AIAA 93-2610

TnE Raplo AND Low Cosr DEvEI-oPMENTHYeRro Rocxer MoroR

K.W. Schulze and S.A. MeyerGencorp-Aerojet, lnc.luka, Mississippi

AIAA/SAEIASMHASEE29th Joint Propulsion

Conference and Exhibit

OFA

June 28-30, 1993 I Monterey, CA

For permlssbn to copy or repubush, contact the Amerlcan lnstltrrt d LrEn8utlcs and Astronautlcs

370 LEntant Prcmenade, S.W., Washlnglon, D.C. 20024

The Rapid and Low Cost Development of aHybrid Rocket Motor

K.W. Schulze* and S.A Meyer**GencorpAerojetIuka, Mississippi

AhstractThe purpose of this project was to verify the low

cost, safe development of a hybrid rocket motor from" off-the-shelf" hardware. The use of this type ofhardware reduces the development cost and increasesflexibility of such a system. Demonsrarion of suchtechnology was accomplished in a period of six monthsfor incorporation into a flight concept hybrid rocketmotor. Multiple test firings were conducted withrelative ease and without risk of hazardous faih:res dueto the inherent safety and reliability of hybrid rocketsystems. The tests yielded design experience urilizingrecycled thermoplastic material as fuel. Thistechnology has potenlial as a low cost propulsionsystem with applications ranging from gas generatorsfor attitude control rhrusters, rocket take-off assist to anexo-atrnospheric skip glide vehicle.

In tro d u cti on

This project was independently developed andsuccessfully demonstrated the viabiliry of thermoplasticfuels use in a hybrid rocket motor. No corporate orgovernment funds were used to develop or test thishybrid rocket motor design. Therefore, cost was a keyfacnr in every aspect of this demonstration.

The static test hybrid motor consists of a fuelsection, contained within a stainless steel combustionchamber, with a high pressrue source of gaseous oxygento support combusrion. All of the required hardware fortesring was used as off the self, instead of built to adesign The use of this type of hardware during theprototype phase of a project reduces the developmentcost and increases flexibility of such a system.Demonstration of such technology was accomplished ina period of less than six months for the static motor testseries. The motor components include a filamentwound graphite/epoxy pressure bottle, a monolithicextruded gaphite nozzle throat, glass reinforced phenolicinsulator sleeve, stainless steel tubing and associatedhardware, solenoid manifold valves, pneuma[ic actuatorvalves, two-st,age regulators, and requiredinsrnrmentation for regression rate model validadon.

BackgroundThe research and development of hybrid rocket

motors is not new. In fact, a majority of work in thisspecialty was accomplished in the 1960's. Renewedinterest in hybrid rocket motor development has gainedconsiderable momentum. Many proJects are focused onhybrid rocket motor development from lab scale testmotors to AMROC's successful testing of a large250,000 lbf-thrust motor, ar Philtips Lab in California.

Inspiration for this project began with a

presentation given by Allan Holzman of UnitedTechnologies Chemical Systems Division at the 27thJoint Propulsion Conference in 1989, otr the safery andsimplicity of hybrid rocket motors. During hispresentation he gave a demonstration by firing severaldifferent solid fuel types using a low pressure oxygensystem with a butane spark ignition. This' wasconvincing evidence that hybrid rocket systems are truly

safe and simple. Suttonl sums up the main advantagesand disadvanLages for hybrid rocket systems as follows.

. S top-stary'restart capabiliries

. Environmentally safe exhaust characteristics formany propellant combinations

. Higher potential reliability than either solid or bi-propellant systems

. Relative to liquid engines, the hybrid requires halfof the pumps and plumbing since only one liquidor gas component must be handled

. Relative to a solid rocket, the fuel grain is strongerand inert and therefore insensitive to over pressuredue to cracks, unbonds and voids which occuroccasionally in solid propellant grains

. Relatively low system cost

. Higher specific l1npulse than solid motors andhigher-density impulse than liquid bi-propellantengines

. The ability to change motor thrust smoothly over a

wide range on demand.

