N2O/O2 blends safe and volumetrically efficient oxidizers

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NomenclatureAc = fuel port cross-sectional area, cm2

Aexit = nozzle exit area, cm2

As = cross sectional area at shock wave position, cm2

A1 = venturi inlet area, cm2

A2 = venturi outlet area, cm2

A* =crosssectionalareaatwhichlocalflowchokes,cm2

Aexit/A* = nozzle expansion-ratio

CF =thrustcoefficient

Cp =specificheatatconstantpressure,J/kg-K

Cv =specificheatatconstantvolume,J/kg-K

Ce = effective exhaust velocity, m/s

c* = characteristic velocity of propellants, m/s

F = thrust, N

oxG =oxidizermassflux,g/cm2-s

totalG =totalmassflux,g/cm2-s

g0 =nominalaccelerationofgravityatsealevel,9.8067 m/s

HVPS=highvoltagepowersupply

Isp =specificimpulse,s

Lport =fuelgrainlength,cm

fM =moleorvolumefractionofonespeciesinabinarygasmixture

Mexit =exitplaneMachnumber

Mw =molecularweight,g/g-mol

ΔMfuel=consumedfuelmass,g

ΔMox =consumedoxidizermass,g

fuelm =fuelmassflow,g/s

oxm =oxidizermassflow,g/s

totalm =totalmassflowthroughthenozzle,g/s

O/F = oxidizer/fuel ratio

O/Factual = actual oxidizer-to-fuel ratio

O/Fstoich =stoichiometricoxidizer-to-fuelratio

Nytrox 87 = nitrous oxide, gaseous oxygen solutionwith 87%nitrous oxide in liquid phase

P1 = venturi inlet pressure, psia

P2 = venturi throat pressure, psia

P0 =chamberpressureorstagnationpressure,psia

pexit = exit plane static pressure, psia

p∞ =ambientpressure,psia

Rg =gasconstant,J/kg-K

Ru =universalgasconstant,8314.4612J/kg-K

rL =longitudinalaverageofthefuelportradius,cm

r0 = initial fuel port radius, cm

Lr =longitudinalmeanoffuelregressionrate,cm/s

Aeron Aero Open Access J. 2019;3(4):171‒196. 171©2019 Whitmore et al. This is an open access article distributed under the terms of the Creative Commons Attribution License, which permits unrestricted use, distribution, and build upon your work non-commercially.

N2O/O2 blends safe and volumetrically efficient oxidizers for small spacecraft hybrid propulsion

Volume 3 Issue 4 - 2019

Stephen A Whitmore,1 Robb L Stoddard2

1Professor, Mechanical and Aerospace Engineering Department, Utah State University, USA2Graduate Research Associate, Mechanical and Aerospace Engineering Department, Utah State University, USA

Correspondence: Stephen A Whitmore, Professor, Mechanical and Aerospace Engineering Department, Utah State University, Logan, Utah 84321, USA, Tel +01-435-797-2951, Email

Received: October 31, 2019 | Published: November 20, 2019

Abstract

Amedicalgradenitrousoxideandgaseousoxygenfluidblend,“Nytrox,”isinvestigatedas significantly safer, but superior performing, alternative for the current generation ofenvironmentallyunsustainable spacecraft propellants. In amannerdirectly analogous tothecreationofsoda-waterusingdissolvedcarbondioxide,Nytroxiscreatedbybubblinggaseousoxygenunderhighpressureintonitrousoxideuntilthesolutionreachessaturationlevel.Oxygenintheullagevapordilutesthenitrousoxidevaporandincreasestherequireddecompositionactivationenergyofthefluidbyseveralordersofmagnitude.Consequently,any risk of inadvertent thermal or catalytic decomposition is virtually eliminated. Thispaper reports on a preliminary test-and-evaluation campaign where an existing smallspacecraftthrusterisfirsttestedusinggaseousoxygenand3-DprintedABSaspropellantsasabaseline.Thebaseline testswere then repeatedusinganoptimizedNytroxbendas“drop-in” replacement for gaseous oxygen. Test parameters compared include; ignitionreliabilityandrequiredenergy,thrustcoefficient,characteristicvelocity,specificimpulse,andfuelregressionrate.Nytroxisshowntoworkeffectivelyandreliablyasareplacementforgaseousoxygen,exhibitingaslightlyreducedspecificimpulseandregressionrate,butwith significantlyhighervolumetricefficiency.Recommended topics for future researcharediscussed.

Keywords: hybridrocket,“Green”propellants,nitrousoxidedecomposition,energyofactivation,3-Dprinting,Nytrox

Aeronautics and Aerospace Open Access Journal

Research Article Open Access

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N2O/O2 blends safe and volumetrically efficient oxidizers for small spacecraft hybrid propulsion 172Copyright:

©2019 Whitmore et al.

Citation: Whitmore SA, Stoddard RL. N2O/O2 blends safe and volumetrically efficient oxidizers for small spacecraft hybrid propulsion. Aeron Aero Open Access J. 2019;3(4):171‒196. DOI: 10.15406/aaoaj.2019.03.00097

r =meanregressionrateoverdurationoftheburn,cm/s

sg =specificgravity

T =venturiflowpathtemperature,K

T0 =stagnationtemperature,K

tburn =burntime,s

t =generictimesymbol,s

η* =combustionefficiency

γ =ratioofspecificheats,Cp/Cv

Φ = equivalence ratio

ρfuel = solid fuel density, g/cm3

ρ*Isp=densityspecificimpulse,N-s/liter

θexit =conicalnozzleexitangle,deg.

IntroductionThe Propulsion Research Laboratory (PRL) at Utah State

University(USU)recentlydevelopedapromising“green”propulsionalternative that has the potential to replace hydrazine for multipleapplications.Theunique

hybrid propulsion technology derives from the novel electricalbreakdown properties of 3-D printed acrylonitrile butadienestyrene (ABS),1,2 discovered serendipitouslywhile investigating thethermodynamic performance ofABS as a hybrid rocket fuel.3Thisconcept has been developed into a power-efficient system that canbe started and restarted with a high degree of reliability.Multipleprototypeground-testunitswiththrustlevelsvaryingfrom4.5Nto900Nhavebeendevelopedand tested.4,5Recently,onMarch25th,2018 a flight experiment containing a medium-weight prototypeof this thruster system was launched aboard a two-stage Terrier-ImprovedMalemute sounding rocket fromWallops Flight Facility(WFF).Thelaunchachievedapogeeof172km,allowingmorethan6minutesinatruespaceenvironmentabovetheVon-Karmanline.DuringthemissiontheUSUthrusterwassuccessfullyfired5timesinahardvacuumenvironment.ThepayloadsectionwassuccessfullyrecoveredbyWFFflightsupport.

