.

What makes you think there are no small jet engines? I know a number of modellers who fly jet engines in the 10-100 lb thrust range, and a number of institutions, including MIT, are developing MEMS turbine generators for powering remote devices, many of which are being developed for Army infantry use:

http://www.enme.umd.edu/SSSC/pdf/update/Transducers%2097%20-%20Paper.pdf
http://www.enme.umd.edu/SSSC/pdf/update/HH2000-Turbine.pdf
http://www.enme.umd.edu/SSSC/pdf/update/MEMS%2099%20-%20Paper.pdf

In fact, turbines reach adiabatic efficiencies of 70% and are able to maintain supersonic turbine speeds while maintaining laminar flow, with Reynolds numbers up around 20,000. The one MIT is working on is an 80 watt turbogenerator, with a turbine disc diameter of about 3-4 mm. The full engine assembly is 1 cm dia.

NASA is also working on a "steam" turbine, using Xenon as the working fluid, which is heated by various processes: fuel cells, solar concentrators, radiothermal devices, etc.
http://www.nasatech.com/Briefs/Jan03/NPO20933.html

The reason for this is because at the microscale, gasses flow with the viscosity of liquids at macro scales.

Ventions is working on small turbopumps (you could call them an MIT spinoff)

http://ventions.com/technologies.html

[colbourne]

This site might be of interest :-

http://sugarshot.org/

It is almost the opposite in that very low grade propellant (sugar and potassium nitrate)  is used and the goal is to reach 100km rather than orbit.

[RanulfC]

NanoLaunch LLC, a corportion of companies that plan on launching sub-orbital and Orbital Nano-sats using Mig-21 and F-15 aircraft as Air-Breathing First Stages.

See:
http://premierspacesystems.com/files/PSS_Web_Site_White_Paper.doc

and:
http://www.co.siskiyou.ca.us/BOS/DOCS/agenda/2010/Questys/MG3928/AS3941/AI7744/DO7745/2.PDF

NanoLaunch LLC consists of:
"Premier Space Systems:Atmospheric launch, Logistics, Business System"
"Space Propulsion Group: Proprietary Rocket Motor Design, Hybrid LOX/Paraffin Motor Development and production"
"Spath Engineering: Propulsion Test, Manufacturing Mgt."
"Whittinghill Aerospace: Systems Integration"

Suborbital testing to begin by "Summer of 2011"

Randy

[Hop_David]

http://www.astronautix.com/graphics/n/nots.jpg

Went to the above url and got:
"Due to a persistent denial of service attack, astronautix.com has been taken off line."

That's horrible! Apparently another good resource biting the dust

[john smith 19]

See Appendix A.  They assume 301s (vac) for the 8 low expansion engines and 327s (vac) for the 1 high expansion engine.
If you're wondering how you get 9 km/s or more with that sort of performance and a 1500/75 mass ratio, join the club.

Mathematics is not my strong suite but I'll take a stab at it.

 ln (mass ratio) =delta V /Isp *g

 ln 20 = 9000 / (Isp * 9.82 )

Well my calculator gives an Isp of 305.93 say 306secs

Now Jeff Greason had a rule of thumb that *average* Isp is (Isp at SL +2* Isp at vac)/3 so the sea level ones = (238 + 2*301)/3 = 280

But the high expansion nozzle has *no* SL Isp listed, suggesting high altitude start so it's difficult to see how to integrate this Isp into the average low expansion Isp's.

This sounds like a vehicle that has been trajectory optimized so it will deliver the required average Isp, but you cannot prove it without a full trajectory simulation.

One curious feature of the design is the use of a nose cone. Kare mentions in the report that air drag is a big factor in small vehicles and reducing it is important. There are much better fairing shapes than the cone but I'm not sure it they have smaller surface area (Haack series). They would also be quite easy to manufacture at this scale by metal spinning. It's not clear if they knew this and still went with the cone (less surface area, easier analysis?) The better shape would buy a smaller seamless tank although possibly more difficult to wrap with aerogel.