Cop),ri.-qht @ 1993 by the American+ Senior Project Engi.neer, AIAA:k:i Pro1ect Engineer

Institute of Aeronautics andmember

Page I

Astronautics, Inc. All rights reserved.

The following are the main disadvantages of hybridrocket motors;

. Mixture ratio and hence specific impulse will varysomewhat during steady-state operation andthrottling

. Nominal steady-state combustion efficiencies in therange 93 to 97 Vo are slightly lower than liquid orsolid systems

. Lower system density impulse and thus a largervolume than solid propellant motors

. Some fuel sliver must usually be retained in thecombustion chamber at the end-of-burn, whichslightly reduces motor mass fraction.

The disadvantages of hybrid rocket systems pose achallenge to the propulsion designer. However, hybridrocket systems offer gfeater overall benefits for specificapplications than either solid or liquid propulsionsystems because of their inherent safety, simpliciry andlower cost.

ApproachThe long term and ultimate goal of this project is

to develop a 300 to 400 lbf-thrust flight concept hybridrocket molor to demonstrate the inherent system safety,flexibility, reliability, and lower cost for a flightapplication. To achieve that goal, an understanding ofthe key design parameters and acquisition of thenecessry Mta to optimtze a flight system was required.A series of static tests was conducted to gain theexperience and datz necessary to develop a low costflight concept hybrid rocket motor. Gaseous oxygenwas selected to avoid some of the problems of liquidoxygen and the sensitivity to injector design. I[ is alsoinexpensive at $0.72 per pound or $15 per K-bottle. Itwas therefore possible to complete a number of tests toacquire the necessary data to validate the burn rate modelfor thermoplastic fuels.

High Density Polyethylene was investigated asone of the viable thermoplastic fuels. HDPE is inertand safe to handle. It is commercially available inmany forms and is easy to machine or cast. It is a goodthermal insulator (Kc = 0.45 Btu/lb-"F) and a- m6lttemperature of 400 to 800 oF. The theoreticalperformance is very good with an Isp vac = 284 (ldf-sec[bm). HTPB has an Isp vac = 359 and as a solidrocket fuel the Isp vac = 279. The other benefits ofHDPE is that the combustion products areenvironmentally safe (HZO and CO2). Recycled HDPEis an extremely low cost fuel at $0.05 per pound ascompared with HTPB at $5.00 to $8.00 per pound.

Hardware for testing was incorporated as- built,instead of built to a design, to keep costs as low aspossible and still acquire the necess ry design data.Preliminary data was obtained by using a pipe chambertest motor (Figure 1) and gradually inco{porated flight

design hardware and materials (Figure 2). Thisminimizes the effects of multiple design changes andvariable interaction and allowed for direct correlation oftest results to a specific design parameter. The inherentsystem safety and robustness allowed for rapidprogression through the test series.

A non-pressure dependent burn rate equation wasused to develop a regression rate model for various typesof thermoplastic fuel sections. Existing technicalliterature does not provide data for predictingperformance of this specific oxidi zer-to-fuelcombinations; ttris data was necessary to further developthe flight design concept. A preliminary model forpolyethylene plastic (Table 1) was used to calculate thepreliminary sizing of the static test chamber panmeterssuch as; fuel section bore diameter, nozzle throatdiameter, and preliminary thmst and burn time.

Burn Rate Model

Several fuel regression rate equations have beenproposed for hybrid rocket motors in the past. Most areattempts at following the behavior observed in solidrocket motors, given by the expresslon;

ro = aPc)n (1)

In hybrid rocket motors, it has been obsenred thatthe semi-empirical regression rate equation is dependenton the mass flow per areil or mass flux rate.

16 - b(Gox)m (2)

However, the following equation accounts for themass flux rate of fuel as;

ro=p t(riro*+pf ro SAy*ApJo (3)

which provides a more accurate description of theregrcssion behavior within a hybrid rocket motor.

