AIAA 97-2802 DEVELOPMENT WORK ON A SMALL

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AIAA 97-2802DEVELOPMENT WORK ON A SMALLEXPERIMENTAL HYBRID ROCKETMatthias GrosseD-83435 Bad Reichenhall33rd AIAA/ASME/SAE/ASEEJoint Propulsion Conference & ExhibitJuly 6 - 9, 1997 / Seattle,WAFor permission to copy or republish, contact the American Institute of Aeronautics and Astronautics1801 Alexander Bell Drive, Suite 500, Reston, VA 22091

2DEVELOPMENT WORK ON A SMALL EXPERIMENTAL HYBRID ROCKETMatthias Grosse*Bad Reichenhall, GermanyAbstractIntroductionThis paper presents development work on a smallunguided, experimental hybrid rocket with an initial massof 65 kg (143 lbs). Hydroxyl-terminated polyisopreneand pressure-fed red fuming nitric acid are used aspropellants. The HERA (Hybride Experimental-Rakete)experimental rocket project was started in order to studycharacteristics of the above mentioned hybrid propellants.The final goal of this investigation was to develop arocket with a range of 20 km. More than 40 test firingswere performed with different experimental testbeds.These tests were divided into five different series: Initial tests Internal ballistic tests to evaluate performance andfuel regression rates Thrust measurements and hardware tests Testing of a robust stationary testbed of the HERAhybrid motor Qualification of the entire propulsion systemBesides the established liquid and solid propulsionsystems, the hybrid rocket is very interesting for futureapplication. The combination of solid and liquid fuelcomponents promises advantages and improvementsconcerning reliability, safety, cost effectiveness and highperformance combined with non-polluting characteristics.NomenclatureaAeAtc*CFGoxIsL*lG mfupcrr0r1dr/dttbαβρΦηc*ηCFηIsregression law constantnozzle exit areanozzle throat areacharacteristic velocitythrust coefficientoxidizer mass fluxspecific impulsecharacteristic motor lengthgrain lengthfuel mass consumptionchamber pressureport radiusinitial port radiusend port radiusfuel regression rateburning timemass flux exponentgeometrie exponentfuel densitymixing ratio oxidizer/fuelcombustion efficiencynozzle efficiencythrust efficiencyProject HERA aimed at the devolepment of a smallexperimental hybrid rocket in a realisistic environment.This procedure has the advantage, that the experimentalprogram was influenced by the development parametersof a flight vehicle. The investigation was conducted andfinanced between 1989-1992 by the author [1] and wassupported by DASA (Bremen/Trauen). Other aspects ofthe project HERA were the study of specific hybridinternal ballistics, performance characteristics, fuelregression parameters and measurement accuracy.The design parameters of HERA were a range of 20 kmand a lift-off mass of less than 70 kg. Further designfeatures for the experimental rocket were: Storable oxidizerUncooled expansion nozzleSimple ignition methodPressure regulated oxidizer feed unitThe HERA rocket consists of a hybrid rocket motor,stabilizing fins and a parachute system. Constructionmaterials were mainly alumium alloys and carboncomposites. The motor nozzle insert is made of graphiteand the nozzle structure of nickel-based super-alloy. Theoxidizer feed unit consists of the components: oxidizertank, oxidizer flow controller and regulatedpressurization system.*Member, AIAA, Development EngineerCopyright 1997 by the American Institute ofAeronautics and Astronautics.All rights reserved.Figure 1: Initial Version of HERA