Low resolution telemetry data was successfully downlinkedand delivered to USU for analysis. Whitmore & Bulcher6 report thedetailsof thisflight testexperiment.Thus,with this spaceflightdemonstration the technology readiness level (TRL) of the arcignitiontechnologymustbeacknowledgedtobeat least level5.Initsmostmatureformthissystemusesgaseousoxygen(GOX)astheoxidizerwith3-DprintedABSasthefuel.TheGOX/ABSpropellantsresulted in a highlymass efficient system,with a flightweight 25N thruster systemachievingvacuumIsp greater than300 seconds.

7,8 Unfortunately,unlessstoredasveryhighpressures,GOXhasalowspecificgravityandisavolumetricallyinefficientpropellant.Ahigherdensityalternativeishighlydesirable.ThispaperwillinvestigatethepotentialtousemedicalgradeNitrousOxide(N2O),gaseousoxygen(GOX)blends,typicallyusedforanesthesiaoranalgesicapplications,as “drop-in” replacement for GOX in a legacy small spacecraftthruster system, previously developed and tested at theUtah StatePropulsion Research Laboratory.As will be described later in thispaper, thesemixturesareofferedasanalternative topureN2O due

toanorderofmagnitudeincreaseinsafetyandhandling,withonlyaminordecreaseinoverallperformance,andasignificantincreaseinoverallvolumetricefficiency.

Background on available green propellant options

A recent study9,10bytheEuropeanSpaceAgencySpaceResearchandTechnologyCenter(ESTEC)hasidentifiedtwoessentialdesignelements to achieving low cost space access and operations; 1)Reduced production, operational, and transport costs due to lowerpropellant toxicityandexplosionhazards,and2)Reducedcostsduetoanoverallreductioninsubsystemscomplexityandoverallsystemsinterface complexity. The ESA/ESTEC study showed the potentialfor considerable operational cost savings by simplifying propellantgroundhandlingprocedures.Developinganon-toxic,stable“green”alternative formost commonly used toxic or potentially-hazardouspropellantswashighlyrecommendedbytheESTECstudy.

Ionic liquid propellants

InresponsetoESA/ESTECreportandother“becominggreen”11 recommendations,forthepastdecadetheUSAirForce(USAF)andtheSwedishSpaceCorporation(SSC)subsidiaryEcological Advanced Propulsion Systems (ECAPS) have been pursuing less hazardousalternatives to hydrazine. The two most highly-developed “green-propellant”alternativesarebasedonaqueoussolutionsof the ionic liquids(ILs)AmmoniumDinitramide(ADN)12,13andHydroxylamineNitrate (HAN).14,15 InAugust 2011, ECAPS announced the resultsof a year-long series of in-space tests of a 1-N thruster comparingtheir High Performance Green Propellant (HPGP) to hydrazine onthePrismaspacecraftplatform.16ECAPSclaimsthatHPGPdeliveredequivalent–tosuperiorperformance,withaspecificimpulseabove230seconds.NASA recently selected theUSAF-developedHAN-basedpropellant AF-M315E for its “Green Propellant Infusion Mission(GPIM).17 In spite of being called “green,” by theirmanufacturers,bothoftheabove-mentionedIL-propellantsaretoxictoorganictissue,and special handlingprecautions are required. IL-basedpropellantsare not truly “green.”Thus, theUSAF has recently begun to refertosuchIL-formulationsmoreproperlyashaving“reducedtoxicity”properties.Inadditionaltopotentialtoxicity,thereexistseveralkeydevelopmentalissuesassociatedwithILbasedpropellantsthatmakethemunsuitableforsmallspacecraftapplications.

ThehighwatercontentmakesIL-propellantsnotoriouslyhardtoignite.MultiplecatalystsystemshavebeendevelopedtoaugmentILignitability; however, room temperature ignition does not currentlyexist. Catalyst beds must be preheated from 350-400ºC beforeand during ignition, and this preheat can consume up to 15,000joules of energy.Catalyst beds and associated heating systems addsignificantlytotheinertmassofthespacecraftandthehigh-wattagepreheat requirementpresentsa significantdisadvantage for systemswherepowerbudgetsarelimited.Duetoveryslowreactionkinetics8 atmoderate pressures (2000-3000 kPa) system latencies associatedwith IL-based propellants are significant for moderate chamberpressuresandmaylimittheusefulnessofIL-propellantsforspacecraftmaneuveringandcontrolsystems.18Afinalmajordrawbackofthesepropellantsisthehighcombustiontemperaturethatsignificantlylimitsthe systemburn lifetimes.Both theLMPandAFMpropellants arepronetouncontrolledenergeticeventsthatresultinhightemperaturechemical fires. In spite of this volatility, the performance of ionic

https://doi.org/10.15406/aaoaj.2019.03.00097




liquidpropellantsisgenerallyquitelow-withvacuumspecificimpulsevaluesnomorethan230secondswhencomparedtotraditionalsolidand or bi-propellant options.The combination of these detrimentalcharacteristicscomparedtotheir“green”advantageshasledsomeintheindustrytoquestionwhetherIonicliquidsaspropellantshavebeen“oversold.”9

Hybrid rockets as a “green” propulsion alternative

The inherent safety and environmental friendliness of hybridrocketsystemshavebeenknownforseveraldecades.19Hybridshavethe potential to act as an ideal “green” alternative formany of thecurrentgenerationoftoxicorhazardouspropellants.Becausehybridsystems only require a single fluid flow path, they are of similarcomplexitytomonopropellantsystems;butwithsignificantlyhigherperformance.Infact,whenproperlyoptimized,hybridsystemshavethepotentialtoprovidethesameperformanceassignificantlymorecomplex bi-propellant liquid systems.While hybrid rocket systemshave been considered for applications ranging from large launchsystemstonanosatellites, theyhavenotfoundarealnichewith thespace-launch and space propulsion industries. Solid and liquid bi-propellant systems have been under development for more thanseven decades and the state of technology development for hybridsystemsisratherimmaturebycomparison.AlowTRListheprimarydisadvantageofhybridsystems.

Thisscenarioisanalogoustothestateofsolarelectricpropulsion(SEP) three decades 153 ago. Systems based on SEPwere alwaysconsidered a higher-risk solution when compared to conventionalsystems. This situation changed when conventional systems nolonger met commercial and/or science-driven requirements, andthistechnologygapdrovetherapidTRLdevelopmentSEPsystems.Technology maturation resulted in the extensive use of arcjetthrustersforGEOcommunicationssatellites,eventuallyresultingintheNSTAR ion systems20 that set in-space ΔV records onNASA’sDeepSpace121andDawnmissions.22Insimilarfashion,astheTRLmatures, small hybrid systems offer the potential to fill an unmetand growing need for advanced propulsion both in-space and aslaunch stages for the emerging SmallSAT market. Hybrid rocketsofferparticularutilityfortheupperstagesofanano-launchvehicle.Althoughahybridrocketwill increase theoverallsystemdrymasscomparedtoasolid-propellantmotor,thecapabilitiestothrottle,shutdownondemand, coast, and relight themotor,will offset any lossin performance of the stage. Such a “smart stage”would not onlyprovide ΔVtoenablethepayloadtoreachorbit;butcanalsoserveasanonorbitmaneuveringsystemthatallowspreciseplacementofthepayload.Suchasystemcouldalsoprovideextensivecapabilitiesforendo-atmosphericmaneuveringforavarietyofdefenseapplications.

