TRIPOLI PREFECTURE # 138 › tripoligerlach › downloads › 2013-03-01.pdf · 2013-01-02 · Tripoli Gerlach was set up as a Research Rocketry Prefecture forTripoli members who

www.tripoligerlach.org

JANUARY 2013 Vol 03 No 01

TRIPOLI PREFECTURE # 138

TRIPOLI GERLACH NEWSJANUARY 2013

RENEWALSJanuary 1st is the beginning of our fiscal year and withthat all Member’s Dues are - due! All Tripoli GerlachMembers were mailed a renewal form. If you haven’tdone so yet please fill it out and send it to our TreasurerDave Rose along with your dollars.

MAGAZINE STAFFEDITOR & LAYOUT:

CONTRIBUTING STAFF:

GUEST EDITOR:

OFFICE:

Tom Blazanin

Chris PearsonDave Cooper

Randy Sobczak

Tripoli Gerlach748 Galloway DriveValencia, PA 16059

Tripoli Gerlach does NOT promote nor certify anyactivities presented here as safe nor appropriate toall readers of this Publication. Information is foreducational purposes only. Tripoli Gerlach, itsmembers & officers, the authors of articlespresented and the Tripoli RocketryAssociation, Inc.are not responsible for reader’s activities, conductor participation in the use and pursuit of any contentpresented.

DISCLAIMERON THE COVER TheTripoli Gerlach MascotRocket stands ready foranother great year of rocketactivities.

This year will see it at threeevents at Black Rock:LDRS, HAMSTER DANCE& BALLS!


JANUARY 2013 Vol 03 No 01

TRIPOLI PREFECTURE # 138

JANUARY 2013 Vol. 03 No. 01PUBLISHED EXCLUSIVELY FOR

THE MEMBERS OF TRIPOLI GERLACHand anyone else interested

All Content Copyright ©2013 by TRIPOLI GERLACH

TRIPOLI

Tripoli Gerlach was founded as a National Prefectureunder the Tripoli Rocketry Association, Inc. Devoted toResearch Rocketry and the Black Rock Desert area ofNevada, we welcome all qualified Tripoli Membershaving a Level 2 certification or higher.

Our Executive Officers are;

Tom Blazanin (003)Prefect

Valencia, PA

Dave Rose (7126)Treasurer

N. Huntingdon, PA

Deb Koloms (9021)Secretary

Watertown, NY

If you have anything to contribute in the way ofinformation, articles, photos or whatever, please sendthem to Tripoli Gerlach Headquarters. Visit our WebSiteon-line at;

[email protected]

[email protected]

[email protected]


PAGE 2

As of January 1st we have not received renewalsfrom the following members:

Waysie AtkinsPaul HolmesBruce KellyDeb Koloms

Robin MeredithChuck RogersMark CanepaDoug Gerrard

Micheal LeenellettBob Schoner

Feel free to use the form on Page 4 to renew. See youin the Summer!

Tripoli Gerlach was set up as a Research RocketryPrefecture for Tripoli members who prefer to launch atBlack Rock in Nevada. It is based in Gerlach, Nevadaand has no actual members residing there. It was neverdesigned, nor intended to be an elite group, but did wantits members to be qualified to conduct ResearchRocketry activities as defined by the Tripoli RocketryAssociation, Inc. To this extent we required personsinterested in joining the Prefecture to hold at least aTripoli Level 2 Certification.

Apparently there are some other parties wishing to formstrictly Research Prefectures and have approachedTripoli for Prefecture status as a Research Prefecture.They too want to require a Level 2 Certification for itsmembers and have setTripoli Gerlach as the example.

It becomes obviously clear that if one Prefecture can doit so can many. This in turn would lead to Prefectures allover the place to require, or I should say, limit theirmembership to a specific group; leaving many Nationalmembers out of the picture in the hobby.

Since Tripoli Gerlach was the first to have a Level 2membership requirement based on its intention andlocation, no one questioned us and the world was atpeace. However, our uniqueness seems to have grabbedthe attention of others who, apparently, wish toduplicate our status. As a Prefecture in 100% support ofthe National Tripoli Organization we need to addressthis.

Our Level 2 requirement is written in our By-Laws andto make a change to the By Laws requires a BusinessMeeting as stated:

Since we only have one Business Meeting a year tomake a By Laws change would require time. We havean opportunity to hold a Business Meeting in July atLDRS 32 where we can act upon this. BUT we need to

VII CHANGES TO BY-LAWSA. The Tripoli Gerlach By-Laws may be changed oraltered, only at a business meeting announcing suchchanges and having a quorum present. Any and allitems within Tripoli Gerlach By-Laws may be changed,except Chapter VIII, DISOLUTION.B. To insure a stable platform any and all changes to theBy-Laws approved by a majority of a Quorum shall notbe able to be re-acted upon for a period of one year.

PAGE 3TRIPOLI GERLACH NEWS JANUARY 2013

do something to assist the Tripoli BoD from havingother parties requesting this requirement.

As Prefect I proposed we change our By LawsMembership Requirement:

To read:

On December 3rd a special vote was taken of allmembers of Tripoli Gerlach to vote YES or NO on thischange. This posting announces our results

Response has been very enthusiastic with just about allmembers voting YES. Actually no one voting votedNO. The general consensus is that this, in no way,would affect the group’s status as a Research Prefectureand would open up the group to new people seriouslyinterested in advancing into Research Rocketry.

Six responses contained very good positive input for usto address at the next Member’s Meeting. One memberabstained, though not because they were against this.

With this our next step is to change any public postingof the old Level 2 requirement and begin accepting anyNational Tripoli Member in good standing into ourPrefecture. The posted public By Laws will reflect thischange and we will operate with it until our nextMember’s Meeting, at which time we will ratify thechange and it will become our official document onceagain.

Four members did not vote.

No members voting voted against this change. Nonegative comments were voiced. All members votingare thanked for their quick and appropriate votes.

II MEMBERSHIP

II MEMBERSHIP

A. Membership in Tripoli Gerlach shall be open to anyand all serious minded rocket enthusiasts, no matterwhere his or her place of residence, holding a currentNational Tripoli Membership and certified within theTripoli Rocketry Association with a Level 2 or higher.

A. Membership in Tripoli Gerlach shall be open to anyand all serious minded rocket enthusiasts, no matterwhere his or her place of residence, holding a currentNational Tripoli Membership .

CALL FOR BY LAWS CHANGE

PAGE 4 TRIPOLI GERLACH NEWS

MEMBERSHIPMembership in Tripoli Gerlach runs from January toJanuary and is open to any and all National TripoliMembers in good standingl.You can join anytime of theyear however, your Membership renews every January.

Of particular interest should be High Power andResearch Rocketry conducted at Black Rock, Nevada.Potential Members should commit to attending at leastone launch at Black Rock. We encourage and supportthe BALLS launch conducted byAPHRA.