Test Objectives

The primary objectives for the first threepreliminary tests were to perfonn a system check-out ofall pressure lines, valves and fittings and to test theignitability of the thermoplastic (Polyethylene) atambient conditions and to maintain a continuous burnwith a steady flow of oxygen. To keep the ignitionsystem simple, &tr Estes "C6" or "A3 " model rocketmotor was used to initiate combustion of thethermoplastic. Two thermoplastic fuel grains ( HDPE& R-HDPE ) and three oxygen mass flow rates (Goxvalve settings of 1.5, 2.5 &.6 turns) were evaluated (see

Table 2) n obtain preliminary burn rate information.

Test # 1 through #3 used the "pipe" chamber as

illustrated in Figure 1, to perform a system check-out ofall pressure lines, valves and fittings and test theignitability of the thermoplastic (Polyethylene) at

Page 2

ambient conditions and maintain a continuous burnwith a steady flow of oxygen. In addition, a meteringvalve was incolporated to control the oxygen flow ratethrough the thermoplastic fuel sections. For Test #2and #3, anozzle throat of 3116 inches dia. was drilled inthe aft closure to evaluate the burn rate at chamberpressure. A center perforated fuel grain with a diameterof 314 inch was used for all three grains. Test #3evaluated the use of recycled high density polyethylene(R-HDPE) thermoplastic. This material is alsocommercially available and is black which allows formore adequate shielding of the steel chamber fromradiant heat effects.

Information gained in the preliminary test serieswas used to size the static test chamber parameters.Test #4A through Test #58 utilized the "static test"hardware illusrated in Figure 2. A monolithic graphite

GK-71) nozzle throat was incorporated with a glassfilled phenolic insulator sleeve to protect the stainlesssteel hardware from the high temperature gases(approximately 5700 oF).

T est ConfigurationThe "pipe" chamber test hardware was a 2-inch

schedule 40 galvanized steel pipe threaded to an AII{SI150 flange and bolted to a closed flange with twothreaded ports. One port was used as the oxygeninjector and ilte second port held the igniter and pressuregage assembly (Figure l). A thermoplastic fuel grainwas bonded and sealed into the 6 inch section of pipe.Oxygen flow was initiated by activating a solenoidmanifold valve which actuated a Nupro pneumatic valvein the oxygen line. The line pressure was set at thetwo-stage regulator at approximately 300 psig.Ignition was triggered remotely via an electric switchand an Estes "C6" motor.

The static test motor consisted of a stainless steelchamber was made by Varian for Ultra high vacuumapptications with a 0.060 inch wall thickness (Figure2). All of the stainless steel components werepurchased as surplus equipment. A 2.36 inch by 1 1.5

inch high density polyethylene fuel grain (TIVAR 100

produced by Poly HI Solidur) was sealed within the

stainless steel chamber. This fuel grain also protectedthe chamber from the hot gas environment. The head

end closure contained four po{S. The main port wasused as the oxygen injector. The second port held theigniter cartridge. The igniter cartridge contained a Estes"A3" size model rocket motor. The two smaller portswere used as the nitrogen purye and pressure transducerpon (0-200 psia) and for a second pressure transducer (0-

1500 psia). Because of the narrow case tube section, aglass reinforced phenolic insulator sleeve was necessaryto protect the metal case. All exposed metal surfaceswere protected using EPDM silicon rubber. The nozzleassembly used a stainless steel flange closure with amonolithic isomolded (rather than extruded) EK-71graphite throat insert. This gfaphite insert was bondedto the stainless steel using an RTV silicone. This thin

layer of RTV protected the stainless steel closure fromthe white hot graphite nozzle.

Material & Cost

The following table is a listing of some of themarn components and costs. The total cost for thehybrid rocket motor was $510. This low cost was theresult of using existing or surplus hardware wheneverpossible. Some machine work was required to modifythe hardware for use as a hybrid rocket motor. Manyitems were also samples from vendors which made thisexperiment a cooperative effort.