[RanulfC]

But the high expansion nozzle has *no* SL Isp listed, suggesting high altitude start so it's difficult to see how to integrate this Isp into the average low expansion Isp's.

As I understand/understood it all engines are started at launch the High-Expasion engine is just run at low throttle till it ramps up as the Low-Expansion engines ramp down. That could be wrong though. Given they are running the Peroxide through a catalyst prior to injection of the Kerosene the engine and propellants act "hypergolic" for starting anyway so it wouldn't be difficult to air-start the High-Expansion engine.

One comment on the slides suggests doing away with the "High-Expansion" engine entirly and just going with an "aerospike" expansion ramp using the piston-pump exhaust.

This sounds like a vehicle that has been trajectory optimized so it will deliver the required average Isp, but you cannot prove it without a full trajectory simulation.

I'm not sure it's on the slide presentation or not but I recall talking with Kare during a presentation and he noted that they had actually run the entire flight through POST and several other trajectory simulation programs. He commented that he still believes that someone with the money could "easily" prove SSTO by doing a version of the Mockingbird

One curious feature of the design is the use of a nose cone. Kare mentions in the report that air drag is a big factor in small vehicles and reducing it is important. There are much better fairing shapes than the cone but I'm not sure it they have smaller surface area (Haack series). They would also be quite easy to manufacture at this scale by metal spinning. It's not clear if they knew this and still went with the cone (less surface area, easier analysis?) The better shape would buy a smaller seamless tank although possibly more difficult to wrap with aerogel.

At the time I suspect the cone was easier to model and since about 2/3rds of the "nose-cone" is tank for the Peroxide I suspect it was just generally a better shape for the time. They were also looking at using a "peroxide-fountian" effect for cooling the inner surface of the nose cone should the aerogel not prove effective enough as heat protection so that may also have factored into the design. You'll note the suggest "upgraded" designs become "longer" and the nose cone shrinks as the vehicle gets bigger.

Randy

[john smith 19]

He commented that he still believes that someone with the money could "easily" prove SSTO by doing a version of the Mockingbird

I think he's right.

John Whiteheads team at LLNL have continued to refine the pumped propulsion system and Xcor have a version available. BTW People think Aluminum is not up to making combustion chambers but the Bell Aerospace chamber on the Agena stage (IRFNA/Hydrazine so hypergolic and room temperature) was Aluminum and gun drilled the throat section to create a solid wall so *no* joins between inner and outer shells. I think the body of the chamber was done the same way. I'm not sure if they even needed to anodize the interior (relying on thermal barrier coatings with *big* thermal expansion mismatches makes me nervous). Armadillo Aerospace also use Al chambers.

Getting hold of 95% HTP is AFAIK still almost impossible but a small scale continuous flow concentration plant (to limit the amount of >70% HTP in the system) should be a relatively straight forward task for a competent chemical engineer.

GNC has had nearly 20 years more work so decent quality is much more affordable and aerogels are fairly widely available.

Insofar as Blue Origin have said anything much about the Shepperd vehicle it seems to follow roughly the mockingbird design but is sub orbital.

[RobLynn]

Getting hold of 95% HTP is AFAIK still almost impossible but a small scale continuous flow concentration plant (to limit the amount of >70% HTP in the system) should be a relatively straight forward task for a competent chemical engineer.

Can buy a distillation plant suitable to make ultrapure 99% H2O2 for $18k:
http://www.tecaeromex.com/ingles/destilai.htm

[john smith 19]

Can buy a distillation plant suitable to make ultrapure 99% H2O2 for $18k:
http://www.tecaeromex.com/ingles/destilai.htm

I stand corrected. Handy to start high and dilute down.

Note the classic concern about high strength HTP (85%+) is with so little water to take the heat away it will exceed the peak operating temperature of the preferred catalyst (Silver?)