We also hold our annual Members Meeting Fridayevening of the BALLS Launch which gives us the bestopportunity to have the most members present. This isthe time we hold our Elections, and conduct ourPrefecture business for the year. The meeting includes aFree Spaghetti Dinner, which has proven very popular -it’s FREE.

In addition to a serious interest in Rocketry activities apotential member should have an interest in the BlackRock Desert area of Nevada. Many membersparticipate in Trekking, our name for getting out andexploring the desert and surrounding areas toappreciate the natural environment.

It is a known fact once you see Black Rock you’ll returnyear after year - or at least make an effort to do so. We dorealize that because of personal reasons it might be hardto attend every year. All we ask of our members is thatthey TRY.

If you are interested in joining us please use theMembershipApplication provided here.

[email protected]

MEMBERSHIP APPLICATIONNAME ______________________________________________________ TRA# __________

ADDRESS _________________________________________________ CERT LEVEL _____

CITY __________________________________ STATE _________ ZIP _________________

E-MAIL _______________________________@____________________________________

PHONE (________) __________-____________________ CAN WE PUBLISH (YES) (NO)

HOME PREFECTURE _________________________________________________________

PERSONAL WEBSITE _________________________________________________________

YEARLY DUESARE $20.00

Make Check for $20 Payable to

13385 Lincoln WayN. Huntingdon, PA 15642

DAVE ROSE As a member I will abide by all rules set forth by the Prefecture as well as those set forth bythe National Organization. I pledge to pursue a commitment to the Prefecture’s designatedLaunches &Activities and support the Prefecture to the best of my ability.

Please fill out this formcompletely. It is necessary tosupply us with your TRAnumber and certificationlevel. Membership runs fromJanuary to December.Renewal will be January 1st,

If at all possible please E-Mailyour photo to be posted to:

[email protected]

SIGNATURE

DATEwww.tripoligerlach.org

Serving Experimental & Professional Rocketry Since 1995

www.rocketmotorparts.com

TRIPOLI

VISIT OUR WEBSITE

2113 W. 850 N. St. • Cedar City, UT 84721 • 435-865-7100

Rocket Motor PartsPropellant Chemicals

Technical TrainingInformational Resources

RCS Rocket Motor Components was founded by Gary Rosenfield in 1995 to offer AeroTechRMS-compatible parts for custom composite propellant rocket motor fabrication. These partsincluded molded phenolic nozzles, propellant casting tubing, phenolic and fiberglass motorcasing material, phenolic and paper liner tubing, o-rings, casting plugs, forward & aftinsulator washers, and delay/smoke charge insulator tubes

TRIPOLI GERLACH NEWSJANUARY 2013PAGE 4

Last issue we gave you the story about the Paiutesplacing various areas on their reservation in Nixon OffLimits to all non tribal members. This means a lot ofextremely interesting “natural wonders” are no longeravailable to visitors to the Black Rock Desert area.

Also in past issues we gave stories of trekking aboutdiscovering a lot of interesting sites and neat things.One of the last areas closed in 2012 was the Great StoneMother near Pyramid Rock. The above photo is a rareshot as seen from the water. The Great Stone Mother ison the right and Pyramid Rock, for which Pyramid Lakewas named, is on the left.

This area is an almost spiritual place. Of course thePaiutes will tell you everywhere is spiritual and they aremostly right. The Pyramid Rock area abounds withTuffa formations and one can spend several days thereexploring - but be careful of the giant meat eatingspiders!

The water is crystal clear with waves that look andsound as if you were at a seaside beach. You can not getto the Pyramid Rock without getting wet as it is anactual island and it is said the giant meat eating spidersare much bigger there.

A place most may never get to see is the Needles areashown in the two photos below. It is a large area of Tuffaformations at the north end of Pyramid Lake. It can beseen quite nicely from the road, in a panoramic view ofnearly the entire area. But things you won’t get to see ishow bizarre this place is up close.

As with most areas; photos do not do things justice.Time was when anyone could visit the area and marvelat the scenery.

TRIPOLI GERLACH NEWS PAGE 5JANUARY 2013

An area offen viewed by people streaming their wayNorth to Gerlach is called the BeeHives. It is also onthe list of banned places. These unique Tuffaformations are right out in the open for all to see butbeing on Reservation Land one can not venture anycloser than their view from the road.

Meanwhile back at the Great Stone Mother site theblack volcanic beach sand is rippled with permanent“foam” A closer look reveals miniature sea shells. Thecenter shell in the photo is big at less than ¼” across.

Two sites banned have petriegyphs that are more thanjust drawings on the rocks. These are actual carvings inthe rocks. Most petriegyphs are painted with somehaving an attempt at stone work. These are actualcarvings set for the ages. Naturally the Paiutes wishthese preserved as sacred.

The most unique feature at this area is on the Pyramiditself and can not be seen from the shore. It is on the westside of the rock facing the town of Sutcliff across thelake. It is a Geyser spouting hot water out about threefeet from the water surface.

The last thing we show is the actual existence of thegiant meat eating spiders of Pyramid Rock andsurrounding area. These suckers are big and mean.When approached, and you approach by accident, theywill actually charge you and hold their ground - andeven lurch toward you in threat. These photos of thespiders, like all the rest, do not do justice. Maybesomeday you’ll be able to see for yourself.

PAGE 6 TRIPOLI GERLACH NEWSJANUARY 2013


Propellants function to impart motion to an objectthrough the conversion of potential energy into useful orkinetic energy. Two ingredients, a fuel and an oxidizer,neither of which will burn satisfactorily without thepresence of the other, are necessary in a propellantsystem.

Two main classes of propellants are recognized on thebasis of physical character: liquid propellants and solidpropellants. Most solid propellants belong to one or theother of two types. Homogenous propellants containboth the oxidizer and fuel in the same molecule and mayalso be referred to as monopropellants. These systems,consisting mostly of nitrocellulose and nitroglycerine ina colloidal mixture, are called double-basedpropellants.

The second type is the heterogenous or compositepropellant, where the oxidizer is a finely groundinorganic salt and the fuel is plastic in nature, bindingthe propellant grain structure together. Black powder,the oldest of propellants, falls into this category since ituses potassium or sodium nitrate as the oxidizer andsulfur as both binder, and with charcoal, as a fuel.

Modern composite propellants first emerged in the late1940's. These incorporated various thermoplastic andthermosetting resins or elastomers with a variety ofnitrates or perchlorates as oxidizers. Perhaps the mostpopular of the composite propellant systems in currentuse consists of ammonium perchlorate as the oxidizerand usually a copolymer or terpolymer of butadienewith other monomers such as acrylic acid oracrylonitrile as the binder.