Ultra High Vacuum Chamber (3Mss) $17.00Varian Conflat Closure with Four ports $18.00Glass filled Phenolic Insulator NCMonolithic Graphite Nozzle NC

@K-7r & HLM85)Stainless Steel Lines, Fittings & Regulators $220.00lvlachine Work $2t0.00Gox, K - bottle (2200 psi) $15.00GN2, Q - bottle (2200 psi) $9.002.36" x !2" F{DPE Thermoplastic Fuel NC2.36" x 12" R-HDPE Thermoplastic Fuel NC2.36- x 12" PMMA Thermoplastic Fuel $21.00

Test Arrangement

Gaseous oxygen (Gox) and nitrogen (GN2) stored

in high pressure cylinders (K - bottles at 2200 psig)

were used and controlled by pneumatically actuatedvalves controlled by a manifold of solenoid valves(Figure 3). The nitrogen gas was used to purge and

cool the chamber after each static test. The flow ofgaseous oxygen was controlled by a metering valve,

which was p e-set by the number of turns. Thismetering valve was calibrated at a setting of 1.5 turns

by using a known volume and pressure of gas and

recording the pressure drop with time. The oxygen linepressure upstream of the metering valve and thechamber pressure were measured and recorded. To keep

the ignition System simple, &r Estes rocket motor("C6" or "A3" model) was used to initiate combustionof the thermoplastic. A thrust adapter was used totransfer thrust loads to an Eaton-Lebow (model 3169)load cell. A horizontal test stand was constructed with a

linear bearing table and load cell mounted to a 314 inchplate strucnue (Figure 3).

InstrumentationData was recorded using the BAKI(ER model2570-

P system data acquisition system (DAS) provided byPacific Instruments, Inc. The model 2570-P is a

portable waveform analyzer for recording and analyzingtransient events. The system capabilities include; 12

channels with & K samples per channel memory and

up to 50 Ksps digital recording rate, remote triggeroption and menu driven set-up and dataanalysis features

including FFT & Spectral Analysis options.Transducer signal conditioning was provided by PacificInstruments, Inc. using two Pacific Instruments model

Page 3

321013215 Transducer Conditioning Amplifiers forfour channels of analog input to the Bakker 2570-PDAS. The digital data was stored on a 1.44 lv{byte IBMformatted disk for frrrther analysis.

Sensors used to acquire data were; one 2 Klbs loadcell (Eaton-Lebow, model 3169), a 0-200psia Pressuretransducer (Eaton-PSD) and a 0- l500psia (Paine)pressure transducer. A 0400psig gage was used forupstream pressure measurements, which was recorded byvideo camera. One high temperature internalthermocouple was planned but not used in this series oftests.

Two video cameras were used to record the eventsand to establish the reference time (preliminary testseries). The test controls, video monitor, Bakker DASwere located 7 5 feet from the test stand behind aprotective wall.

ResultsThe Seven tests that were conducted are outlined in

Table 2 and were divided into a preliminary series and astatic test series.

The polyethylene thermoplastic (Test #1) wasignited easily and burned for approximately 20 seconds,producing a smoke free plume eight feet in lengthbefore shut-down. The test was a success and yieldedthe first opportunity to estimate fuel regression rate ofapproximately 0.0281 in/sec for a Gox valve setting of6 turns.

Test #2 used the same chamber as in Test #1, buthad a3ll6 inch nozzle throat in the threaded aft closure.The Gox metering valve was Set to 2.5 turns whichyielded a 100 psig AP between the oxygen injectorpressure and the chamber pressure. Ignition wasinitiated with a Estes "C6" model rocket motor. Thepolyethylene successfully maintained a burn withindications of peak chamber pressure of 220psig.Pressure was maintained for 0.6 seconds until thegalvanized steel nozzle throat eroded to match the fuelgrain bore diameter (1 inch dia.). Motor burn was

maintained for ll.4 seconds yielding an averageregression rate of approximately 0.0155 in/sec. Flametemperature was approximately 4500'F based on theerosion of the galvanizeA steel closure.