Hybrid rocket low-power arc-ignition system

Historically,duetothelackofareliablenon-pyrotechnic,multiple-use ignition method, hybrid rockets have never been seriouslyconsidered as feasible for in-space propulsion. Hybrid rockets are“safe”duetotherelativepropellantstability;however,thisstabilitymakes hybrid rocket systems notoriously difficult to ignite. Thehybridrocketignitionsourcemustprovidesufficientheattopyrolizethesolidfuelgrainattheheadendofthemotor,whilesimultaneouslyproviding sufficient residual energy to overcome the activationenergy of the propellants. Conventional solid-propellant ignitionsystems use pyrotechnic or “squib” charges to ignite a secondarysolid-propellant motor whose high-enthalpy output rate initiates

the full motor combustion. Such high-energy devices often comewith a suite of environmental and objective risks, and operationalchallenges. Pyrotechnic charges are extremely susceptible to theHazards of Electromagnetic Radiation to Ordnance (HERO),23 and largepyrotechnicchargespresentasignificantexplosionhazardthatisincompatiblewithmanylaunchopportunities.Mostimportantly,fornearlyallapplicationspyrotechnicignitorsaredesignedas“one-shot”devicesthatdonotallowamultiplerestartcapability.Thusthegreatpotentialforrestartableupperstagesorin-spacemaneuveringsystemsusing hybrid propulsion remains largely unrealized.An operationalhybridsystemwithmultiplerestartcapabilitydoesnotcurrentlyexist.

This restartability issue has been overcome by leveraging theunique electrical breakdown properties of certain 3-D printedthermoplastics like acrylonitrile butadiene styrene (ABS).23 Theauthorsdiscoveredthatfuseddepositionmodeling(FDM)processedABS possesses unique electrical breakdown properties that canbe exploited to allow for rapid on-demand ignition. Under normalconditionsABSpossessesaveryhighelectricalresistivityandisnotconsideredtobeanelectricalconductor;however,asFDM-processedABSissubjectedtoamoderateelectrostaticpotentialfieldthelayeredmaterialstructureconcentratesminuteelectricalchargesthatproducelocalized arcing between material layers. Joule heating from theresultingarcproducesasmallbuthighly-conductivemeltlayer.Thismeltlayerallowsforverystrongsurfacearcingtooccuratmoderateinput voltage levels-between 200 and 300 Volts. Additional Jouleheatingfromthestrongsurfacearcingcausesasufficientfuelmaterialtobevaporizedandseedscombustionwhensimultaneouslycombinedwithanoxidizingflow.

Figure 1a shows a typical pyrolysis event, where the ablatedhydrocarbonvaporresultsfrominductivearccarvingapathacrossthefuelmaterial.Figure1bshowsatypical3-Dmotorheadendlayoutwithflowimpingementshelvesandembeddedelectrodes.Figure1cshowstypicalignitionsystemelectronicsschematic.

Shortly after this discovery, the authors of this paper madeseveral unsuccessful attempts to reproduce a similar phenomenonwith other hybrid fuel materials including Hydroxyl-TerminatedPolybutadiene (HTPB), acrylic, paraffin, and extrudedABS.Theseexperiments also demonstrated that electrical breakdown of FDM-processed ABS occurs at voltages significantly lower than occurwithamonolithicallyfabricated(machinedorextruded)article.Theextruded or machinedABS does not electrically break down (arc)untilvoltagelevelsexceeding2000voltsareinputacrossthematerial.ThisvalueisroughlyanorderofmagnitudehigherthanforasimilarFDM-processed test article. The observed arcing properties areartifactsofboththegraincompositestructureduetoFDMfabricationand the electromechanical properties ofABS. The associated arc-ignitionsystem,developedby theUtahStateUniversityPropulsionResearchLaboratoryoverthepastfiveyears,hasresultedinapower-efficient ignition system that can be started, stopped, and restartedwithahighdegreeofreliability.1AkeypropertyoftheFDM-printedfuelisaninternalstructurethatenablesverylow(>3J),rapidonset(>100microsecond), non-catalyzed ignition. The system is entirelyimpervious to the previously-described hazards associated withelectromagneticradiation(HERO).Thelowenergyignitionprocessrequiresapre-programmedsequenceofeventsthatmakesinadvertentignitionavirtualimpossibility.1Whitmore SA, et al. Restartable Ignition Devices, Systems, and Methods Thereof. United States Patent Publication, Pub. No. US 2015/0322892A1;2015.





a) Arc-Pyrolys of Fuel b) Motor Head-end Ignitor Layout c) Tyical Arc-Ignition System

Figure 1 3-D printed hybrid arc-ignition system details.

Choice of non-toxic oxidizers for hybrid 223 space propulsion

Figure2showstheavailableoptionsforhybridrocketoxidizers.Inthepracticalrealmonly4oftheseoptionscanbeconsideredasbeingevenreasonablyconsiderableas“green.”Theseare1)LiquidOxygen(LOX),2)GaseousOxygen(GOX),3)HydrogenPeroxide(H2O2),and4)NitrousOxide.(N2O).Figure3comparestherelativeperformancesoftheseoxidizerswhenburnedwithastandardindustrialformulationofABS2inahybridmotor.ThesecalculationswereperformedusingtheNASAChemicalEquilibriumProgram (CEA).24 Plotted are (a)characteristic velocity c*,(b)flametemperature,(c)specificgravity,and(d)densityvelocityρ*,whichistheproductofthemeaneffectivepropellant density ρ and c*.Clearly,usingLOXandGOXasoxidizeroffers the bestmass efficiency c*; with peroxide offering the bestvolumetric efficiency ρ*. Unfortunately, LOX which offers bothoutstandingvolumetricandmassefficiencyasanoxidizer,mustberuled-outduetothelackofstorability.