3The initial rocket design (fig. 1), based on theoreticalcalculations and initial tests, was continually improved inan iterative process, using the test firing data and resultsof 6-DOF-simulations. These simulations producedlaunch parameters for maximum range and the resultingimpact dispersion, as well as, wind weighting functionsand structural loads on the vehicle. Aerodynamic datawere obtained with the aid of a semi-empirical computercode. The final rocket design (fig. 2 and tab. 1) shouldachieve a maximum range of 19 km and a burn-outvelocity of Mach 2.2. Final testing could not be carriedout, due to budgetary problems.The initial version of the HERA motor consisted of fourchambers with cylindrical grains mounted to a singleexpansion nozzle. Figure 3 shows the nozzle sectionmade of nickel-based super alloys and the graphite insert.Table 1. Characteristics of Final HERA VersionLaunch massOxidizer massFuel massLaunch thrustAverage chamber pressureBurning timeCaliberLength65.4 kg26.4 kg7.0 kg3250 N32 bar21.6 s0.195 m2.90 mFor the development of the hybrid propulsion five testseries were planned to investigate important aspects ofthe hybrid rocket design (tab. 2 and 5).Table 2. Test SeriesTest SeriesParachutePressurisationSystem Tests of Motor Components Combustion Efficiency Fuel Regression Parameters Ignition Behavior Determination of BestOperating Conditions Complete Fuel Utilization Reproducability Nozzle Material andPerformance Ignition BehaviorTestbedFirings Combustion Stability Qualification of Oxidizer FeedUnitQualificationFirings Test of the Entire PropulsionSystemInitial FiringsOxidizerTankInternalBallistic TestsMain Valve andIgnition DeviceObjectivesHybrid MotorFinsThrust TestsFigure 2: Final Configuration of HERAFigure 3: Nozzle Assembly of RocketAmerican Institute of Aeronautics and Astronautics

4Hybrid Propellant Combination Hydroxylterminated Polyisoprene /RFNAThe selected solid fuel hydroxyl-terminated polyisoprene(KURARAY LIR-503) is a common industrial product.Diphenylmethandiisocyanate (MDI) was selected as thecurative due to its favorable toxic characteristics. Thefuel was cast into paper phenolic cartridges. Properties ofthe fuel are listed in table 3. The selected oxidizer wasRFNA (83% HNO3, 16 % N2O4 and 1 % H2O). Thestochiometric mixture ratio with hydroxyl-terminatedpolyisoprene is 5.107. The maximum theoreticalcharcteristic velocity is 1540 m/s with a O/F-mixtureratio of 4.3.Use of mixing diaphragm and special mixing chamberwas given up, due to it complexity. The following motorcomponents and materials were selected for the rocket: Hollow cone spray injectors High-density, fine-grade graphite nozzle inserts Fuel grain design using cylindrical ports and aninitial L/D of 18 and initial port radius of 17.5 mmTable 3. Properties of Hydroxyl-terminatedPolyisoprene (LIR-503)Reduced Sum Formula(cured state)Prepolymer:Molecular WeightFunktional GroupNumber of OH-groupsper MoleculeJod NumberViscosity (38 C)Density (20 C)CH1,586N0,0022O0,003525000OH2.5368800 poise0.92 g/cm3The propellants hydroxyl-terminated polyisoprene andRFNA are not hypergolic, therefore all test firings used amixture of 3 parts Fe-acetonylacetate and 2 parts 2,6dimethylaniline as igniter. 15-20 g of the ignition mixturewas melted and applied to the grain surface, where itsolified very quickly (fig. 4).Ignition MixtureFuel GrainFigure 4: Application of Ignition MixtureTest Firings and SetupsInitial TestsTwo different motors were built for the testing of motorcomponents, i.e. injectors, nozzle configurations andmaterials. The first motor used a regenerative cooled steelnozzle coated with zirconium(IV)oxide and a solid conespray injector (fig. 5). The second motor used anuncooled nozzle structure made of Haynes alloy with aZrO2-coated steel nozzle insert. A hollow cone sprayinjector was used. Both motors used a mixing chamberlined with graphite and a mixing diaphragm. The motorcase was made of aluminum. Nine firing tests wereperformed.Figure 5: Initial Test Motor 1Internal Ballistic TestsThe internal ballistic test serie was used to determine thefuel regression rate as a function of oxidizer mass flux,chamber pressure and grain port radius. Additionally, thecharacteristic velocity and combustion efficiency wereobtained. The data were employed to select the operatingconditions of the HERA motor. A simple test motor (fig.6) was used for evaluating the internal ballistic characteristics of hydroxyl-terminated polyisoprene/RFNA. Thetest conditions span an interval of oxidizer mass fluxfrom 4-15 g/s*cm², oxidizer mass flow from 95-210 g/s,final port radius ratio l.1-1.8 and chamber pressure from8 –18 bar. More than 20 tests were performed. Theyexhibited very smooth and stable combustion behavior.Figure 6: Internal Ballistic Test MotorFuel G rainH ollow C one S pray InjectorAmerican Institute of Aeronautics and AstronauticsN ozzle Insert