Test #3 used recycled polyethylene thermoplasticand a Gox valve sening of 1.5 turns. A pressure of 190

psi was maintained for 1.6 seconds before the galvantzedsteel nozzle throat eroded to match the fuel gfain bore

diameter. The burn time for this test was 18 secondswhich yielded an average fuel regression rate of 0.0057in/sec.

Test #4A &.4B used the same HDPE fuel grain(TIVAR 100). The burn time was l4.l seconds for aGox valve setting of 2, resulted in a chamber pressure

of 200 psi. The regression rate for test #4A was .005

in/sec. Test #48 had a 15 second burn time for a goxvalve setting of 4. The regression rate was calculated at

.0075 in/sec. For the gox valve settings on tests #44and #48, Do nozzle throat erosion was observed.

Test #5A & 58 used a new HDPE fuel grain

GIVAR 100). The burn time was 15 seconds for a Goxvalve setting of 6, resulted in a chamber pressure of 207psi. The regression rate for test #5A was .00155 in/sec.Test #58 had a burn time of 18 seconds for a Gox valvesetting of 6. The regression rate was calculated at.00155 in/sec. For the Gox valve settings on tests #5Aand #58, the nozzle throat erosion rate was estimated at

.0385 in/sec. At these Gox valve settings, the graphitematerial was reacting with the high temperature oxygengas even though the combustion mixture was less than

stoichemetric (3.42 OIF ratio). Figure 4 is an exampleof the thrust trace measured on Test #5B. It was

interesting to note that the loadbike which occurredafter ignition, was due to a delay charge in the Estesmodel rocket motor. At the end of Test #5A & 58, the

hybrid motor was shut down for 2 seconds and restarted

with a smooth transition to steady state operation.

During the posttest examination of the testhardware, it was noted that there was minor burnthrough to the metal case without heat damage to the_

case. On the preliminary tests, there was evidence ofback-side burning but was quickly extinguished. Thiswould have been a catastrophic failure in a solid rocketmotor, however the hybrid rocket motors are veryinsensitive to unbonds or delaminations from the case

wall. and are therefore much Safer. The phenolicinsulator sleeve performed very well during all fourtests, which totaled 54 seconds.

Data Analysis

Data from the experiments were analyzed toevaluate the average fuel regression rates, combustionefficiencies, oxidizer-to-fuel ratios , nozzle erosion rates,

and motor performance (Table 3). The mass flow rate

of fuel is the change in mass of the fuel grain measured

before and after the test divided by the test burn time;

thf = Am + At (4)

The mass flow rate of oxygen was determined fromthe equation,

fro* = Cd Pox Aj y

where the upstream pressure was measured duringthe static test and the temperature assumed to be at

ambient conditions. An effective average regression rate

for the fuel was determined bY;

ro = F [ ( rhox + pf ro SA )/Ap 10 (6)

Page 4

This equation assumes that all regression occursonly in the bore and does not occur at ttre exposed endsof the grain. The characteristic velocity is dehned as,

C*=JtrAgdt+ Itrt*ro*)dt e)

This was determined from the average chamberpressure, measured throat diameter, and experimentallyderived total mass flow rate

IIII = ffiox + mf (8)

Critical Design IssuesThe critical design issues observed from the seven

static tests are;

. severe nozzle throat erosion due to high temperature(>5000oD oxygen rqrction.

. case metals must be insulated from hot gasenvironment

r v&rying combustion efficiency due to changingmixture ratio

. High pressure oxygen storage vs chamber pressureand burn time

. Non-linear scale effects for small hybrid motors

. Higher inert system weights for small scale hybridmotors

. Non-uniform regression rates (a,xial & radial)

These issues pose a challenge to the propulsiondesigner. However, more work in these areas will bringmaturity to hybrid motors that will offer greater overallbenefits than either solid or liquid rocket motors.