Figure 2 Available options for hybrid oxidizers.35

Recent work by theWhitmore et al.,25,26 (1) and (2), and otherorganizations27,28haveadaptedhybridrocketsforusewithhighgrade(90%)hydrogenperoxideH2O2.Hydrogenperoxideisaveryefficientanddensepropellant,andmakesagoodoxidizer.Unfortunately,unlessusedinveryhighconcentrations(>98%),likeIL-basedpropellants,hydrogen peroxide is notoriously hard to ignite in a hybrid rocket.Mosthybridapplicationsusinghydrogenperoxide relyoncatalyticignition,andunfortunately,catalystdevelopmentremainsarealissue.In all current applications, systems require significant propellantpre-heatandconditioning.Noneof theexistingsystemsareable to2Stratasys ABSplus-P340.

achieve a reliable cold-start. Significant ignition time latencies areexperienced.Whitmore29 reports on the development of amodifiedversion of the previously-described arc ignition system where theinjected peroxide stream is thermally decomposed by leading theperoxide flow with a flow of gaseous oxygen that initiates weakcombustion of the fuel grain.During this pre-lead event, sufficientheat is released to thermally decompose the injected peroxidestream. The liberated heat and oxygen from decomposition drivefullcombustionalong the lengthof thefuelgrain.Gaseousoxygenpre-leadsassmallas500msreliablyinitiatecombustion.Multipleon-demandrelightsareprovidedwiththissystem.Unfortunately,aswiththecatalyticignitionsystems,thereexistconsiderablelatency-greaterthan1500ms-inordertoreachfullcombustion.

Unfortunately, the significant latencies and pre-requirementsmakePeroxideanunsuitableoxidizerforin-spaceapplications.Oncecombustionbegins,theperformanceoftheperoxidehybridsystemsisquitegoodwithvacuumspecificimpulsevaluesapproaching300seconds. Thus, hydrogen peroxide hybrids are more suitable forlaunchvehiclestageswhere ignition latenciescanbeabsorbed intothe mission timeline, and the high propellant/catbed heating loadscanbesuppliedthroughavailablegroundpower.Asdescribedearlier,the low-powerarc-ignitionsystemisakeyenabling technologyforin-space hybrid propulsion.To date, however, the vastmajority ofdevelopmentofthissystemhasreliedontheuseofgaseousoxygenastheoxidizer.Gaseousoxygenisanexcellentoxidizerandtheproposalteamhassignificantexperiencewithtestingofsmallhybridthrustersystems usingGOX. It is entirely “green” and can be quite safelyworked with at pressures below 2000 psig, as long as appropriatesystemscleanlinessstandardsareadhered-to.30Unfortunately,GOXeven when stored at high pressure has too low of a density to bepractical for long-term spacemissions requiring evenmoderateΔV levels. Thus, by process of elimination the proposers are limitedto theuseofnitrousoxide as theprimaryoxidizer for thisproject.NitrousOxidepresentsafinal,reliablealternativeforagreenhybridrocketoxidizerforspacepropulsion.Nitrousoxideisbyfarthemostcommonly used hybrid rocket oxidizer.N2O is an inexpensive and readilyavailablepropellant,andhaslongbeenconsidered“standard”oxidizer for hobby-rocket hybrid enthusiasts. Nitrous has the clearadvantageofbeingnon-toxictohumantissueandisclassifiedasnon-explosive,non-flammablebytheUS.OccupationalSafetyandHealthAdministration(OHSA).31

Hazards associated with using nitrous oxide as a rocket propellant

Nitrous Oxide exists as a saturated liquid below its criticaltemperature of 36.4ºC, and propulsion applications typically must


http://usglobalimages.stratasys.com/Main/Files/Material_Spec_Sheets/MSS_FDM_ABSplusP430.pdf




deal with N2Oinbothliquidandvaporform.AlthoughinliquidformN2O isquite inertandnearly impossible todetonate, invapor formN2O canexperiencea rapiddecomposition reaction.

32This stronglyexothermic reaction releases up to 1.864 MJoule per kilogram ofmaterialthatisdecomposed,alongwithlargeamountsofnitrogenandoxygengas.Generally,thedecompositionreactionactivationenergyEaislarge,andthevapormustbeheatedtotemperaturesgreaterthan800ºCinordertoinducedecompositionofpureN2Ovapor.Whenasaturated N2Ovapormixture iscontaminatedbyasmallamountofhydrocarbonmaterial, the relative stability of the vapor is loweredand the dissociation activation energy drops dramatically, and

decomposition reactions canoccur at near room-temperatures for ahighlycontaminatedmixture.33Ineffecttheadditionofhydrocarbonmaterial to nitrous oxide has the effect of catalyzing or “seeding”the decomposition event. Figure 4 shows the concept, where theactivationenergybarrier issignificantlyloweredbythepresenceofahydrocarboncontaminant.Becausenitrousoxideisahighlypolarmolecule,itisanexceptionallygoodsolventandreadilypicksupanddissolvesevenminoramountsofhydrocarbonmaterialsthatmayliealong theflowpath.This physical property further exacerbates thepotentialsafetyhazardsassociatedwithnitrousoxideoperations.

Figure 3 Performance comparison of 4 candidate hybrid oxidizers.

Figure 4 Hydrocarbon seeding of reduces N2O decomposition activation energy barrier.

The notorious July 26, 2007ScaledComposites fatal accident36 in Mojave, CA during Spaceship Two propulsion systems testing,

directly resulted from a runway decomposition of contaminatedN2O. The nitrous oxide was stored in an unlined composite tankand effectively acted as a solvent dissolvingminor amounts of thecomposite shell. These “hydrocarbon seeding” contaminants actedas a catalyst to significantly reduce the activation energy for thedecompositionreaction.Becausethetestwasbeingperformedduringa record heat wave with recorded ground temperatures exceeding120ºF(49ºC),asignificantvolumeofvaporwasbeingpumpedduringthegroundoperationstest.It isbelievedthatasmallsparkfromanungroundedpumpprovidedasmallbutsufficientquantityofenergytoinitiatetheconflagration.Threepeoplediedintheaccident.

Evenwhenstrictcleanlinessstandardsareadhered-to,theuseofpure N2Oasanoxidizerpresentsasignificanthazardforpropulsionapplications due to the close coupled nature of the oxidizer tank and the combustion chamber.As the motor burns and theN2O is depleted,asignificantvolumeofullagevaporwillcollectinthetank.Tank depletion also implies a significant drop in the internal tankvaporpressureduetoadiabaticcooling.Thisdropinvaporpressureprovides the opportunity for a reduced pressure drop across the injector,allowinghotcombustiongasses toenter thefeedlinesandpossiblythelowerportionofthetankitself.Theresultisasignificantpotentialforadecompositionreaction,resultinginafireorexplosion.





Mitigation of the N2O decomposition hazard

Fortunately, it appears that adaptationsof proceduresdevelopedby the medical and dental anesthesia community offers a strongmitigationtothisdecompositionhazard.34Duringuseofnitrousoxideforanesthesia,agaseoussolutionof50%byvolumeN2Oand50%O2 isadministeredtothepatient.TheN2O content provides the anesthesia properties, while the O2contentkeepsthepatientfromasphyxiating.In amanner directly analogous to the creationof soda-water usingdissolvedcarbondioxide; theN2O/O2 hybridsolution,referredtoasNytrox, iscreatedbybubblinggaseousoxygenunderhighpressureinto the nitrous oxide until the solution reaches saturation level.Oxygen concentrations as low as 30% can be used for anesthesiaapplications without encountering issues associated with patientasphyxiation.