5Thrust TestsIn this series, the grain was fired for the full durationunder HERA operating conditions. The thrust, specificimpulse, thrust coefficient and efficiency weredetermined. Furthermore, these tests produced fuelregression rate data. Nine test firings were performed (fig.7). A modified version of motor 2 from the initial testingwas utilized (fig. 8).Figure10: Testbed FiringQualification FiringsThe objective of this testing was the qualification of thepropulsion system for flight testing (tab.4). The test standwas fully instrumented for monitoring of the all firingdata (fig. 11).Figure 7: Thrust TestH o llo w C o n e S pra y In je ctorFu el G ra inG ra ph it N ozzle In se rtH a yn es A llo y N o zzle S tructu reFigure 8: Configuration of Thrust Test MotorTestbed FiringsThe testbed motor was a simplified version of the HERAmotor with a maximal burning time of 6 s. The nozzleentrance volume was simulated by an additional chamber(fig. 9). Nozzle and chamber assembly were of mild steel.A 2 mm thick fuel layer was sealed to the chamber wall.The investigation of ignition behavior, combustionstability and performance of the testbed with the rocketoxidizer feed unit was performed with this version. Fourtest firings were performed (fig. 10).M an ifoldFu el L ayerM ixin g C ha m be rH o llo w C o ne S p ra y Injecto rN o zzleFigure 9: Configuration of Testbed MotorFigure 11: Preparation of Qualification TestTable 4: Qualification Parameters ofInitial HERA MotorChamber PressureTank PressureThrustSpecific ImpulseCombustion EfficiencyAverage Mixture RatioBurning TimeAmerican Institute of Aeronautics and Astronautics17.0 bar24.0 bar1780 N2020 Ns/kg0.934.037.5 s

6Table 5: Test Configuration MatrixMotorSprayInjectorNozzle Throat Nozzle InsertArea (cm²)MaterialMixingChamberL* (m)NominalFuel GrainBurning Time Dimensions(s)r1 r0 lG (cm)NominalThrust(N)Initial TestFull Cone1.8 – 2.5Steel ZrO20.550different400InternalBallistic TestHollow Cone*1.77Nimonic AlloyNone323.2*1.75*62400Thrust TestHollow ollow Cone*7.55-8.16Mild Steel0.263.2*1.75*721700Initial HERAHollow Cone*7.55Graphite0.337.53.4*1.75*721780* average diameter of droplets: 0.4 mm by an injector pressure difference of 10 barDiscussion of ResultsSelection of Operating ConditionsChamber Pressure (bar)The optimum chamber pressure and oxidizer mass fluxfor the investigated motor configuration given throughthe fuel grain geometry, nozzle throat area and injectorparameter were examined. The injected oxidizer massflow is plotted against chamber pressure for the internalballistic and thrust test series in figure 12 (averagedvalues). A nearly linear correlation up to a mass flow of170 g/s was observed. There was no increase in chamberpressure above 170 g/s. The fuel mass flow was lowerthan expected. The mixture ratio becomes oxidizer richand c* declines. The average oxidizer mass flux above170 g/s was greater than 11.5 g/s*cm².1510IgnitionA simple ignition method was used as mentioned above.The grain surface was coated with an hypergolic ignitionmixture. It proved to be inadequate in achieving a reliableand predictable ignition and caused several test failures.The mixture was applied only to the first few centimetersof the grain port length for the initial and internalregression test series. There was no pressure overshootobserved, but several tests quenched. Anotherundesirable effect of this ignition method was anincreased fuel regression due to massive thermal ignitionpeak during the first second of burn (fig. 13). This effectcontributed to the regression data errors.Internal Ballistic andThrust Tests4.7 - 10.7 g/s*cm²11.5 - 15.4 g/s*cm²Oxidizer Mass Flux5004080120160200Oxidizer Mass Flow (g/s)Figure 12: Chamber Pressure versus Oxidizer MassFlow (single port grain)Therefore the average oxidizer mass flux of the rocketwas set to 9 g/s*cm². The expected average chamberpressure for a single port configuration without mixingdevices for 170 g/s is 16.5 bar. Firings with the HERAmotor testbed confirmed this expected value.Figure 13: Chamber Pressure versus Time Test(Internal Ballistic Test)For further testing the ignition mixture was applied tonearly half the length of the port to avoid quenching.Correct ignition was achieved, but ignition pressurevaried widely. Extreme ignition pressure peaks destroyedseveral test motors. An ignition pressure peak higher than100 bar caused the destruction of the HERA motor in thequalification test series.Combustion StabilityThe initial and internal ballistic tests displayed verysmooth combustion (fig.14 a 15).American Institute of Aeronautics and Astronautics