This document will examine the design andconstruction of composite propellant rocket motorsusing hydroxyl terminated poly butadiene (HTPB) andammonium perchlorate (AP)

As solid propellants have certain advantages over liquidpropellants, composites may be more desirable forsome applications than the familiar black powderformulations. All solid propellants possess a highdegree of reliability by virtue of design. Once ignited, a

PRINCIPLES OF SOLID PROPELLANTROCKETS

solid rocket normally operates according to a presetthrust program, which is primarily determined by theconfiguration of the propellant grain. The amount ofthrust which may be obtained from a given grain designis largely determined by the propellant composition.Composite propellants burn at higher temperatures andpressures than black powder, with a net result that poundfor pound, they can deliver about two and one half timesthe power of a black powder motor.

Fundamental to the design of any solid propellant rocketis a simple geometric principal: The burning surface of asolid propellant recedes in parallel layers. Because ofthis, solid motors are self-stabilizing. That is, shouldsmall convex or concave irregularities arise on theburning surface, as would happen if an air bubble wastrapped in the propellant grain, such irregularities woulddisappear as burning proceeds. This is significantbecause as bubbles are encountered, the burning surfaceof the propellant and consequently the internal pressureof the motor and the burn rate of the propellant areincreased. When this exceeds the design parameters ofthe motor, rapid overpressurization occurs, leading to acatastrophic failure.

The burning rate, r, of a solid propellant is the linear rateof consumption in a direction normal to the burningsurface. It typically ranges from 0.1 to 2.0 inches persecond and is primarily influenced by combustionpressure, propellant composition; and to a lesser degreeby the ambient grain temperature and the velocity of gasflow past the burning surface. Burning rate may beexpressed by the following equation:

Equation #1

PROPELLANT CHARACTERISTICS

AP COMPOSITE BASICS

The burning rate, r, is in inches per second; the pressureof combustion is in pounds per square inch; and 'a' and'n' are constants.

Propellant composition and pre-ignition temperatureare the determinants for the value of the constant 'a',which ranges between .002 and .05. The pressure orburning rate exponent 'n' is solely a function of thepropellant formulation with negligible influence of the

Back in the Pre-Research days of Tripoli, long before Prof. Terry McCreary’s Experimental Composite Book,Randy Sobczak, of PlasmaJet wrote a paper on AP Composite Basics. It covered all aspects of AP CompositeMotor Making before reloadable motors existed. It was actually published in The TRIPOLITAN, though fewpicked up on it and, with the author’s permission, is republished here for your review.

r = aP cn


bulk temperature. Typical values for the burning rateexponent range from 0.2 to 0.5, but in some cases mayvary between 0 and 0.9. The burning rate exponent isof critical importance in maintaining the stableoperating pressure of any rocket motor.

During combustion, a rocket functions in a state ofdynamic equilibrium. A stable operating pressure ismaintained by a delicate balance between the rate atwhich gas is being generated by the burning propellantand the rate at which it is being expelled through thenozzle. This is affected by an area ratio between thepropellant burning area and the nozzle throat area.This ratio is known as Kn, where:

Equation #2

and the specific formulation of a propellant and itsburning rate exponent, will determine the operatingpressure of a given motor. The relationship between andoperating pressure of a given motor is expressed by theequation:

Equation #3

where 'P' is the pressure in pounds per square inch, 'B' isa constant for a specific propellant, and 'n' is the samepressure or burning rate exponent as appears in the burnrate equation (Equation #1).

Small changes in the value of 'n' can lead to significantchanges in the operating pressure of a rocket motor andconsequently in the propellant's burn rate which wasshown in Equation #1 to be pressure dependent. If 'n' is0.3, (1-n) is .7 and 1/(1-n) is 1.42. A 20% increase inburning area would cause a 30% increase in pressure.But if 'n' should be 0.8, (1-n) is .2, and 1/(1-n) would be5. For such a propellant, that same 20% increase inburning area would cause a 148% increase in pressure.It becomes evident that propellants with high exponentsare to be avoided, as small variations in the burningsurface such as when bubbles become trapped withinthe propellant grain or cracks are present in the grain,lead to greatly magnified variations in chamberpressure.

The quantity of energy available from a rocketpropellant is determined by the chemical nature of theoxidizer and fuel molecules, as well as by the chemicalnature of the reaction gas products. This is mostconveniently expressed as specific impulse, Isp, which

SPECIFIC IMPULSE

is an effective measure of the performance of variouspropellant systems compared to one another. Thehigher the Isp value, the more efficient the propellant.Specific impulse may be considered to be the amount ofthrust, which is available for each pound per second ofpropellant burned. It is the reciprocal of the specificconsumption of propellant and is expressed in poundsof thrust per pound of propellant used per second. Thisis found by dividing the thrust, or total impulse (It) bythe weight flow rate expressed in pounds.

Equation #4

The range of specific impulse for most ammoniumperchlorate composite propellants may vary from near170 to approximately 230, with a common figurearound 200. In comparison, the Isp of black powder isbetween 70 and 80, roughly two and one half times lessthan that of a composite propellant. The total impulse ofa rocket motor describes the total amount of energystored in that motor. Thus, the total impulse of a motorcontaining 2 Lbs of propellant having a specificimpulse of 210 would be: It= Isp x Wt or 210 sec. x 2.0lbs = 420 lb-sec. Depending on the grain design, thismotor would be capable of producing 420 pounds ofthrust for 1 second, 105 pounds of thrust for 4 seconds,or any combination of thrust x time which would comeout with a product of 420 lb-sec.

Regardless of the propellant system used or the Isp of agiven propellant, the design of a nozzle is afundamental criterion in the construction of a rocketmotor. The rocket nozzle functions to transform theheat energy of combustion into the kinetic energy of ahigh velocity gas stream with the maximum possibleefficiency. Nozzle theory is based upon the laws ofthermodynamics, gas dynamics and fluid dynamics. Abasic understanding of nozzle function is of paramountimportance to the proper design of rocket motors.Derivations and discussion of the fundamental lawspertaining to the conservation of matter and energy, andof the dynamic processes involved, are presented indepth in the proper texts for the interested reader topursue. The following discussion of combustionprocess and nozzle theory will provide the basis forgeneral understanding.

The combustion of composite propellants occurs indifferent phases, with the oxidizer particles

MOTOR DYNAMICS

COMBUSTION OF COMPOSITEPROPELLANTS

Kn =Ab

tA

P = B(Kn)1

1-n

Isp =I

Wt

t


decomposing in the midst of the decomposing fuelmatrix.Ammonium perchlorate itself does not melt, butrather undergoes an exothermic decompositionresembling that of homogenous propellants. Adjacentstreams of fuel-rich and oxidizer--rich gasses rise fromthe surface, and immediate reaction is not possible untilmixing by diffusion is complete. The combustionprocess takes place in three distinct zones, the foam,fizz, and flame zones. At the combustion surface, thegas velocity is relatively small and possesses littlekinetic energy. It is in the flame zone that the finalreaction occurs and the majority of the heat and gaseousproducts are evolved. There, the high pressure of theexpanding gasses forces the gas particles to the rear,causing a slight decrease in potential energy at thenozzle entrance, but an increase in velocity.