Flight Design ConceptIn order to minimize the overall system inert

weights for a gaseous oxygen hybrid rocket motor, itwas necess ry to incorporate the high pressure oxygenvessel as part of the flight design concept To achievema:rimum performance for a high pressure vessel, leadto the selection of a graphite/epoxy filament woundpressure bottle with an aluminum liner. lJulizing a 5micron diameter filament in a 12,000 filament per tow(Amoco T-900) p're-impregnated with Fiberite LRF 545resin yielded a performance capability of 8 ,2W psi witha margin of safety greater than 3. Total bottle weightwith aluminum liner and gel coat was less than 4pounds. The T-900 fiber is an expensive aerospacegrade fiber that was selected based on it's availability atthe Productivity Enhancement Labs managed byThiokol Corporation at Marshal Space Flight Center.Because of the experience base at this faciliry using thismaterial and the low level of material required (less than5 pounds) an extremely robust design was afforded withlittle or no risk of stnrctural failure. More cost effectivematerials and realistic production designs would greatlyreduce the cost of this component. However, the

filament winding process is an automated low-costmanufacturing process that can easily support increasedproduction rates and supports the low-cost objective ofthis project.

A butane type igniter would reduce the set-up timeand allow for start-stop-res tart capability during arajectory flight. The case must be designed to allowfor easy loading of the fuel grains and the insulating ofall exposed case materials. The nozzle design should bea one piece molded nozzle with a silicone-carbide throatinsert. The nozzle would be secured by a spanner ringthreaded to the aft end of the case.

Future ActivitiesThe long term and ultimate goal for this project is

to fly a prototype hybrid rocket motor. To achieve thatgoal this project was undertaken to demonstrate theability to reliably initiate and maintain motor burnutilizing recycled thermoplastic fuel and gaseousoxygen. Having accomplished that aspect of theproject, system modifications were evaluated to allowfor flight of the hybrid motor. The key to the flightdesign would be the pressurized bottle containing thegaseous oxygen. The performance-to-weight advantagesassociated with a graphite/epoxy filament woundpressure bottle, along with the relatively compact sizeiead to it's selection for the flight concept. However,more testing is needed to gather data on PMMA,Recycled HDPE and possible alternatives. Completionand testing of the high pressure (8,200 psi) oxygenstorage bottle. Further testing is also needed to betterunderstand thermal protection within the motor. Thenext step will be to design a hybrid rocket system thatwill fit within a 5.7 5 inch diameter sounding rocket,which will be launched sometime next year.

ConclusionsThis project was successful in demonsrating that

valuable data can be obtained from "off-the-shelf"hardware and that this data can be used to develop andoptimized design of a low cost Hybrid flight motor.

Gained valuable experience in the design of hybridrocket motors. More work is needed to optimize a

design.

This project has successfully demonstrated theinherent safety of the hybrid rocket motor using gaseous

oxygen.

Demonstrated the viability of thermoplastic fuels.However did not collect enough data to validate the burnrate model for thermoplastics.

The project was successful in validating computerdesign models at extremely low costs, relative to typicalaerospace approach.

Page 5

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Nomenclature

Burn ratn exponentBurn Rate ConstantPort Area Q"2)Throat Area (in2)Injector Area (in2)Characteristic velocity (ftlsec)Discharge coefficientInitial bore diameter (in)Delta time (sec)

Change in fuel massChange in bore radiusThroat DiameterSpecific heat ratio32.17 4 (tbm - ftllb f- sec2)

Oxygen mass flux (lbm/in2lsec!Oxygen mass flow rarc (lbm/sec)

Fuel mass flow rate (lbm/sec)Total mass flow rate (lbm/sec)

Oxidizer/Fuel ratio (w0Motor chamber pressure (psi)

Pressure - Oxygen injector pressure

Nozzle throat efficiencyThrust efficiency3.14159Fuel regression ratelburnrate (ir/sec)Gas constant for oxygen

pf Fuel densiry ltUs/in3)Se Su{ace area of bore Gn2)Tox Static temperature - O*ygen injectorl, Exit divergence factorOD Outer DiameterID Inter DiameterHDPE High Density PolyethylenePMMA Polymethyl-methylacrilate (Plexiglas)R-I{DPE Recycled High Density Polyethylene

Acknowledgments

Aerojet Corporation for technical expertise and forsupport to present this paper at the 29ft JointPropulsion Conference. A special thanks to JackSohl and Phil Crimmins for their support in thisunique project.

