What does everyone think about throwing together a quick and dirty static hybrid rocket to get a feel for the technology? | Hybrid engine discussion | Forum

Post edited 8:48 pm – March 23, 2010 by brmj

I've been thinking it shouldn't be too hard to build a rudimentary hybrid rocket just to get a feel for the technology, play with nozzle designs and fuel/oxidizer chemistrys and have some physical hardware and impressive youtube videos to show off in the very near term. This ought to be something that can be built with off the shelf or salvaged parts, with the exception of the nozzle and perhaps a few other parts that will have to be machined. We can run it off of gaseous oxygen at first for availability reasons, build the walls out of metal pipe of the appropriate diameter, and generally design it to be dirt cheap and quick to build at the expense of efficiency. This ought to be something we could do pretty quickly. I hesitate to call it a weekend project, but to someone with the right tools and experience, somehting much like this ought to be very quick and easy.<
Something a bit like the following is bassicaly what I had in mind.

What does everyone think?

I agree with this. This is something that could be accomplished in a relatively short period of time. To proceed with this, I think we should determine a set of objectives and constraints. What exactly are we trying to accomplish. What altitude should we strive for? Will we take any sort of pictures at the apogee?

I am willing to spend time helping to design this. It will give good experience towards OHKLA.

Post edited 5:44 am – March 26, 2010 by rpulkrabek

Luke Maurits said:

I am confused.  Isn't the call here for a static rocket engine?  There would be no altitude or apogee with this.  If we were to do this, we should aim for particular thrusts and burn times.

Don't be confused. It was me that was confused. Somehow I forgot to read the word "static". The drawing now makes more sense that the oxidizer isn't connected the way I thought it would be. Whoops. Sorry about that. 

In any case, I still like this idea. What I am looking forward to the most is being able to confirm the theoretical thrust values determined through FEA with actual measured values. I say let's do this!

The motor my team and i designed using N2O and HTPB resulted in about 350N check out some of the videos http://my.fit.edu/pantherii (go to galleries) We used speaker wire cut so that it would arc  in a GOX environment which would start the burning, which would then burn through a ziptie, which released the N2O into the Chamber. This method seemed the best for ignition at this scale.

This thread needs to be revived.  Testing fuel grain porting and design can be done, get a 6 foot length of a scalable diameter, section it, and ship it to those parties with the means to drill out the blanks into the desired patterns.  As for the nozzle, we need some drawings so stock material and a machinist can be found to get prototype nozzles made up.  As for the combustion chamber, I think we need to decide on a set diameter for a prototype rocket and work on coming up with materials so that we can start building.  I can put the test rig together, 2500 lb load cells are on ebay for $25, I've got spare stock steel, land, and static testing experience, hell I can drill out the fuel grains. 

Set, go.

rpulkrabek said:

Yes, this needs to be done. Let's make sure to have an IRC meeting with the goal of making a decision on the diameters. From this we can determine a nozzle geometry to test and have some CAD drawings (2D that we can hand over to be manufactured).
 

The sooner we build, the better. These tests need to be done, before we can have any sort of final design. At the moment there are too many unknowns with the rocket to understand how it will actually fly. 

How does every one feel with the test rocket having an internal diameter of 2.5 in.?

Luke

HTPB is on the ITAR list and can’t be exported from the USA without completion of an export licence
application to the US
state department. This would be near impossible to obtain for a project such as
CSTART, but there are alternatives in hydroxyl functional polymers that can be
processed at room temperature. Such as Polyurethane elastomers, or
Polypropylene Glycol (PPG) which offers better performance than HTPB. HTPB is
far from an optimum fuel for hybrids, it was only chosen by the industry due the
massive quantities on hand for manufacture of solid propellants, as well as the
industry know how for processing.

I have had nearly 30 years exposure to hybrid propulsion now, much of this
coming through an commercial agreement with AMROC in the late 80s. And I can
tell you now that the Aerospace communities’ dependence on LOX is one of major
barriers to developing the technology, with a massive amount of the issues
being caused by irregular mixing of oxidiser & fuel. A large amount of
energy has to go into vaporising the LOX, which has required either different
forms of pre-combustion chambers or injection of hypergolic fluids (TEA). This
has plagued the technology with inefficient & unstable combustion, and when
combined with solid polymer fuel grains of low regression rate; you get large
complex and poorly loaded combustion chambers. These are precisely the reasons
that ventures by the likes of Lockheed Martin etc fail to produce large hybrids
that reduce the cost or deliver on the real promise of the technology, which is
simplicity, safety & economy.