There are three basic types of rocket nozzles: subsonic,sonic, and supersonic. It is the supersonic nozzle whichis of interest, consisting of three parts; a convergentsection, a throat of specific diameter and therefore area,and a divergent section. Nozzles of this type are oftencalled DeLavel nozzles, after their inventor, and may bethought of as two cones joined at their verticles by ashort, straight throat section, with all transitions beingsmooth so as to avoid disturbances in gas flow.

When gas enters the converging portion of the nozzle,the decreasing cross-sectional area causes the flow tospeed up. The maximum velocity which can be obtainedin the converging portion of the nozzle occurs at thethroat, and corresponds at that section to the local sonicvelocity. In practice, this will not occur unless the ratioof chamber pressure to throat pressure reaches a certainminimum value, the critical pressure ratio, whichcorresponds to thirty-two pounds per square inchabsolute or twice the ambient atmospheric pressure.Once this chamber pressure has been reached, thevelocity of gas at the throat will always correspond tothe critical throat velocity regardless of further chamberpressure increases.

Once the exhaust gas has reached sonic velocity, severalof its major flow properties change. This may be used toadvantage by the addition of a diverging section to thenozzle. Gas velocity increases into the supersonic rangeand pressure decreases, as expansion of the exhaustgasses takes place over the entire length of the divergentsection. Optimum expansion occurs when the pressureof the exhaust gasses at the exit plane of the nozzle isequal to the ambient pressure. This will be found at aspecific cross-sectional area of the nozzle exit, Ae, of agiven rocket, which may be related to the throat cross-

NOZZLE THEORY

sectional area as the nozzle area expansion ratio.

Equation #5

Thrust is lost when the area ratio varies in eitherdirection from the optimum. There are upper and lowerpressure limits for all propellants. Very high chamberpressures, above 6000 psia for most propellants, causeerratic and rapid burning, frequently leading tocatastrophic failure of the motor. On the other hand,many propellants will not support combustion at lowpressures. This may be advantageous as a safety feature,but must be taken into account when designing the graingeometry so as to minimize the unburned propellantresidue or "slivers" at motor burnout.

grain having a fixed surface area exposed tocombustion, there exists a maximum effective throatarea which will maintain a chamber pressure highenough to support combustion. Most solid rocketsemploy nozzles which will maintain chamber pressureswell above this critical limit.

For a particularpropellant

It is hoped that the preceding discussion provides thebasis for an appreciation of the intricacies involved withdesigning solid propellant systems. Let us now examinethe practical application of theory.

The mechanical requirements for nozzle fabrication arequite stringent. The material utilized must exhibit goodmachinability or ease of fabrication while retainingexcellent resistance to erosive change under the mostextreme conditions. The throat section is usually madefrom graphite or some non-ablative material which issurrounded by insulation to keep the outside structuralmaterial cool. The inside of the exit cone in large motorsis made from such materials as asbestos-phenolicbacked by an external structure to contain the nozzlepressure. Fiberglass phenolic laminates and ceramicshave all seen application as nozzle materials in small

NOZZLE DESIGN

E = AA

e

t

SLIVERS


rocket motors, but problems with cost, fabrication, anderosion have limited their use. By far, the most simpleand economic solution has been the full diametergraphite nozzle, machined from readily available rodstock which fits snugly into the inside diameter of themotor. The major drawback of this system is that thegraphite acts as a heat sink and in larger or long burningmotors may cause charring of non-metallic motorcasings. This difficulty may be circumvented bysuspending a graphite nozzle insert of significantlysmaller diameter than the motor case in a hightemperature epoxy which provides insulation for thecasing. Thermosetting high temperature injectionmolded plastics combine the erosion resistance ofgraphite with the insulating properties needed to protectthe motor wall, and are currently enjoying widespreadusage. Their major disadvantage is the high initial costof tool and die fabrication.

When designing a rocket nozzle, consideration must begiven to the angle of convergence, the angle ofdivergence, the nozzle area ratio, and the nozzle throatarea with its relationship to the propellant burning area,Kn.All of these parameters must be precisely calculatedfrom complex equations in order to optimize motorperformance and efficiency. However, some usefulgeneralizations may be drawn. The convergent conehalf-angle, varies between 15 degrees and 45 degrees,with 30 degrees being a good compromise and 45degrees being a more space and material efficientchoice. For the divergent cone, a nozzle half-angle,between 12 degrees and 18 degrees, has been foundexperimentally to be optimum, with 15 degrees being agood choice in high thrust motors.

The nozzle area expansion ratio, E, that is the ratio ofnozzle exit area to throat area, deserves more attention.Expansion of the exiting gasses is ideal when theexternal pressure is equal to the nozzle exit pressure, andthe motor delivers maximum thrust. An underexpanding nozzle will discharge the exhaust at apressure greater than the external pressure, because theexit area is too small. Thus, the gas expansion isincomplete within the nozzle and continues outside. Thenozzle exit pressure is higher than the atmosphericpressure. An over expanding nozzle is one in whichgasses are expanded to a lower pressure than theexternal pressure due to it having an exit area which istoo large. Separation of the gas jet from the nozzle wallwill result, reducing the exhaust velocity, therebyleading to a loss of thrust. The formation of shock wavesis also of concern with improperly designed nozzles. Itis best to design a nozzle with the optimum expansionratio or one which under expands the gas jet slightly. A

ratio of 1:3 to 1:4 is appropriate for composite motorsystems operating in the 300 psia to 400 psia range. Athigher pressures, the area ratio would increase. Thiswould also be the case for sounding rockets operating athigh altitudes where the exhaust gasses must expandfurther so that they can match the lower atmosphericpressure at the nozzle exit plane. The upper stage nozzleof such vehicles often have exit diameters ranging fromfive to seven times the throat diameter.

The final parameter which must be considered in nozzledesign is the throat area. As has been previouslydiscussed, the propellant characteristics of burn rate, "r"and pressure exponent, "n" must be considered whenchoosing the area of the throat. By substituting Equation#2 into Equation #3, the relationship of the throat area tothe operating pressure and burning area isdemonstrated:

Equation #6

THROAT DIAMETER

The operating pressure is subject to the weight andstrength limitations of the motor casing, while theburning area is a function of grain geometry.