Catalyst pack prep seems to be a bit of a black art, as Armadillo Aerospace found out. IIRC attempts at using catalytic converter type preforms gave poor results despite being highly temperature resistant and designed to act as the base for catalytic coatings.

It's an area where small scale tests using *carefully* documented materials and processes could yield big improvements.

[RanulfC]

The same company cited above actually makes a FIVE-element catalyst pack that they claim is very efficent and long-lasting. Another company General Kinetics Inc. (http://www.gkllc.com/) has done some good work on both injected catalysts and advanced catalyst packs. They also have some really informative papers on H2O2 applications.

Randy
(Edit: Corrected element screen number on catalyst pack: See:
http://www.tecaeromex.com/ingles/cata-i.htm)

[john smith 19]

The same company cited above actually makes a tri-element catalyst pack that they claim is very efficent and long-lasting. Another company General Kinetics Inc. (http://www.gkllc.com/) has done some good work on both injected catalysts and advanced catalyst packs. They also have some really informative papers on H2O2 applications.

Randy

Those chamber pressures (along with the comment on the PDF about their use on an interceptor) suggest they were part of one of the pumped propulsion projects at LLNL.

That would make them quite a good starting point for someone wanting to try a mockingbird design.

Thanks for the link.

[RanulfC]

The same company cited above actually makes a FIVE-element catalyst pack that they claim is very efficent and long-lasting. Another company General Kinetics Inc. (http://www.gkllc.com/) has done some good work on both injected catalysts and advanced catalyst packs. They also have some really informative papers on H2O2 applications.

Those chamber pressures (along with the comment on the PDF about their use on an interceptor) suggest they were part of one of the pumped propulsion projects at LLNL.

That would make them quite a good starting point for someone wanting to try a mockingbird design.

Thanks for the link.

No problem though I can't see or find any evidence that GK folks worked on or with LLNL. They DO have a good number of LLNL reports on reciprocating piston pumps on this page though:
http://www.gkllc.com/lit-misc.htm

The Interceptor rocket motor is the LR-40 made by Reaction Motors International, the same folks who built the rocket motors for the X-1 and X-15 rocket planes.
Here is the LR-40 presentation:
http://www.gkllc.com/lit/gk-authored/AIAA-2001-3838_History_of_RMI_Super_Performance_90_Percent_H2O2-Kerosene_LR-40_RE-pitch.pdf

It's actually an "in-line" turbopump design and doesn't use a piston-pump and with a thrust of 10,200lbs is probably far to much for a "Mockingbird" vehicle

On the other hand some of the smaller motors might work very well...

Randy

[Proponent]

Attached is a paper the title of which mirrors that of the thread.  If it's already been posted in this thread (I haven't been following it), please let me know, and I'll remove this post to avoid pollution of the thread.  [No doubt someone's thinking:  "Yeah?  So why don't you remove all your posts then?" ]

[john smith 19]

The same company cited above actually makes a FIVE-element catalyst pack that they claim is very efficent and long-lasting. Another company General Kinetics Inc. (http://www.gkllc.com/) has done some good work on both injected catalysts and advanced catalyst packs. They also have some really informative papers on H2O2 a

No problem though I can't see or find any evidence that GK folks worked on or with LLNL. They DO have a good number of LLNL reports on reciprocating piston pumps on this page though:
http://www.gkllc.com/lit-misc.htm


Randy

If you click on the pictures of thrustors on their home page they download a 2 page spec sheet. It mentions why they were developed there. 

"Interceptor" is being used in 2 slightly different contexts. The RMI engine was for a plane, the thrustors were for an interceptor missile.

LLNL was very much the birthplace of SDI and the idea of kinetic energy kill vehicles under the Brilliant Pebbles programme.

Thrustors of this thrust level are normally pressure fed but *much* larger due to lower chamber pressure (100-150psi) needing some kind of pump.