7Fp (b a r)c(N )T h r u s t3 0 03 0Chamber Pressure No.2 (bar)Ignition Peak (80 bar)12840C h a m b e r P r e s s u r e1/ 21 5 05 ,0T im e1 0 ,0(s )Figure 14: Parameter of a Initial Test versus Time56789Time (s)1 50 ,04Figure 16: Chamber Pressure versus Time TestCharacteristic Velocity and CombustionEfficiencyThe combustion efficiency defined as ηc* c*ex / c*thcharacterize the completeness of combustion. Thetheoretical characteristic velocity and the experimentalresults of the internal ballistic and thrust test series areshown in figure 17. The theoretical performance wascalculated with computer code [2], which is based on themethod of Bathelt and Volk [3]. The ignition mixturevalues were included in the calculations. A quadraticpolynom was fitted to the experimental results ( c*ex 986.35 194.1Φ-23.0Φ²). The average combustionefficiency was 0.91. This value was calculated using themean values of the polynom and the theoretical curve.The maximum c*ex of 1400 m/s was at a mixture ratio of4.3. The average ηc* of the testbed firings was 0.92.The chamber pressure at the nozzle entrance was used forcalculation purposes. A pressure drop between injectionhead and nozzle entrance for this grain configuration(without mixing diaphragm) was typicaly 0.3 bar. In theinitial test firings (with mixing diaphragm and chamber) avalue of 1 bar was measured.The first thrust test firings showed some roughcombustion, believed to have been caused by themounting of the complete setup on flexible thrustmeasurement facility. Modifications of the eliminated therough combustion. The testbed firings showed anunstable operation with very high amplitude andoscillations at 8 Hz. (fig. 16).1600Characteristic Velocity (m/s)Figure 15: Typical Smooth Internal Ballistic Test140012001000Theoretical800Internal Ballisticand Thrust TestsTheoretical Performance:600Chamber Pressure 15 bar4002000246Mixture Ratio (-)Figure 17: Experimental & TheoreticalCharacteristic Velocity versus Mixture RatioAmerican Institute of Aeronautics and Astronautics8