All of the many variations of grain geometry fall intothree broad classes. Regressive grains have a largeinitial surface area which decreases as burningprogresses. Neutral burning grains will maintain aconstant burning area and progressive grains exhibit anincrease in burning surface as propellant is consumed.Each of these categories have inherent advantages anddisadvantages. For example, a moonburning graingeometry with regressive characteristics would beuseful in long burning, relatively slow traveling rocketsgoing to high altitude. The high thrust of this initiallyprogressive geometry would be desired at thebeginning of the flight when vehicle weight is high. Asthe rocket gains altitude, its mass is decreasing aspropellant is consumed.At the same time, the frictionalresistance or drag decreases as the atmosphere thins,and the grain geometry becomes regressive as it burnsout to a sustainer phase. In this way, a subsonic velocitywith its lower drag coefficient may be maintained, thusoptimizing vehicle altitude. When constant thrust isdesired, neutral burning grains are called for. This ischaracteristic of an endburning charge consisting of asolid cylindrical section of propellant which isinhibited on all surfaces, except at one end, so that itwill burn like a cigarette. The end is often machined

GRAIN GEOMETRY

P = B( )AA

e

t

11-n

TRIPOLI GERLACH NEWS PAGE 11

into a cone shape to increase the initial surface area.Such charges have a constant burning area (unless theend is coned) and have a very long burning time, withvery limited applicability for composite rockets.Hollow rod or rod and tube geometries will also provideneutral burning characteristics while exposing morepropellant surface area and providing higher thrustlevels than are available with endburning designs.

Progressive burning characteristics are found ininternal burning case-bonded charges. An internalburning charge, as its name implies, burns outwardlyfrom the internal perforation. It may have single ormultiple ports in a variety of shapes, which provide

these motors with high levels of thrust over moderatelyshort burn times. Such motors are useful in boostinglarge or heavy payloads. These propellant grains may beseverely stressed during motor function, particularly atthe internal perforation. Moderate tensile strength 100-150 psi) and good elongation (30-70%) are necessaryfor cases bonded charges.

Thrust programs may be designed so as to combine thecharacteristics of progressive, neutral or regressiveburning in a single motor. For an internal burning stargrain shape, the initial surface area can be made nearlyequal to the outside or final area of the grain.

During burning, the surface will increase slightly, thendecrease to a minimum, then increase gradually until thepoints of the star reach the outer surface. At that point,there will be left semi-lunar sections, or slivers, whichwill burn out with ever decreasing area. Progressive-regressive profiles are characteristic of moonburngrains first designed by Bill Wood, which utilize anoffset core, which is approximately 25% of the graindiameter. Initially, such motors are progressive core-burners up to the point where the expanding corereaches the case wall. Then, the remaining propellant,which is now in the shape of a crescent moon, burnsregressively. A "D" shaped grain consisting of a solidrod with a thin slice cut off lengthwise will exhibit asimilar thrust profile.

A working knowledge of the interrelationships alreadydiscussed between operating pressure, propellant burnrate, and propellant surface area leads to a secondmanner in which the thrust program of a given motormay be varied by changing the throat diameter. By usingan ablative nozzle material, which erodes at a knownrate during motor function, the Kn of a motor may bechanged in conjunction with the surface area ofpropellant burning. Thus, it is possible to design aneutral burning motor by combining a progressive graingeometry with an ablative nozzle, thus maintaining a xxconstant Kn burn rate, and operating pressure.

PROPELLANT GRAINCROSS SECTIONALS

Internal Burning Core

Bates Grain

Internal Star

Moon Burn

Rod & Tube

Endburn

Casing

Inhibitor

Propellant

NEUTRAL

PROGRESSIVE

REGRESSIVE

TIME

GRAIN GEOMETRYBURNING CHARACTERISTICS

Endburn

Internal Burning Core

Bates Grain

Internal Star

Moonburn

Rod & Tube


TRIPOLI GERLACH NEWSPAGE12

MATERIALS

MOTOR CASINGSThe combustion chamber or motor casing functions as asimple pressure vessel. High strength and low weightare of primary concern when choosing a casingmaterial. For large rockets in the early years, steel wasbeen a frequent choice. A case failure can projectshrapnel quite a distance and cause great damage.Aluminum is a lighter metal, but the thicker wallsnecessary to fabricate a casing of the same strength assteel, results in no net advantage other than in a casefailure aluminum has proven to have a low projectilecharacter. It is obvious with all things consideredaluminum is your metal of choice. Of the metals,titanium is by far the strongest and lightest, but it is verydifficult to fabricate and is an expensive alternative.

Aluminum motor cases are the rule for reloadablemotors. They hold up well to repeated use and arerelatively easy to fabricate.

Smaller composite rocket motor casings, referred to assingle use motors, are generally made from phenolicpaper or cloth, or fiberglass. These materials areexceptionally strong and very light, and possess theadded advantage of having a decreased hazard potentialfrom shrapnel in the event of motor detonation.Fiberglass casings may be manufactured from glasscloth or by filament winding where plastic or morecommonly epoxy impregnated fiberglass is wound overa mandrel to form a tube. When the resin is cured, themandrel is removed to make the casing. When usingfilament, it is desirable to maintain a 55-60 degree angleof winding so as to prevent the formation of micro-porosities extending through the walls of the finishedcasing.

The phenolic based materials are lighter and lessexpensive than their fiberglass counterparts, as well asbeing more resistant to the high flame temperatures ofcomposite propellants. They are also much weaker thanfiberglass and therefore contraindicated for use in highpressure motors. In some cases, thin wall phenolic tubesare used as rigid liners into which propellant grains maybe cast, then machined and loaded into motors.

Many motor casings will incorporate a liner asinsulation. Most often it is fabricated from the samebinder as the propellant and filled with inert materialssuch as titanium dioxide, silicon dioxide, and/or otherhigh temperature resistant materials.Asbestos free highdensity gasket material, (1/32") available at automotivesupply stores, is an easy to use alternative when desired.

EPOXIES

OXIDIZER

Epoxy compounds have received widespread utilizationas the material of choice for sealing bulkhead and nozzleclosures of small rocket motors. These materials are welladapted for withstanding the high heat and high pressureenvironment of motor function. An almost infinitenumber of formulations are possible based on thespecific primary resin, modifying resin, and additivessuch as reactive diluents, bonding agents, surface activeagents, fillers, and curatives.

An excellent and readily available epoxy may be foundat local hobby shops which cater to radio controlairplane enthusiasts: SIG One Hour Cure Epoxy is amedium viscosity clear epoxy suitable for gluing fulldiameter nozzles and delays.

When "floating" a nozzle of substantially smallerdiameter than the motor casing, a reinforced, filledepoxy is required. These materials are commonly usedfor potting or encapsulating electronics components andare available in various viscosities and thermalconductivities. Biwax Corporation's Formula 411 workswell for this application.