Figure5plots thevapor/liquid/isothermdiagram for a saturatedN2O/O2solution.Figure5(a)plotsthevaporandFigure5(b)plotstheliquidphasemassconcentrationsofoxygeninsolutionasafunctionofsaturationpressure.Isothermcurvesfortemperaturesvaryingfrom-30ºCto30ºCareplotted.The0ºCisothermishighlightedasthesolidbluelineforboththeliquidandvaporsegmentsofthechart.Readingthisdiagram,at0ºCand86atmospheres(1250psig), thereexistsa“sweetspot”wheretheconcentrationofgaseousoxygenintheullageisamaximum,approximately37%,whiletheoxygenfractionintheliquid phase remains relatively low, approximately 13%.Note thatthepressurerequiredtoholdtheO2issolutionissignificantlyhigherthan the natural vapor pressure of N2O,which is approximately30atmospheres at 0ºC. This optimal point allows for the maximumproportionofvapordilutionwhilemaintainingahighdensityfortheliquidfluid.TheO2insolutionprovidestwoimmediatesafetybenefits.

Figure 5 O2, N2O solution vapor/liquid isotherm plots.34

First, the oxygenmixture in the ullage significantly dilutes thenitrous oxide vapor, and significantly diminishes any potential fora decomposition reaction. Figure 6 plots the minimum energy Ei requiredforapointsourcetostartaself-sustainingdeflagrationwavein nitrous oxidewith varying initial concentrations of oxygen. Forpurenitrousoxidevaporthisenergyisonlyabout400-500miliJoules;however,onlya10%O2 concentration increases Eitoavaluegreaterthan5joules,anorderofmagnitudeincrease.A35%O2 concentration-easily achievable at pressures above 100 atmospheres-increasesEi togreater than1000 joules,byafactorofmore than4000!Infact,analytical studies performed by Karabeyoglu35 have demonstratedthatblendedN2O/O2vaporwithat least20%concentrationofO2 is virtuallyimpossibletoigniteusinganyconceivableignitionSecond,the presence of O2inthesolutionsignificantlyincreasesthe“quenchdiameter,”thediameterofametalpipethatwillquenchanypotentialdecomposition reaction and ensure that any potential deflagrationwave will not propagate. Figure 7 shows this behavior. Note thatthequenchdiameterforpurenitrousoxideisapproximately1.4cm(0.6 in),andgrowstomorethan4.5cm(1.77 in)foronly15%O2 in solution.Thisdifferenceisafactorofmorethan8.7intermsoftheallowablepipingcrosssectionalarea.Thisallowablegrowthisquitesignificantinthatitallowsforsubstantiallyhighermassflowlevelsinthesystemwithnoincreaseindeflagrationrisk.

ThekeypropertiesassociatedwithusingN2O/O2 mixturesinlieuof pure N2Oaresummarizedas

a. BlendedOxidizersystemswouldbemuchsafer thanpureN2O becausevaporphasehaslargeO2 concentration.

b. Typicalsystemwith30%oxygeninvaporphaserequiresfourtofive orders ofmagnitude larger ignition energy comparedwithpure N2O.

c. Safe partial self-pressurization possible at high densities. Thispropertygreatlysimplifiesthedesignandeliminatestheneedforaheavy,separatepressurantsystemusingheliumornitrogen.

d. Blended Oxidizer allows improved Isp performance comparedwith pure N2O.

e. Critical control variables are temperature and pressure, whichdetermineO2 massfractioninliquidandvaporphases.

f. Atagivenpressurelevel,NytroxdensityhigherthanGOXbyafactorof3or4,andallowsforasignificantimprovementintheoverallvolumetricefficiencyofthepropellants.

The calculations of Figure 5 were performed using the Peng-Robinson36 2-phase state-equation for binary solutions. Theimplemented numerical algorithm follows the procedure laid outby Karabeyoglu.35 The mixing rule used to combine the binarycomponentsisbasedonthemodelofZudkevitch&Joffe.37Forafluidgiventemperature,thealgorithmsearchesfortheequilibriumpressurelevelthatmatchesthefugacityofthevaporandliquidphasesforeachofthebinary(O2, N2O)fluidcomponents.





Figure 6 Minimum ignition energy for N2O/O2 mixtures at 3 pressure levels.35 Figure 7 Quenching diameters for N2O/O2 mixtures at 3 pressure levels.35

Performance of the N2O/O2 oxidizer mixtures

Figure 8 plots the densities of the vapor and liquid phases, ascalculated by the previously discussed Peng-Robinson model.ReferringtoFigure5,at0ºCa90%massconcentrationofN2O in the liquidsolutioncorrespondstoavaporpressureofapproximately75atmospheres(1100psia).Atthisvaporpressurethesolutiondensity

isapproximately800kg/m3.Atapressureof120atmospheres(1470psia), the percentage of nitrous oxide in the liquid solution dropsto only 70%with a corresponding density of only 590kg/m3. Thisbehaviorseemscounter-intuitive,butisthenatureoftwophasebinarysolutionswhere the nitrous oxide and oxygen components becomemutuallydissolved

Figure 8 Density of Nytrox vapor and liquid phases at vapor pressure for 6 different isotherms.

Although the solution ofO2 into nitrous oxide slightly reducesthe density of the oxidizer, the overall effect includes moderateenhancements of the Isp and a significant reduction of the optimalO/F ratio.This performance-trademakes theN2O/O2 solution only slightlylessvolumetricallyefficientthanwhenpurenitrousoxideisused.Figure9presents theseperformancecomparisons.Plottedarethe(a)characteristicvelocityc*,(b)vacuumIsp,(c)specificgravity,and (d) density Isp ρ*,which is theproduct of themeanpropellanteffective density and the specific impulse. The plotted curves arefor5differentoxidizerswhenburnedwith3-DprintedABS;GOX, pure N2O, 90% N2O/10% O2, 70% N2O/30% O2,and50%N2O/50%O2.ForsimplicitytheNytroxblendswillbereferredtobythemass-percentageofnitrousoxideinthefluidblend;respectively,Nytrox 90, Nytrox 70, and Nytrox 50.

The values plotted on Figure 9were calculated using the CEAprogram,24 assuming chamber pressures varying from 100 to 500psia.Thevacuum Isp calculations assume a 40:1nozzle expansion-ratio.Thespecificgravitycalculationassumesastoragepressureof1250 psig (86 atms.),andfueldensityof1.04g/cm3.AlsoplottedonFigures8(b)and8(c)aretheIsp and ρ*Isp ofHydrazine.NotethatthehybridmassIspperformancesignificantlyexceedsthatofhydrazine.Thedensityperformanceρ*Isp of the Nytrox 90solutionisgreaterthanhydrazine, whereas the Nytrox 70isslightlylower.AsexpectedusingGOXastheoxidizerresultsinthemostmass-efficientsystem,butthelowGOXstoragedensityresultsinthelowestdensityimpulse.