8Specific Impulse and Thrust CoefficientThe experimental specific impulse results were fitted bya quadratic polynom. The specific impulse efficiency,defined as ηIs Is ex / Is th , was 0.90. A specific impulse of1910 Ns/kg was achieved with a mixture ratio of 4.3. Thenozzle or thrust coefficient effectiveness, defined as ηCF ηIs /ηc* ,was 0.99. The theoretical calculations, basedon a frozen equilibrium flow assumption, probablyunder-predicted the nozzle performance (Ae/At 4.49),which would lead to this high value. The thrustcoefficients are shown in figure 18.as function of the oxidizer mass flux above and belowthis value are given in figure 19 and table 6 (port radiusas parameter ). The value of the empirical oxidizer massflux exponent (0.6 - 0.9) below about 9 g/s*cm² is wellknown from other authors. Above 9 g/s*cm² the exponentis unusually small. The correlation results display asurprisingly strong dependency on the port radius (tab.6).Table 6:Correlation Parameter and Relative Errors1,61,21,00,80,60,4Thrust TestsTheoretical Coefficients for15 bar Chamber Pressure,Frozen Equilibrium,Sea LevelTheoreticalAe/At 4.49Exp. Ae/At ror0.0057 135 %0.0078 41%0.8 22 %0.254 56 %0.81 50 %1.10 28 %0,58Mixture Ratio (-)Figure 18: Experimental & Theoretical ThrustCoefficient versus Mixture RatioFuel Regression Rate (mm/s)Thrust Coefficient (-)1,4Parameter0,4r 26 mm24 mm0,3alpha 0.254 beta 1.120 mmalpha 0.8 beta 0.810,24812Oxidizer Mass Flux (g/s*cm²)Fuel RegressionThe overall average regression rate is based on theassumption that the surface of the fuel grain regressescylindrically throughout the firing.It is defined as:dr/dt ( mfu/πρlG r0²)0.5 – r0))/tb .A data analysis program [4] was used to determine thefuel regression rate as a function of the oxidizer mass,chamber pressure and grain port radius. The replacementof the time dependent motor parameters by their timeaveraged values for this grain configuration results onlyin negligible errors in the regression correlationparameters [1], [5], since the final to initial grain portradius ratio was small only 1.8. The fuel regressionresults gave the best correlation with oxidizer mass fluxand port radius as parameters:dr/dt a Goxαrβ .The correlation constants a, α and β are subject torelatively high errors (tab. 6). This results from thelimited accuracy of measurements of the parametersinvolved. The ignition mixture also falsified the averageregression rate (fig. 13). The small size of the test motorsalso leads to additional errors in the measurements [6].Figure 19: Fuel Regression Rate versus OxidizerMass FluxAnother approach to evaluate regression behavior is toanalyse tests over a small range of average oxidizer massflow, chamber pressure and with a constant initial grainport radius. This procedure simulates a one time firingtest. The average oxidizer mass flux was above 9 g/s*cm²for these tests. A plot of the final port radius versusburning time for these tests shows a constant fuelregression rate (fig. 20 ).The increasing fuel mass flow resulted in a decreasingO/F mixture ratio (fig. 21). This is confirmed byobservation of the plume color of long burning tests,which change from transparent to brilliant yellow and aslightly increased chamber pressure (max. 10%). Theseresults support the statement that above 8-10 g/s*cm²the port radius is the determining factor in the fuelregression rate.It was observed that a significant change in fuelregression behavior accured at an oxidizer mass flux of 810 g/s*cm². Therefore, two regression correlation valuesAmerican Institute of Aeronautics and Astronautics

9Final Port Radius Ratio (-)2,0It is interesting to note, that the scatter of the data pointsin the graphs (c* 2.5 – 3.3 %, Is 4.6 % and dr/dt 3.1– 7.9% ) are on the same order as the estimated errorscalculated by the method of error propagation.Oxidizer Mass Flow 143 - 185 g/sChamber pressure 13.3 -17.2 barInitial Port Radius17.5 mm1,81,6Especially the parameters oxidizer mass flux and fuelregression rate, show significant errors, due toconceptional uncertainty in port diameter and active fuelregression time. Furthermore, the length dependent fuelregression gave regression rate errors. (fig. 22).1,41,21,0051015202530Injection Point35Time (s)Figure 20: Final Port Radius Ratio versus TimeDisturbedZoneHomogeneousZone76.Figure 22: Port Length Dependent RegressionMixture Ratio (-)543211,0Table 8: Estimated Errors of Derived Motor ParametersOxidizer Mass Flow 143 -185 g/sChamber Pressure 13.3 –17.2 barInitial Port Radius17.5 mm1,21,41,61,82,0Final Port Radius Ratio (-)Figure 21: Mixture Ratio versus Final Port RadiusRatioParameterValue2σσ-Error (%)Oxidizer FluxFuel Mass FlowTotal Mass FlowMixture Ratio4-15 g/s*cm²20- 40 g/s110 – 250 g/s3.5 –7.06.0 – 8.00.5 – 3.01.3 – 2.71.5 – 4.0CharacteristicVelocityFuel Regression Rate1400 m/s2.0 – 4.60.25 – 0.45 m

Internal Ballistic Tests The internal ballistic test serie was used to determine the fuel regression rate as a function of oxidizer mass flux, chamber pressure and grain port radius. Additionally, the characteristic velocity and combustion efficiency were obtained. The data were employed to s

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