Scaling effects worked out at LLNL indicated turbo pumps scale badly below 5000lbs so a reciprocating pump was the obvious alternative, given the defense application. Smaller thrustors -> smaller missile (or more propellant) -> better tracking of maneuvering target.

[RanulfC]

Thanks Proponent, interesting paper!

John Smith 19:
Ok I understand what you're getting at now. I DO suspect that piston-pumps would be better for a smaller vehicle.

I just wish there were more "vendors" for smaller, high thrust rocket engines.

Oh have you looked at the website for Frontier Astronautics?

Randy

[john smith 19]

Thanks Proponent, interesting paper!


John Smith 19:
Ok I understand what you're getting at now. I DO suspect that piston-pumps would be better for a smaller vehicle.

I just wish there were more "vendors" for smaller, high thrust rocket engines.

2 words. "Limited market." Most of these are used on satellites where people are *very* cautious on new hardware, especially commsats where a station keeping thrustor failure will cost *billions* in lost revenue over say a 10 year life. Today they are *all* pressure fed hypergolics with NTO and MMH/UDMH for performance and are not really volume restricted in the way missiles are. No pressure to evolve the technology.

They may also be worried about the bogeyman of HTP decay in the tanks. GK mentioned in one of their papers defending HTP that the Syncomm2 satellite ran 1963-1969 using HTP attitude control with 1960's quality HTP. We *should* be able to do much better today with higher concentrations and lower impurities (both very good for long term storage) *provided* proper materials selections are made.

HTP takes a performance hit at the *same* chamber pressure (but does not cost c $60/lb or need a full body suit to work with it)

I'm not sure if you fully understand that smaller chamber *needs* higher pressure and no one commercially wants that as higher pressure -> heavier tanks, *unless* you go to a pumped solution, which lowers tank pressure (LLNL pumped demonstrator used MMH and Titanium tanks 0.5mm in thickness) but is untried on orbit. For small vehicles it's all about mass per unit of surface area, either the tanks or the chambers.

A comment in one of Whiteheads papers from LLNL was to the effect that the exploration community had relied on these for their main delta V engines on large space probes (to places like Jupiter and Saturn) but that it should start looking to their own thrust chambers (at higher pressures) to give better efficiency in future large missions.
   
In the big aerospace world the nearest thing for this was the Rocketdyne LEAP thruster, using mixed oxides of Nitrogen fuel.  This also ran c500-1300lbs of chamber pressure, which suggests a pump drive, but no details of the drive system were given.

The DoD runs a thing called the "Space Test Program" which test US and foreign space hardware. Potentially this allows new vendors a "qualification" flight so you can demonstrate your hardware works IRL.

Oh have you looked at the website for Frontier Astronautics?

yes. I think I met them some years ago at Space Access, when they are getting started. They've made impressive progress.

[RanulfC]

Interesting little web-site I ran across:
http://home.earthlink.net/~apendragn/atg/qp/

Randy

[strangequark]

HTP takes a performance hit at the *same* chamber pressure (but does not cost c $60/lb or need a full body suit to work with it).

Nope, it just reacts hypergolically with test engineers. Also, since when does hydrazine require a full body suit to work with?

[mmeijeri]

Also, since when does hydrazine require a full body suit to work with?

Do you mean because it is easy to work with once it has been loaded and sealed? Unlike, say cryogens.

[john smith 19]

Nope, it just reacts hypergolically with test engineers. Also, since when does hydrazine require a full body suit to work with?

Probably since the exposure limits for MMH were set at 0.01 ppm and UDMH at 0.5ppm of air. Gasoline at a filling station is around 1000 ppm.

For reference a nerve agent is roughly 60ppm of the mass of a whole person.

Those pictures of techs fussing around some satellite in a clear room are (AFAIK) taken when the sats are *empty*.

This is what NASA calls the gear they use to handle these chemicals and this is what it looks like.

http://www.wolfhazmat.de/astrosuit/nasa_01.htm

It gets the job done but it's scary and expensive.