ConverselyusingpureN2Ogives thebestvolumetricefficiency,but results in the lowest specific impulse and requires significantly





moreoxidizerinordertoreachoptimalIsp.Thecurvecorrespondingto the Nytrox 90mixture(at75atmospheresvaporpressure)givesthebest compromisewithadistinctρ*Isp optimumoccurringat anO/F ratio of approximately 4.2.As described previously, increasing the

pressure to 84 atmospheres dilutes the nitrous oxide slightlymore,butallowsthemaximumpercentageofdissolvedoxygeninthevaporphase, and is an important considerationwith regard tooperationalsafety.

Figure 9 Performance of 3 N2O/O2 concentrations compared against pure N2O and GOX as oxidizer.

Experimental apparatus, instrumentation, and test procedures

This section reports the experimental apparatus and proceduresdeveloped for this testing campaign. The hardware details andlaboratory procedures for manufacture of the Nytrox solution arepresentedfirst.Thisdiscussionisfollowedbydescriptionofthetestarticle, experimental apparatus, and procedures used for the hot-fire comparison tests. Because the procedures, established by adhoc experimentation during this testing campaign are essential tomanufacturing theNytroxmixturewith thepropergasproportions,thefillandpercolationprocedureswillbedescribedindetail in thefollowing paragraphs.As a cost-saving measure, many of the testhardware itemsusedwereborrowedfromprevious testarticlesandadaptedforthiscampaign.ThesmallhybridthrusterusedforthistestserieshadpreviouslybeenoptimizedforusingGOXastheoxidizer;was adapted without change for the Nytrox systems. Due to thepreliminarynatureofthisrestingcampaign,presentedNytroxresultsarebynomeansconsideredtobeoptimal.

Nytrox solution processing equipment and fill procedures

Forthisstudyhighly-purifiedgradesofnitrousoxideandgaseousoxygenwereusedinordertoensuretheresultingNytroxmixturewasfree from contaminants and any possible catalytic agents. The gas

supplier3 quotes the N2Opurityat99.7%byvolume;withtheprimaryimpurities being traces of oxygen, nitrogen and water vapor. TheGOXpurityisquotedas99.4to99.7%,withthemainimpuritybeingargon.Argon is not liquefiable at normal temperatures, and sinceargon’scriticalphaseconstantsaresoclosetooxygen,itspresenceisconsiderednegligiblewithregardtothemixingproperties.Also,sinceargonisinert,thereisnopotentialforcatalyticeffects.

The basic procedure consists of filling the run tank with thedesiredweightofN2O,connectingthefilledtanktoaGOXsupply,andallowingtheGOXtobubbleupthroughtheliquidnitrousoxide.Adiptubeis requiredontheruntanktoallowGOXtopercolateupthroughtheliquidphasenitrousoxidewithoutinvertingthetank.Thedip tubealsoallowsdirectdeliveryof liquid-phaseNytrox forthe hot fire tests. During passage through the liquidN2O, oxygendissolves into solution and also droplets of nitrous oxide are carried up into thegasphase.Thenet result is that thevolumeof liquid inthe cylinder steadily diminishes until equilibrium vapor and liquidphaseproportionsarereachedforthefluidtemperature.Theobjectiveof the developed procedurewas to generate aNytrox solution thatpossesses amaximumconcentrationof oxygen in the vapor phase,whilemaintainingahighN2Oconcentration in the liquidphase.Asdescribedearlier,anddepictedbyFigure5andFigure8,at0ºCthisoptimumoccursat approximately86atmospheres (1250psig).Theresultisa“Nytrox 87”solutionwithavaporphaseO2 concentration

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of 37%, and a liquidphaseO2 concentrationof only13%.For thisequilibrium condition the liquid-phase Nytrox 87 solution has a density of approximately 0.785g/cm3. This value is compared to aliquid-phase density of pure N2O of 0.907g/cm3at0ºC,whichisonly15%higher.Usingtheidealgaslaw,GOXatthesametemperatureandpressurewouldhaveadensityofonly0.120g/cm3,ormorethan6.5timeslessthandensethantheNytrox 87solution.

Nytrox tank fill apparatus and fill procedure

ThehighgradeN2OisdeliveredinaK-sizetank.TheGOX-supplyalso comes delivered in aK-size tankwith an internal pressure of2000psig.ToensuresafetyduringtheNytroxmixingprocedure,thepipesandfittingsasprocuredwereexclusivelyratedforNitrousOxideservice.Also,allpersonnelpresentduring themixingprocessworethepropersafetyequipment.TheNytroxwasmixedinacommercialNOS®tankwitha10-lbmfillcapacity,anddesignedforautomotiveapplications.4Thisparticularunitcomeswithapre-installeddiptube,hasadesignburstpressureof8000psigandafactoryinstalledburstdiscratedto3000psig.Safetyofusingthistankwasverifiedsincethepressuresdesiredwerewellbelowtheburstdiscpressure.Allservicelineswerefabricatedfrombraidedstainlesssteel,andarespecificallyrated for nitrousoxide service.Tobegin theNytroxmanufacturingprocedure,theNOSruntankisfirstfilledwiththedesiredamountofliquidnitrousoxide,typically5-7lbs.(1100-1,550grams).Toprotectallpersonnel in theadventofanunlikelydecompositionevent, theN2Ofillprocedureisperformedinawirecage.Figure10showstheNOSruntankfillapparatus.TheNOSruntankwasplacedinanicebathtolowerthetanktemperatureto0ºC,whiletheN2OK-servicetank was kept at room temperature. The temperature differencecreatedbytheicebathlowersthevaporofthefluidintheNOSruntank,creatingapressuredifferencethat initiates influidflow.Afterensuringthattheneedlevalveandbothtanksareclosed.

An electronic scale with serial output is used to measure thenitrous oxide mass moved from the service tank to the run tank.Beforefillingtheemptytank,weightwasrecorded,andthescalewastared.TheNOSruntank/icebathcomboisthenplacedonascaleusedtomeasuretheweightofN2Oaddedtothetank.Withtheneedlevalveclosed,both theN2O tankand theNOSrun tankwereopened.TheneedlevalvewasthenopenedslowlytoallowflowofN2O into the NOSruntankataslowrate.Oncethescaledisplayreadsthedesiredmass,theneedlevalvewasclosed,followedbytheN2O service tank andNOSruntankvalves.Slowlydisconnectingthefilllinefromeachbottleallowslinestoventduringremoval.