The Swedish & the French achieved more with Hybrid technology in the 70s
and the 80s using combinations of Nitric Acid & hypergolic fuels, producing
large sounding rockets capable of carrying payloads to over 100km. Unlike many
"experts" I would not rule out Nitric Acid or Hydrogen Peroxide so
quickly, they both have very important upsides and are not as hazardous as
proponents of LOX & NOX would have you believe. They are also not as
difficult to procure and can be more economical than NOX, with Nitric Acid at
98% being a commodity chemical & Peroxide being producible by concentrating
commercial stock (50% up to +85% by simple evaporation). I have worked with
Nitric Acid a few times historically (White Fuming Nitric Acid or WFNA) &
had no problems; it is no less safe than dealing with a highly compressed
oxygen gas such as NOX. You just need to read an MSDS, understand its
properties, develop procedures & STICK TO THEM!. If people cant do this
then they probably have no business trying to build rockets, as safety needs to
be paramount at all times. I hear a lot of “experts” spout the Acid &
Peroxide are highly dangerous & toxic line habitually, without actually
putting it into fair context. Its sobering to remember that both of these
chemicals are handled in the mega-tonnes per year by fairly non-skilled
labourers, Peroxide at 70% is used in the paper pulp industry for instance. And
if spilled can be diluted by demineralised water & will convert to steam
& oxygen which is eco-friendly!. WFNA will breakdown into nitrates
(fertiliser) & Nitric Oxides (which are more toxic but a common by product
of most combustion process), but it can be cleaned up absorbent media fairly
easily.  The poor reputation of Peroxide
steams from the early German experience of the 1940’s when poor handling
practice (ie pouring it out of buckets), combined with the impure manufactured
product lead to catastrophes. Most of this has taken on urban legend type status
in the propulsion community, but as the likes of Beal Aerospace & to a lesser
extent Armadillo showed it is mostly unfounded.   

WFNA is double the density of NOX and 85% Peroxide is close to the same
(1540kg/m3 & 1380 kg/m3 respectfully), plus they are room temperature liquids.
Peroxide can be stored in plastics such as HDPE, U-PVC & Polycarbonate,
which opens up the possibility to make very lightweight plastic lined composite
propellant tanks. Where as LOX & NOX require alloy lined tanks or exotic
new composites, which are outside the capabilities of the experimental
community.

Ignition & mixing becomes far more efficient as energy requirements to
vaporise room temperature oxidiser is far less than cryogenic ones, ignition
can be achieved by using catalyst (Peroxide) or pellets of hypergolic
substances (or mixtures in the fuel grain) for Nitric Acid. Peroxide has the
added benefit of being decomposable to super hot steam of oxygen and water,
which ideal for ignition & mixing with the solid fuel grain.

So in summary I wouldn’t just dismiss these oxidisers from your list of
candidates so quickly, but instead do a more thorough investigation of them
based on your long term project goals. Rather than assessment based on the ease
of obtaining a propellant as the main focus.  

On NOX..

NOX seems like a good oxidiser for small hybrids, however the current mantra
that it is the safest oxidiser is proving not to be true. The big issue is the
NOX vapour which is highly unstable, and can reach and sustain a detonation
wave with a low activation energy. In fact some hybrids have exploded purely
from heat wash back from the injector face; it can also detonate from adiabatic
compression (i.e hammering or bubble compression in lines) as the guys at
Scaled Composites found out. Proper design can eliminate most of this risk
& I would strongly advise pressurising the tank ullage on your hybrid motor
with helium gas, which will help inert the vapour left in your tank after
liquid depletion. And make sure you get hold of and understand the safe
practice for handling oxygen systems before playing with NOX, I have seen too
many people out at launches who don’t respect NOX to the level they should.
Exposure to any containments, dust, grease fabric etc can lead to auto-ignition
with limited energy input (i.e static discharge), & be aware NOX is a
strong solvent to many plastics. If you are using it in a hybrid with a plastic
fuel & your ignition fails leading to venting of NOX through your chamber,
be prepared to toss your fuel grain & replace it.