[mmeijeri]

As I understand it the hypergolics are only dangerous during loading and safing. Once sealed, they're perfectly safe and more convenient to handle than cryogens. That is part of the reason why some people have suggested the use of hypergolics for air-launch. Of course, the cost of the loading and safing may be higher than the savings gained by air-launch.

[strangequark]

Probably since the exposure limits for MMH were set at 0.01 ppm and UDMH at 0.5ppm of air. Gasoline at a filling station is around 1000 ppm.

For reference a nerve agent is roughly 60ppm of the mass of a whole person.

Those pictures of techs fussing around some satellite in a clear room are (AFAIK) taken when the sats are *empty*.

This is what NASA calls the gear they use to handle these chemicals and this is what it looks like.

http://www.wolfhazmat.de/astrosuit/nasa_01.htm

It gets the job done but it's scary and expensive.

I know what a SCAPE suit looks like John . Granted, I may have been a little disingenuous, but my point is that hydrazine in small quantities doesn't require anything more than pvc gloves, eye protection, and a fume hood. People see what HTP requires in smaller quantities (or what was required in the 60s...) and generalize upward, which is not necessarily valid.

I guarantee you that if NASA used HTP in large quantities, they would require PPE just about as bad. Peroxide isn't toxic, but it is damned corrosive and reactive. This forum seems to have a large attraction to HTP, and I suspect it's because it's the devil they don't know. I'm not saying it's a bad propellant, but it's not all peaches and cream either (the Isp on that combo is terrible anyway).

Martjin is correct too, in that hypergols are pretty reasonable to work with if you have everything sealed and are doing automated transfers. There's even a system in work for CubeSats using monoprop hydrazine. The trick is that the system is delivered preloaded, and fully hermetically sealed.

[mmeijeri]

Well, you know me quark, now I have to rush to the defence of peroxide!  Seriously though, is Isp that bad? I thought it was only slightly worse than MMH/NTO. What combination were you thinking of?

[strangequark]

Well, you know me quark, now I have to rush to the defence of peroxide!  Seriously though, is Isp that bad? I thought it was only slightly worse than MMH/NTO. What combination were you thinking of?

Haha, sorry I was trying to make a joke that may have fallen flat on the internet. I meant the Isp for peaches and cream would be terrible. HC with Peroxide is decent for Isp.

Honestly, what it comes down to is that I don't think HTP really gains you all that much, compared to more developed systems. It's not that it's bad, it's that the trade isn't positive enough.

[john smith 19]

Honestly, what it comes down to is that I don't think HTP really gains you all that much, compared to more developed systems. It's not that it's bad, it's that the trade isn't positive enough.

The primary drivers for the switch in the 1960s were performance and long term (*decades*) storage. Part of why France achieved orbital launch years before the UK. Everything else was irrelevant primarily to the military planners who wanted these systems.

At the same chamber pressure and nozzle size conditions the amines win. This is undisputed. The storage position is more debatable.

HTP does degrade back to water an O2 with a pressure increase. Permanently sealed tanks without pressure relief would be highly dangerous. However a number of materials are known to be *much* less affected by the process and much less prone to catalyzing it. Tin served as an inhibitor and and it is possible to "passivate" surfaces by flushing systems with HTP before filling, as was done with Inhibited Red Fuming Nitric Acid or IRFNA (another popular oxidizer before the hydrazines).

General Kinetics noted the Syncomm 2 satellite used HTP form 1963-1969 without problem, as did the X15.

BTW NTO is pretty corrosive and will happily attack soft seal materials, and your lungs. 

If you're military or LLNL then Hydrazine is fine. It's hazards are known and factored into your costs.

Outside these groups any one looking to experiment will look at the $60/lb costs of Hydrazines, the WMD level of toxicity and consider the performance is (usually) not worth the risk.

In the context of "Micro rocketry" implying smaller developers might like to make attempts the costs and danger (I would suggest) far outweigh the performance benefits.