Nytrox/O2 percolation procedure

Allgas-mixingprocedureswereperformedintheBattery Limits and Survivability Testing (Blast)Lab,USU’son campus jet engineandrockettestfacility.Thisservicebunkerhas1-footthickconcretewalls with two 6” thick Plexiglas viewing pane from which testconductors can view hazardous operations directly in an indoors shirt-sleeve environment. Conveniently the Blast lab is located directlyacrossthestreetfromthePRLfacility.Figure11showsthephysicallayout of the assembled system (a), and (b) shows the percolationapparatus block diagram. Two different service lines are used formixing theNytrox.Thefill line fromtheN2O to theNOSrun tankwasapproximately8ft.longconsistingofthefollowingcomponents:four2ft. linesections,oneN2Ofilter toensurecleanlinessofN2O, 4Anon.NOS 10lb Super Hi-Flow Nitrous Bottle Kit.HolleyInc.;2019.

onebackflowprevention (check)valve ratedat3000493psig, andoneprecisionflow-adjustment(needle)valveratedat2000psig.ThefilllinefromtheO2totheNOSruntankwasapproximately4ft.longconsistingofthefollowingcomponents:two2ft. linesections,onebackflow prevention (check) valve rated at 3000 psig, a precisionflow-adjustment (needle) valve rated at 2000 psig, and a pressureregulatorratedat3000psig.

Figure 10 N2O fill apparatus.

After ensuring that the needle valve and both tanks are closed,the O2filllineisattachedtoboththeO2 tankandtheNOSruntankwith thecheckvalveallowingflow into theNOSrun tank.TheO2 tankisthenopenedandthepressureregulatorissettoadownstreampressureof1250psig.TheNOSruntankisthenopened.SinceaveryslowflowofO2isdesiredintotheNOSruntank,theneedlevalveisopened just until O2flowcanbeheard.Thisconfigurationisthenleftto allow the O2 topercolate through theN2O currently in theNOSruntankandreachpressureequilibrium.OncepressureequilibriumisreachedandthepressureintheNOSruntankisconfirmedandtheneedlevalveandbothtanksareclosed.Asbefore,theservicelinesareslowlydisconnectedfromeachbottleallowingthelinetoventduringremoval.OncetheNOSruntankisfilled,afinalmassisloggedbeforestoringtheNytroxforfuturetesting.Tofurthermitigateanypotentialriskofrunawaydecompositionreaction,theservicedNOSruntankisstoredpotableinafreezerunittokeepvaporpressureslowandensureaminimal amount ofN2O vapor in the tank ullage.By decreasingthetemperatureoftheNOSruntank,theactivationenergybarrierisraisedevenfurthertopreventanyaccidentfromoccurring.Internalfreezertemperaturesarekeptaround–10ºC.

Hot fire test apparatus and instrumentation

This section details the hardware, instrumentation and testproceduresusedtoperformthehotfireevaluationtests.Thehotfiretestingcampaignwasperformedusingthepreviously-describedBlastLab test cell.Whitmore and Bulcher7 andWhitmore8 describe the


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analyticalmethods, test apparatus, instrumentation, testprocedures,andanalysismethodsusedtoderivethepresenteddatainmuchfuller

detail.Thissectionsummarizesthemajorconclusionsofthistestingcampaign.

a) Physical Layout

Figure 11 Nytrox percolation apparatus.

Thrust chamber

The legacy GOX/ABS small spacecraft thruster of Refs.7,8 was adaptedforuseinthetestingcampaign.Figure12presentsthedetailsofthethrustchamberassembly.Figure12(a)presentsa2-Dschematic.Figure 12(b) presents a photograph of the disassembled system.Depicted are the major components; i) graphite nozzle, ii) nozzleretentioncap,iii)motorcase,iv)3Dprintedfuelgrainwithembeddedelectrodes,v)insulatingphenolicliner,vi)chamberpressurefitting,andvii)single-portinjectorcap.The38-mmdiameterthrustchamberis constructed from 6061-T6 high-temperature aluminum, andwasprocuredcommerciallyfromCesaroniInc.5Table1summarizesthethrustergeometryandotherspecifications.Theelectronicarc-ignitionsystemforthisthrusterwasdescribedpreviouslyandisdepictedbyFigure1.

Hot fire thrust stand apparatus and instrumentation

Figure13showstheflightweightmotorassembled(a)andmountedto the test load balance, (b) shows the piping and instrumentationdiagram(P&ID)ofthetestsystems.Teststandmeasurementsincludeventuri-based

GOX massflow measurements, load-cell based thrustmeasurements,chamberpressure,andmultipletemperaturereadingsatvariouspointsalongtheflowpath.Thedifferentialventuripressuretransducer was installed to increase the accuracy of the sensed pressure drops. The thrust-stand support members allow bendingalongthedirectionofthrusttopreventthemfrominterferingwiththemeasuredload.Theentiretestassemblyismadeusingcommerciallyavailable T-slot6 extruded-aluminum components. Figure 14 showsthe instrumentation deck layout.On the top of the instrumentationdecktherearethreeNIDAQunitsshown;(lefttoright)USB6009,USB6002,USB9213.TheNIUSB6002isusedtoreadandwritedata from several bridge transducers and acts as the controller forthehigh voltage signal using a single TTL-level (3.2-volt) digitalcommand.TheNIUSB6009servedasanadditionaldevicetoread5Anon.Cesaroni Pro-X, A Better Way to Fly.Pro38®hardware.6Anon.T-SlottedAluminumExtrusions;2016.

andwritedatasinceallthechannelsontheUSB6009wereused.TheNIUSB9213servedasareadandwritedeviceforthethermocoupleprobe inserted in theflow to record the temperatureof theNytrox.AlldataacquisitionandcontrolprocesseswereprogrammedontoacontrollaptopcomputerusingtheLabVIEW®programminglanguage.Communication from the laptop to the instrumentation systemwasachievedbyusinga30-ftamplifiedUSB2.0extensioncable.

Theignitionsystempowerprocessingunit(PPU)isbasedontheUltraVolt® D-series line of high-voltage power supplies (HVPS).7 AspreviouslypicturedinFigure1,theHVPSprovidestheinductiveignition spark that pyrolyzes sufficient ABS material to seedcombustion.TheD-seriesHVPSunits takea15-voltDC inputandprovideacurrent-limited(7.5mA)highvoltageoutput-upto1000Vor6Wattstotaloutput.Previousexperiencewiththisignitionsystemhasdemonstratedthatignitioncanbereliablyachievedusingaslittleas3watts.Dependingontheimpedanceonthearcpathbetweentheignitorelectrodes,thedissipatedvoltagetypicallyvariesbetween10and400volts.Totalenergyofignitionistypicallylessthan3Joules.Ignitionenergyresultswillbereportedlaterinthispaper.

Directly aft of the thrust chamber lies the solenoid actuatedGOXrun-valve.Thesolenoidflowvalveisactuatedviaadigitaloutcommandfromtheinstrumentation.TheNationalInstrumentsUSB-6002initiates“IgnitionControl”sendingpowertothesolenoidvalveviathesolid-staterelayandHVPSTTL-levelactivatesignalusingtheNI6002as thecontroller.The24Vpowersupply isused tosupplypower to the solenoid valve and HVPS; whereas, the 15V powersupplyisusedtopowerthetransducers.Thethermocouples,venturiinlet,differential,andchamberpressure transducers,alongwith theloadcellallhavetheirsignalsconditionedusingNationalInstrumentsDataAcquisition(DAQ)units.