Guys I am about to fire the motor for my project over the next few months,
it is a NOX/ Sorbitol hybrid with a thrust of over 2000N for 7 seconds &
would be considered an N class HPR designation. I went the Sorbitol direction
for a few reasons, mainly because of density impulse, ease of fuel casting
& availability of Sorbitol at the time (as I have 25kg on hand). I'm just
redesigning the motor head & injector enclosure to incorporate a new
throttle valve assembly I have come up with, so once I have completed this I
will upload some pictures to the blog page & point you there. Its Important
to note though that we (ASLI) are going to move away from NOX to Peroxide in
the next generation propulsion and onward, as its our belief that NOX will not
scale in a useful manner. In fact there is some evidence that large NOX hybrids
are going to be dangerous as the volume of NOX vapour in the pre-combustion
chamber of the motor size grows, so to will the risk of large scale detonation
waves propagating and thus motor case failures. I for one will not be signing
up for a ride on Space Ship 2 just yet guys..

I am happy to commit some expertise to your program in a kind of consultancy
role, but I’m pretty busy working on my own program & there is some stuff I
am doing that I won’t share as I have commercial aspirations.

But I can certainly help the group to correctly design and test a useful
propulsion system..

Cheers – Jamie

Hi Paul,

Hi Paul,

 Okay I had a look over the CAD diagrams, and the concepts contained in them
are moving in the correct direction. However there are issues with the sealing
designs, as well as lack of thermal insulation of the combustion chamber. Face
sealing with gaskets is normally not applicable to rocket motors, as any error
in tolerance will allow super hot gas leaks to burn holes through your casings.
A good example is when myth buster built a rocket to bust the myth about the
Mulsim rocket man, the solid fuel motor they built used face seals. When ignited
it formed a leak and burnt through in seconds, leading to a sideways vent
forming off axis thrust which sent the rocket spinning out of control.  What I would suggest for this project is that
you develop a disposable combustion chamber using ablatives & composites; I
can give you pointers on how to do this. What this will do is eliminate a lot
of the complexity involved in trying to design a reusable combustion chamber
based on metals, which requires a lot of work to make adequate seals & is
much heavier. Also the brace to hold the fuel grain in is not required, the
post combustion chamber will perform that task for you.

For your nozzle the best way to go for your flight weight motor is to
produce a moulded silca (or glass) phenonlic resin one, with perhaps a graphite throat insert.
For a test stand motor you could go with a steel or Inconel housing with a graphite throat
insert, or a solid graphite nozzle braced in an alloy holder (which is what I
have done for the reusable hybrid for ASLI). The next generation motor will be
using moulded silca/ phenolic pre/post combustion chamber & nozzle as well
as a disposable composite motor case. A suitable phenonlic resin for making
said parts is called CELLOBOND & is a product of Momentive see here: http://ww2.momentive.com/Produ…..px?id=5634

It is two part liquid phenonlic resin which you can fill with fibers (or
microspheres) & additives (carbon) to cast into ablative parts suitable for
rocket motors; I have made quite few nozzles from this material for solid
propellant motors over the years.

You should be able to find a distributor in the US pretty easily..

For injectors – DONT USE BRASS it is a catalyst when hot to Nitrous Oxide
and can produce a detonation once it warms up under firing. Best to use either
copper or stainless steel; alloy can be used if you provide some film cooling
on the injector face (i.e a few injector holes parallel to the injector face).
NOX injectors are difficult to accurately model BTW the main issue is the
saturated liquid/ vapour behaviour properties of the substance, be prepared for
sometimes wild variations in flow rate & motor performance. When design
allow a large pressure drop across your injector as NOX is notorious for
chugging (fluctuations of back pressure through the injector), if your chamber
pressure is too close to your tank pressure. I use over 250psi generally to
ensure good margin of safety for variable NOX pressures, which is a problem due
to its strong temperature dependence.  

Some questions

What level have you gone too to size this rocket can I ask? What software
are you using for aerodynamic & trajectory simulation?. Have you developed
any tools to size your propulsion & look at critical issues such as fuel
regression rate?.

I am happy to share a spreadsheet I have developed for sizing hybrid motors,
it performs all the first level analysis of a design based on your desired
thrust/ burn time & diameter inputs. It will output critical figures for
your proposed motor such as: Fuel grain regression & sizing, nozzle sizing,
injector sizing & NOX tank sizing dependant loading temperature. It can easily
be adapted to any fuel combination if you have the regression rate & thermo-chemical
information

 We use this at ASLI for our first
level performance trade studies, before committing to more thorough analysis,
which includes a accurate Matlab package we developed for modelling the Nitrous
Oxide self-pressure feed model.

You should also download the program RPA v1.2, it will allow to you to build
thermo-chemical models of different propellant combinations.

Hi There,

You can check out http://www.asli2007.blogspot.com or http://www.asli2010.blogspot.com

Propulsion work is not public yet, but we will be updating progress there soon.