Note John Whiteheads "Mockingbird" SSTO was an HTP design and they have mentioned that their design work for the JPL funded micro pumped systems were developed in high pressure Helium and HTP because it let them do more runs in a *normal* laboratory environment rather than suiting (it's not the pump, it's the gas generator that's the issue).

For an experimental "micro rocket" I'd say the first objective is to build something a small team can handle and find out it's payload *after* it's built.

Outside military systems and space probes storage time is not a very relevant parameter to consider. If you did the *obvious* answer is to go to solids, which the USAF starting with Minuteman and the USN with Polaris.

With the exception of Orbital Science Corp I don't see anyone rushing to follow.

[mmeijeri]

Honestly, what it comes down to is that I don't think HTP really gains you all that much, compared to more developed systems. It's not that it's bad, it's that the trade isn't positive enough.

Well, that depends on the application as John suggests. For use in an RLV HTP seems more practical than hydrazine, since you would have to load and safe the vehicle many times. HTP (or NOFBX once proven safe and reliable) would probably also be preferable for a crew taxi, except in the short run, which is why I assume Boeing and SpaceX went for hypergolics.

But I think you are right for spacecraft that don't return to Earth. I see no rush to replace traditional hypergolics with something less toxic or corrosive in that field. To the degree there is an urge to use HAN / ADN it seems to be driven more by a desire for higher Isp than monopropellant hydrazine than by a desire for better handling characteristics.

[Ed LeBouthillier]

Randy said:

> Interesting little web-site I ran across:
>
> http://home.earthlink.net/~apendragn/atg/qp/

I've also started a blog to develop ideas related to small launch vehicles. You can find it here:

http://orbitalaspirations.blogspot.com/

Cheers,
Ed L

[strangequark]

(or NOFBX once proven safe and reliable)

Don't hold your breath. From the information available on that propellant, I suspect a bipropellant would be more reliable, safer, and less complex. If someone from Firestar would care to correct me, I am willing to learn, but that is my impression thus far.

But I think you are right for spacecraft that don't return to Earth. I see no rush to replace traditional hypergolics with something less toxic or corrosive in that field. To the degree there is an urge to use HAN / ADN it seems to be driven more by a desire for higher Isp than monopropellant hydrazine than by a desire for better handling characteristics.

That is correct. The easy handling characteristics of HAN are lagniappe, and it's the Isp that is really driving that. For this topic, that is almost a negative, because the reaction temperatures drive you to exotic materials.

As far as pulling this back to micro-rocketry, my propellant choice, if I were head designer, would be biprop N2O with Ethanol. They're widely available, pretty safe to handle (if you respect the N2O), and have respectable performance. It's not hypergolic or catalytic, but if you weren't picky about ignition time, you could use an acoustic igniter with the nitrous and not have to use pyro. Alternatively, thermal decomp of nitrous isn't too hard, and the required heater is a pretty inexpensive, off the shelf part.

[mmeijeri]

As far as pulling this back to micro-rocketry, my propellant choice, if I were head designer, would be biprop N2O with Ethanol.

Any specific reason to prefer N2O over HTP or ethanol over IPA?

[A_M_Swallow]

As far as pulling this back to micro-rocketry, my propellant choice, if I were head designer, would be biprop N2O with Ethanol.

Any specific reason to prefer N2O over HTP or ethanol over IPA?

HTP can be made from water so it can be made by an ISRU process on the Moon.

[tnphysics]

As far as pulling this back to micro-rocketry, my propellant choice, if I were head designer, would be biprop N2O with Ethanol.

Any specific reason to prefer N2O over HTP or ethanol over IPA?

As for N2O, it

-is non toxic (though an anesthetic in high levels, where high means percent)
-is not decomposed by most metals AFAIK

[Enthalpy]

There I suggest a small launcher that puts 300kg on Leo:
http://saposjoint.net/Forum/viewtopic.php?f=66&t=2527&start=20#p35066

By far not the smallest possible launcher, but small in the sense that it's built around the smallest engine that enables reasonably a single stage to orbit (SSTO) - an RD-0146, heir of the RL-10 but with a roll actuator.