The primary use of ammonium perchlorate, NH4CLO4,is as an oxidant in solid propellants. It is also used inexplosives, mines, shells, timing devices, andpyrotechnics. It is produced from anhydrous ammonia,aqueous hydrochloric acid, and sodium perchlorate.Ammonium perchlorate is a white crystalline solid witha molecular weight of 117.49 and specific gravity of1.95. It is slightly soluble in water. Pure ammoniumperchlorate is stable below 65.6°C and undergoes anendothermic reaction at 240°C, followed by twoexothermic steps at 275°C and 470°C. Contaminationwith metallic salts such as those of copper, chromium,and iron catalyzes the second decomposition step so thatit occurs at progressively lower temperatures as theimpurity concentration is increased. Ammoniumperchlorate is a strong oxidizer which is not explosiveunless contaminated. It constitutes an extreme firehazard in contact with oxidizable substances, organicmaterials, ammonium compounds, cyanides, sulfur andsulfur compounds, powdered metals, phosphorus andmetal salts. Strong acids may react with perchlorates togenerate perchloric acid, a dangerous explosive ifallowed to contact oxidizable materials. Ammoniumperchlorate crystals have piezoelectric properties, andm a y g e n e r a t e a c h a rg e u p o n s t r e s s d e -formation.Ammonium perchlorate contains 34%

THE PROPELLANT SYSTEM



available oxygen, considerably less than that of thesodium or potassium salts. Nevertheless, because of theweight fraction of solids in their combustion gasses,propellants containing it have overall performancecharacteristics exceeding that obtainable with either ofthe other two oxidizers. It also has the advantage of notproducing smoke.

Ammonium perchlorate propellants produce hydrogenchloride and other chlorine compounds duringcombustion. In high humidity or a moist atmosphere, thehydrogen chloride will condense into a dangerous fog ofhydrochloric acid. The exhaust gasses of these motorsare toxic, as well as being highly corrosive to manymaterials.

Ammonium perchlorate is produced in three ordinancegrades. The fine classified grade is available in 55 micronand 90 micron sizes, both coated with tricalciumphosphate (TCP) as an anti-caking agent. Regular-ClassI is 200 micron and Coarse-Class II is 400 micron in size.The latter two grades are offered with or without the TCPand may be rotary rounded, producing spheroidal grains.

The shape of the grains and particle size of the oxidizerare of critical importance in a propellant formulation,influencing the burning rate, processing properties andthe physical properties of the propellant. In general, adecrease in particle size results in an increase in burningrate, with the most significant effect in the submicronrange up to about one hundred microns. The effects ofcrystal size are sometimes so significant that a wholeseries of propellants can be made with the samecomposition by merely varying the particle size.

In practice, most composite propellants are multi-modal,consisting of several different sizes of oxidizer inspecific ratios. The larger 200µ and 400µ grains arerounded to spheres so as to present the smallest possiblesurface area per volume of oxidizer. The smaller crystalsof ground oxidizer will then fit into the intersticesbetween the larger particles. The net result is a propellentwith high solids loading which is not fuel rich and thusmaximizes the Isp and mechanical properties of thatpropellant.

PARTICLE SIZE

MULTI-MODALPROPELLANT SYSTEMS

BINDERHydroxyl terminated polybutadiene (HTPB) is a longchained clear liquid rubber polymer. First used as abinder and fuel in solid propellants by Aerojet in 1962,HTPB is chemically a polyurethane because it is cured

BINDER SYSTEM

with isocyanates. Reaction sites for cross linking areprovided by hydroxyl (-OH) radicals at several pointsalong each chain, as well as at terminal ends. It is thethree dimensional matrix of the cross-linked rubberchains which impart the important mechanicalproperties to a propellant. The ability of a propellant towithstand high strain rates is directly related to the lowtemperature properties of the binder, such as elongationand brittle point. In high solids loaded propellants, amodulus of 400©700 psi with good elongation andtensile properties is required, particularly when casebonding. With a glass transition temperature near orbelow –100, HTPB has excellent characteristics.

Boiling point 300-CSpecific gravity @25-C.90Viscosity (Brook) @25-C6000Strain capacity @-65-F25-35%

The actual mechanical properties of a specific propellantare a function of the exact formulation, i.e. by the sizedistribution and amount of solids, the ratio of binder tocuring agent, and the amount of plasticizer.

As previously mentioned, HTPB is cured byisocyanates. Some require an elevated temperature(oven cure) of 125°+F to activate, while others such asisophorone diisocyanate (IPDI) or PAPI; are active atroom temperature. Such curatives are usually present inthe range of 8-10% of the rubber content of thepropellant, based on calculation of the activity of theparticular agent used. These curatives are all toxiccompounds, with some more so than others. Among theroom temperature curing agents, toluene diisocyanategives the shortest pot life and is the most toxic. PAPI-901 and N-100 are two good choices for low toxicity andadequate pot life for room temperature curatives. Caremust be taken to ensure that all propellant componentsare kept dry, as isocyanate groups will react with water,producing a substituted urea and liberating CO2 in agassing reaction.

A number of very low viscosity plasticizing agents maybe added to a propellant for improved wettability whichwill allow higher solids loading and consequentlyimprove performance. These agents will improve themechanical properties of a propellant, retard oxidationand embrittlement to a certain extent, and when used as asignificant portion of the binder system (25-30%), willallow for some propellants to be pourable.

PHYSICALDATA

CURATIVE

PLASTICIZER

TRIPOLI GERLACH NEWSPAGE14 JANUARY 2013

Dioctyl adipate (DOA) is a high quality grade of D1-2-ethylhexyl adipate which is used as a diester fluid forsynthetic lubricants. This colorless liquid has low acuteoral toxicity, but is considered as a high health hazarddue to its mutagenic and carcinogenic effects.

Dioctyl azelate (DOZ) is a similar product with aslightly higher molecular weight and lower toxicity.

Isodecyl pelargonate, (IPDI), is another synthetic oil ofeven lower viscosity than DOA, which is an excellentplasticizer.

In effect, most any low viscosity type of oil may be usedas a plasticizing agent. The advantage of theaforementioned products lies in their wetting ability andultra low viscosity.

Most published propellant formulations will contain abonding agent. These compounds react with the surfaceof the ammonium perchlorate crystals (frequentlyreleasing gaseous ammonia) and facilitate actualbonding of the rubber to the crystal. Without such anagent, the oxidizer crystals are simply retainedphysically within the propellant, captured in effect by athree dimensional matrix of rubber. Without a bondingagent, crystals of oxidizer which are on cut or machinedsurfaces of a propellant grain will be lost duringprocessing, leaving a fuel rich surface.

TEPANOL (Dynamar Bonding Agent/Processing AidHX-878) is one such bonding agent and commonlycomprises 0.25% of the entire propellant formulation.Bonding agents greatly improve the mechanicalstrength and properties of a propellant, but are not ofsignificance in small rocket motors.

Finely divided metals are added to almost all compositepropellant formulations. These fuels provide a varietyof benefits and enter into some very complexinteractions during combustion. Spheroidal aluminumis perhaps the most commonly used metal in compositepropellants, found in various formulations from near1% up around 18%. The ballistic performance ofaluminized propellants is greatly increased, raising theIsp in the range of up to 10% when compared to thesame formulation without metal. It must be noted,however, that this effect is not significant in smallrocket motors, where the metal is not in the combustionchamber long enough to be consumed, and is mostlyexpelled in the exhaust gasses in the molten form. Theaddition of aluminum to a propellant will also serve to

BONDINGAGENTS

METALLIC FUELS

dampen acoustic oscillations, thus minimizing thepossibility of grain fracture at ignition, and also makingignition easier, especially in small motors. The net effectof aluminum of burn rate is usually not large and may bepositive or negative.