North Alabama Ballistics Lab (NABL) and SamSchlueter, Facility Director, for use of the facilitiesat nearly anytime.

Pacific Instnrments and to Don Seyk, VP of Marketingand Doug Allen, PD Mgr., for the gracious use ofthe BAKKER 2570-P data acquisition system.

American Automated Engineering , Inc. for hightemperature insulation materials and a specialthanks to Darryl Thornton for his interest andsupport.

Sign Great Lakes Carbon Colp. for Graphite NozzleThroat Materiats GK-71 &, HLM85).

Loctite & Permatex for high Temperattue SiliconAdhesive and Filler samples.

References

1) G.P. SUTTON, ROCKET PROPUI^SION ELEMENTS

6th Edition, ISBN 0-471-52938-9, 1992Wiley & Sons, Inc.,

2) P.N. Estey and G.R. Whittinghill, HYBRIDRCf, KET MCrIOR PROPEIIA]VT S ELECTION

AUTERNATTVES

ArAA 92-3592American Rocket Company (AMROC)Camarillo, CA

3) B. Greiner and Dr. R. A. Frederick, Jr., RESULTS

OFLABSCALE HTtsRID ROCKET MOTOR

IN\CSTIGATIONAIAA 92-3307Propulsion Research CenterUniversity of Alabama in Huntsville

4) K.E. Adams and F.'W. Jordan, COMPUTER STATIC

FIRING ANALYSIS TECHMQT]ES OF SOLIDPROPELLANIT ROCKET MOTORS

Atlantic Research CorporationGainsville, VA

t C.E. Thies and F.W. Jordan, EFFECTS OF NOZZLE

CONFIGURATION ANID DEFECTS ON MOTOR

EFFICIENCYArAA 81-1379Atlantic Research CorporationGainsville, VA

o NASA SP-8076, SOUD PROPELLAIIT GRAINDESIGN AND INTERNAL BALUSTICS

lvlarch 1972National Aeronautics and Space Adminisfation

Page 6

Gox -=+

Thermoplastic Fuel

Figure 1: Pipe Chamber Gonfiguration

Thermoplastic Fuel

Steel End Capwith

3/1 6" No zzle

Glass-FilledPhenolic lnsulator

EK-71 GraPhiteNozzle lnsert

lgniterAssembly

lgniterCartridge

Gox -t

GN2 -RTV Insulation

304 StainlessSteel Chamber

Figure 2= Static Test chamber Gonfiguration

ANSI 150Flanges

2" Schedule 40Steel Pipe

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-7 .200e-01 2.325e+00

Preliminary Performance Modelfor Polyethylene Thermoplastic

Table:1

otF sF" C' (f t/sec )

0 .000.250.501 .001.502.002.503.003.505.00

10.0015.00

1 00.00

0

144lo/198243255252245239224197178

10

0

330437 644438545257 2756415495SAPA5J5b503144313981

10

Note o: Frozen Fiow, Expansion = 4, Pc = 100 psia, Sea Level

Hybrid Rocket Motor Test MatrixTable: 2

Test No. Fuel TWeightdelta Gor Valve

Burn RatePressure NozzleChamber erosion

Time (sec in/sec siq) ratePipe Chamber#1#2#3

HDPEHDPE

R.HDPE

2011.418

1 4.1151518

0.3910.0910.3g1

b̂2.51.5

0.0 2910.01550.0 057

0.0 050.00750.00950.0095

105 0.0395 0.017

#44#48#sA#sB

HDPEHDPE

HDPE

HDPE

0.091 20.170 44.177 60.264 6

100190207207

0

0

0.0020.002