Much smaller is possible of course. As you probably don't want to develop optimized pumped engines just for the micro-launcher, solids or pressure-fed oxygen+fuel would be the choice.

Marc Schaefer, aka Enthalpy

[Enthalpy]

The physical limit is the atmospheric drag, so I've made a quick-and-dirty spreadsheet to evaluate it.

The spreadsheet finds properly that air drag is tiny at normal-sized launchers, like 50m/s performance wasted by Vega, which is compact but accelerates strongly. That explains why Energiya or the SSTL were as streamlined as a cathedral.

The limit is around 100kg lift-off mass, with D=0.3m and 10+8m/s² acceleration and a good Cx=0.25. This wastes significant 270m/s.

Or around 10kg lift-off mass, with D=0,14m and 10+6m/s² acceleration and a good Cx=0.25. This wastes 500m/s, which hurts. Leaves you maybe 200g in orbit, depending on the propellants.

BUT if you start from an old MiG-21 or MiG-25 or F-4 Phantom, at 18,000m height, air is 10 times thinner, so your launcher can be 10 time smaller in each dimension, or 0.01kg. The aircraft even provides the 400m/s lost to air drag. And the launcher can have small wings, allowing the engine to push less, hence expand its gas more efficiently. Reasonable for small launchers.

Then, an other limit is the mass of the tanks or the casing, as thinner walls get impractical. At D=0.2m and 50b, good steel would be around 0.4mm thick. Good aluminium, at 2mm, is less efficient but can go a bit farther.

At solids, you need a minimum insulator thickness inside the casing, like 10mm.

In other words: the limit is set by the sensors, the transmissions, the ability to control a craft with fast reactions...

[Ed LeBouthillier]

The physical limit is the atmospheric drag, so I've made a quick-and-dirty spreadsheet to evaluate it.

The spreadsheet finds properly that air drag is tiny at normal-sized launchers, like 50m/s performance wasted by Vega, which is compact but accelerates strongly. That explains why Energiya or the SSTL were as streamlined as a cathedral.

Would you mind discussing the techniques you used for determining the aerodynamic losses? I've been looking at a good way to visualize the relationship between size and aerodynamic losses. I'm curious about how you went about discovering it.

Thanks,
Ed L

[Rick M]

This site might be of interest :-

http://sugarshot.org/

It is almost the opposite in that very low grade propellant (sugar and potassium nitrate)  is used and the goal is to reach 100km rather than orbit.

We had two motor CATOs on our 120,000' DoubleSShot static motor tests but believe we know the cause after post anomaly testing and will soon be doing our next test. Ed did some calculations that showed we could get a sugar powered rocket to orbit; a 'SSTO' meaning Sugar Shot to Orbit with 5 stages and a tiny payload ( :

[meekGee]

Nice looking solid motor core.

I love the guy running away from the rig (at T-5) like Wile-E-Coyote after lighting the fuse on an ACME rocket.

[RanulfC]

Necro-Threading but thought I'd update that a similar discussion is now here:
http://forum.nasaspaceflight.com/index.php?topic=31135.0;all

Randy

[cpooley]

Microlaunchers is based on starting with a roughly 120 kg GLOW and a 100-200 gram spacecraft to escape, to photograph a NEO and play back stored images over rest of 1st year via laser diode.

Most of the losses relative to small size are with the 1st stage.  The 3 or 4 stage rocket can be partitioned to have the 1st stage just clear the atmosphere.

In vacuum, even very small rockets are nearly as good as "conventional" size.

Starting with escape makes sense because there's no future in crowding LEO space, and LEO is not exploration anymore.

[QuantumG]

Welcome back Charles. Perhaps you could work with Ed LeBouthillier and put together one of his awesome diagrams to explain the concept.