When considering metallic fuels other than aluminum,those with low molecular weight are desirable. Thosewhich might be of benefit to propellant application maybe determined by considering the density and the heat offormation of the metal oxide. Beryllium heads the listand has been reportedly used, but has the problem ofproducing toxic combustion products. Boron, lithiumand silicon all have higher heat content per gram thanaluminum, and are potential additives. Magnesium hasalso been used, but imposes a hardship on the binder dueto its lower heat content and lower density.

Numerous compounds are utilized to modify the burnrate of propellant systems. Most exert a positivecatalytic effect, while some such as oxamide decreasethe burn rate by insulating the heat wave and slowing theprogression of combustion. Addition of inertcompounds (chalk) or substituting less active oxidizers(ammonium chloride, sulfate) for a portion of theAPin apropellant will also slow the rate of burn.

By far, the majority of modifiers are catalysts which insome manner enhance the rate of burning. The effectswhich these compounds have on the dynamics ofcombustion is an exceedingly complex area of research.

At this point, it will suffice to say that catalysts exerttheir effects in relationship to combustion pressure andconcentration. An effective range may be determined,above which the increase of catalyst percentage hasdiminishing returns without significant increase in burnrate.

The following is a partial listing of some frequentlyutilized burn rate catalysts with brief comments abouteach one:

Manganese Dioxide (MnO ) - Positively catalyzes solidphase reactions. MnO is a strong positive catalyst forthe decomposition, but is a negative catalyst for thedeflagration ofA.P.

Iron (III) Oxide (Fe O ) - An excellent and readilyxxxxxxavailable catalyst. It will increase burn rate more thanMnO2 will at same level. Fe O promotes the complete

BURN RATE MODIFIERS

PROMOTERS2

2

2 2

2 2


decomposition ofAPat 270-280 degrees.

Chromium (III) Oxide (Ci O ) - Primarily enhances thelow temperature decomposition of AP, allowing thatreaction to go to completion. Chromium oxide exerts amuch greater effect on burn rate than manganesedioxide, and at the 2% level is superior to iron oxide inlow pressures, up to about 600 psia.

Copper Chromite (Cu Ci O or CuO - ÀCu Cr O ) - hasa significant effect, but varied on burn rate.Analysis hasshown that copper chromite catalysts differ fromcompany to company and may not even be the samefrom batch to batch. Propellants containing copperchromite become brittle and do not age well.

Cupric and Cuprous Oxide (CuO and CuO ) – Bothcatalyze the low and high temperature decomposition ofAP, and promote ignition. Cupric oxide (CuO) issuperior to copper chromite and even chromium oxideas a burning rate promoter.

2 2

2 2 7 2 3

2

Ferrocene (Dicyclopentadinyliron, C H Fe) and itsderivatives Catocene and N-Butylferrocene – Theseliquid burn rate promoters are based on two fivemembered cyclopentadinyl groups with a ferrous ion(Fe) sandwiched between. These compounds interactwith aluminum during combustion. Decreasing theparticle size, and thus increasing the surface area ofaluminum to react, will increase the burn rate ofpropellants containing these compounds.

The burn rate of a propellant may be decreased by thesubstitution of up to 20% of the oxidizer withammonium chloride or ammonium sulfate. Theaddition of zinc powder to a propellant will also slowthe rate of burn, while also generating a dense, blackexhaust. A fuel rich propellant, or one to which an inertcomponent such as calcium carbonate has been added,will also burn slower. The use of an inhibitingcompound, which will contribute to the combustionreaction, however, seems to be the more sensibleapproach.

10 10

BURN RATE INHIBITORS

Back in the “Way Back” days of early Tripoli severalpersons were researching and producing CompositeRocket Motors. It was a time of free pursuit. People likeGary Rosenfield, with AEROTECH, Scott Dixon, withVULCAN SYSTEMS, and John Rakonnen and hisPRODYNE were producing and selling CompositeMotors on the open market.

Ranked with this group of innovators was Randy Sobczakand PLASMAJET. Randy produced some veryaggressive motors that to this day are very sought after.

R a n d y b e g a nexperimenting withcomposites in the late70’s and teamed up withJohn Krell to formPLASMAJET.

To this day severalPLASMAJET Motors,tested by Tripoli TMTin the early year, keeppopping up and despitet h e i r a g e , w o r kexceptionally well.

RANDY SOBCZAKPLASMAJET

Randy Sobczak in the early PLASMAJET days testingpropellant at Smoke Creek, Nevada

PLASMAJET Motors tested and approved by TripoliTMT were:

These included PLASMJET’s leading core-burners,moonburners and smoket formulartions.

All expired around 1997

F-17G-39E-40

H-36H-83H-64

H-124H-64H-95

H-95H-116H-121

H-116H-121I-124

Randy & John Krell at Lucerne

TRIPOLI GERLACH NEWSPAGE16 JANUARY 2013

Head End Ignition is a way to ignite a rocketMotor internally from an on board computer oraltimeter. It is great for staging and evenapplicable to ground ignition as well. Thesystem is based on a forward closure that has adelay column recess or what is called a smokewell.

The forward closure will require a hole tappedlarge enough to accommodate a modelairplane engine Glow Plug, preferably a longtype Glow Plug. You will also need to drill andtap a side hole for a screw to secure the“device” into the forward closure so it will notblow out at ignition. The photos will show this.

The list of items required to accomplish thisare show in the photo to the right:

- Pyrodex (do not use black powder as itmust be compressed).

- Double Sided Masking tape

- Glow Plug, preferably the long type.

D -Any standard Forward Closure with a delayor smoke well.

- Piston (3) (Must be Fabricated)

- Extension Ring(2)

- Charge Holder(1)

A

B

C

E

F

G

(Must be Fabricated)

(Must be Fabricated)

H

I

J

K

- Two Wrenches

- Small Washer

- Large Washer

- Bolt & Nut

HEAD END IGNITIONOliver Schubert, after developing the No-Match Ejection System, developed an easy to build Head EndMotor Ignition System that will work on any type and size motor that contains a Smoke Well or a DelayCharge Well. The concept is presented here with no dimensions since it is applicable to all size motors.The photos and accompanying text should present the concept in an easy to understand method. - youwill need access to a metal lathe.

The three cross-sectioned pieces in thedrawing to the left need to be machinedfrom aluminum stock. Dimensions are notincluded as they will vary with motor size.

What is important is the extension Ringand Piston. What ever diameter you areworking with the length of the collar and

Piston (both marked in red)must be 50% the depth ofthe of the Charge Holderwell (marked in blue). This

The next very importantitem is the Bolt. If it is toosmall, the resulting core ofthe pellet is too small to pass

compresses the Pyrodex by1/3rd creating a nice solidpellet.

1 2 3

JANUARY 2013 PAGE 17TRIPOLI GERLACH NEWS

D.

F

Using a wrench on top and one on the bottom tighten the nut down on the bolt until the large washer is flush withthe Extension Ring. Once the unit is compressed let it sit for a minute before loosening and removing the nut & bolt.With the Extension Ring and Bolt removed you should have a semi-solid grain of compressed Pyrodex lookinglike .

Head End Ignition is a way to ignite a rocket Motor internally from an on board computer or altimeter. It is great forstaging and even applicable to ground ignition as well. The system is based on a forward closure that has a delaycolumn recess or what is called a smoke well.

A

D

B

E

C

F

The three cross-sectioned pieces in the drawing to theleft need to be machined from aluminum stock.Dimensions are not included as they will vary withmotor size.

What is important is the extension Ring and Piston.What ever diameter you are working with the length ofthe collar and Piston (both marked in red) must be 50%the depth of the of the Charge Holder well (marked inblue). This

The next very important item is the Bolt. If it is too

compresses the Pyrodex by 1/3rd creating anice solid pellet.

small, the resulting core ofthe pellet is too small topass all the gases throughfrom the rear of the pelletwhere the burning begins(you may end up with alarge "bang").

Determined the size of thecore experimentally bystarting with a small boltand going bigger until thepellet burns without

JANUARY 2013 TRIPOLI GERLACH NEWSPAGE18

Remove the last backing from the double sided tape and place the unit, PyroDex side down, into the well of the forwardclosure as in photo . Make sure the Container dimple and the hole drilled in the closure for a set screw are aligned. Thenpress to secure the Conatiner into place. Once set complete securing the Container with a set screw.

I

Besides the little hole drilled and tapped for a set screw theForward Closure must be modified on top by drilling andtapping a hole to accommodate the Glow Plug as shown inthe center photo .

Fill the Glow Plug hole with a small amount of PyroDex,not more than 1/8th inch. Install the Glow Plug and theclosure is ready for motor assembly.

J

The hole where the glow plug screws into has to be thelength of the glow plug thread plus ~1/16" so there is roomfor a bit of pyrodex powder. You may need to manufacturea disk that increases the back wall thickness of the rearclosure and press-fit it in, unless you’re making your ownclosure, before cutting the threads for the glow plug.

This unit is very reliable and has worked 100% of the timein repeated testing and actual use in several rocketryprojects such as Tom Blazanin’s DARQUE SOL FlyingSaucer Project and more recently in Ken Good’sInternally Staged DRAKE Project.

While this article is basically designed to present theconcept of a simple Head End Motor Ignitor it does coverthe bases. No dimensions are given as stated since motorsizes vary but this system is easy to understand and build.A small metal lathe is required and a free afternoon is allthat’s needed.

Any inquiries on this system can be answered by Oliver byE-Mail at: [email protected].

I J K

Place the unit, PyroDex side down, onto apiece of double sided masking tape as shownin photo . This will actually hold thePyroDex secure for final placement.

Note the “dimple” on the side of theContainer. This will mate up with a set screwin the forward closure and secure the unit inthe closure itself.

With the tape in place trim off the excess asshown in photo .

G

H

G H

A glow plug needs low voltage (1.5V) and high amperage (~3-5A). An e-match (which is what an altimeter isdesigned to fire) needs 9V at about 1A. Oliver’s NoMatch unit uses the output from the altimeter to turn on it's ownlow voltage battery that is capable of providing the 3-5Aneeded for the glow plug.Aside from the voltage difference,the 9V batteries we are using are not capable of providing the high amperage.

In a nutshell, the NoMatch changes the 9V 1Afrom the altimeter to 3V 5Afor the glow plug. FYI: I'm actually givingthe glow plug too much voltage which - if applied to long - will burn out the glow plug, but there is no standard 1.5 Vbattery available that can provide the high amperage required for the glow plug. Since altimeters turn on only for 1sec, this is not an issue.

Find out more about NO MATCH at: www.lvhq.net/nomatch

JANUARY 2013 PAGE 19TRIPOLI GERLACH NEWS

The quite and low profiled TomBlazanin lives in a semi-doublewide mobile home he calls TheCompound. It has an attachedgarage with outside access onlybut Tom has put it to good use.

Called Area 748, after hisaddress, he has organized theentire place for working onRockets and making motors.

Area 748 is laid out with toolsand supplies surrounding acentral portable work area thatdoubles with a built in table saw.Completed rockets are hungfrom the ceiling and the entireplace is in constant evolution.

With a metal lathe, drill presses,saws, compressor, mixers andtools; Area 748 is equipped tohandle any rocket venture and is openly shared with otherrocket people in need of space and tools..

As a partner with Dave Rose in GRAPHIX & STUFF andDT Research Rocketry a lot of the equipment and materialis joint ventured, which works out well.

With only a single car garage to work out of, additionalstorage space is naturally required.Area 748b (below)

unusually

shares space in GRAPHIX & STUFF’s work room to storeChemicals, Casters & Liners, other Tubing andEquipment.

AREA748c (below right), used to store additional rockets,shares a room with GRAPHIX & STUFF’s ApparelInventory. It is also used, when needed, as a place to drynewly painted rockets as well.

LETS VISIT TOM BLAZANIN

www.researchrocketry.com

JANUARY 2013 TRIPOLI GERLACH NEWSPAGE20

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READ THIS BEFORE OPENING PACKAGE:

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The Mass of This Product Contains the EnergyEquivalent of 85 Million Tons of TNT per Net Ounce ofWeight.

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: Some Quantum Physics Theories SuggestThat When the Consumer Is Not Directly Observing ThisProduct, It May Cease to Exist or Will Exist Only in a Vagueand Undetermined State.

TheSubatomic Particles (Electrons, Protons, etc.) ComprisingThis Product Are Exactly the Same in Every MeasurableRespect as Those Used in the Products of OtherManufacturers, and No Claim to the Contrary MayLegitimately Be Expressed or Implied.

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The EntirePhysical Universe, Including This Product, May One DayCollapse Back into an Infinitesimally Small Space. ShouldAnother Universe Subsequently Re-emerge, the Existence ofThis Product in That Universe Cannot Be Guaranteed.

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PRODUCT WARNING LABELSAs concerned citizens, and those particularly interested in science, we applaud the recent trend towards legislation thatrequires the prominent placing of warnings on products that present hazards to the general public. Yet we must alsooffer the cautionary thought that such warnings, however well-intentioned, merely scratch the surface of what is reallynecessary in this important area. This is especially true in light of the findings of 20th century physics.

We are therefore proposing that, as responsible citizen/scientists, we join together in an intensive push for new laws thatwill mandate the conspicuous placement of suitably informative warnings on the packaging of every product offeredfor sale in the United States ofAmerica. Our suggested list of warnings appears below.

TRIPOLI PREFECTURE # 138 › tripoligerlach › downloads › 2013-03-01.pdf · 2013-01-02 · Tripoli Gerlach was set up as a Research Rocketry Prefecture forTripoli members who

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