Hot fire test procedures

Initially a set of baseline tests was performed using gaseousoxygenastheoxidizer.Thistestserieswillensurethatthesystemhas7Anon. High Power 8C-30C Series, Single Output High Voltage DC/DC Modules.UltraVolt,Inc.,2016.


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beenreturned to thestatus thatexistedduring the testingcampaignof Ref.7 Key parameters to be measured during this baseline testseries include ignition power, thrust, chamber pressure, massflow,fuel regression rate, and specific impulse. Following the baselinetests,theGOXtankwasswappedfortheNOSruntankfilledwiththeprocessedNytrox.Otherthanthechangeinoxidizer,thetestassemblywillremainidentical.FortheNytroxtests,thedownstreamregulator

setting is adjusted tode liver this blendedmix at exactly the samechamberpressureasfortheGOXtests.Specialattentionwasplacedonestablishingtherequiredignitionpower,andtheresultingthrust,specificimpulseandfuelregressionrates.TestswereperformedusingtheNOSruntankatroomtemperature;andalsowiththetankchilledbyanIce-bathtoensurethatliquidsolutionisinjectedintothemotor.

(a) 2-D Schematic (b) Thrust Chamber ComponentsFigure 12 Test article thruster assembly.

a) Mounted to Test b) Piping and Instrumentation

Figure 13 Thruster systems test apparatus.

Theprocedures followedwere the same forbothGOXbaselineandNytroxTests. Before themotorwas assembled, the fuel grainweightandportdiametersatboththetopandbottomwererecorded.Thenozzlethroatandexitplanediameterswerealsologged.Finally,theNOSruntankweightandpressurewerelogged.Oncethepre-testmeasurementswere recorded themotorwas assembled.Themotorassemblyleadswereconnectedtothecart,oxidizerfeedlineattached,andmotorassemblymountedtotheteststand.TheteststandwasthenmovedtotheBlastLabtestareafortesting.Insideofthetestarea,A/Cpower was then connected to the test stand and connectivity checked usingthedesignatedlabtestcomputer.ThefeedlinefromeithertheGOXortheNOSruntankwasattachedtothethrustersystems.Theentire feed line was then leak checked to ensure proper connections, andtheregulatorfortheoxidizerfeedpressurewassetto310psig.

For this testcampaignpower to the ignition“spark”powerwas

active for a total of two seconds, pre-leading the opening of theoxidizer run valve by 1 second. The oxidizer run valve was pre-programmedtoopenforaprescribedamounttime,andfortheseteststhis timevaried from1 second tomore than4 seconds.Themotorwould snuff immediately after closure of the run valve. Typically,one fuel grain allows for 8 seconds of burn time, so a typical testserieswouldallowfourtestsof2secondseachonasinglefuelgrain.Followingeachburn,thepreviouslydescribedweightandgeometrymeasurementswererepeatedandlogged.

Analytical methods

Computational sequence for calculating the GOX massflow

For the GOX mass flow sensor, the massflow calculation wasrather straight forward, and the compressible venturi massflowequations are derived frommaterial presented byAnderson.38 The





enteringstagnationpressureiscalculatedfromthesensedinletP1 and throat P2absolutepressurelevels,andtheventuriinletA1 and throat flowareasA2

( ) ( )

( ) ( )

21 1

11 2

20 2

2 21

1 22

AP P

AP

AP P

A

γ γ

γ γ

γ γ

+ +

⋅ −

=

⋅ −

(1)

Once the true inlet stagnation pressure is calculated, then theachieved massflow is calculated using the unchoked compressiblemassflowequation

2 1

1 11 0

0 0

2 11ox d

P Pm C A PRg T P P

γγ γγ

γ

+ = ⋅ ⋅ − − ⋅

(2)

Figure 14 Instrumentation deck layout.

ThecalculationofEq.(2)requiresatemperaturemeasurementT,and thisvaluewasmeasuredusinga thermocouple installedon theventuri flow block. The flow discharge coefficientCd accounts for frictionalflowlosses.Bothventuriflowmeterswerecalibratedusingcoldflowteststhatcapturedthetotalmasspassedthroughthesystem.PreviouslyRefs.,7,8performedextensivecoldflowtestsandmeasuredthe discharge coefficient for GOX flow to be approximately 0.95.SincethetestsetupfortheGOXbaselinetestsdidnotchangefromtheoriginaltestsofRefs.,7,8theVenturiwasnotcalibratedusingcoldGOXflowforthiscampaign.

Computational sequence for calculating the Nytrox massflow

In contrast to theGOXflow,due to the two-phase, binaryfluidnature of the Nytrox solution flow, deriving meaningful massflowmeasurements from the venturi sensor is rathermore complicated.Multiple models have been previously developed for two phasenitrousoxidemassflows.TheseincludemodelsdevelopedbyZilliacand Karabeyoglu,39 Dyer,40 Whitmore & Chandler,41 Zimmermanet al.,42 and Waxman et al.43 It is likely that these models, eachdevelopedfor theflowofasinglesaturatedliquidareapplicabletothe two phase binaryfluid injector problem, but a solid theoreticalfoundationforthisadaptationhasyettobedeveloped.Thus,forthispreliminaryproof-of-concepttestingcampaign,theNytroxmassflow

throughtheventuriwasmodeledasasimplecompressiblegasflowwithacalibrateddischargecoefficient.Here,anidealgasisassumedwith the gas properties derived from vapor phasemole fraction ascalculated by thePeng-Robinsonmodel.36The associated ideal gasthermodynamicpropertiesare

a. Molecularweight,

2 2 2 2w f w f wNytrox N O N O O OM M M= ⋅ + ⋅M M

(3)

b. GasConstant,

ugNytrox

wNytrox

RRM

= (4)

c. SpecificHeatatConstantPressure,

2 2 2 2 2 2f w p f w wN O N O N O O O OpNytrox

wNytrox

M C M MC

M

⋅ + ⋅ ⋅=M M

(5)

d. RatioofSpecificHeats

p pNytrox NytroxNytox

v p gNytrox Nytrox Nytrox

C C

C C Rγ = =

− (6)

InEqs(3)-(6)thesymbolMf representsthemolefractionofagivenvapor species and Ru represents the universal gas constant. Usingthese values for Rg and γ, theNytroxmassflow is calculated usingEqs.(1)and(2)

Computational sequence for fuel regression rates, instantaneous O/F, and equivalence ratios

Although the inline venturimeasures the oxidizermass flow inreal-time,theteststandcouldnotmeasurereal-timefuelmassflow.Thus, for this testing campaign the “instantaneous” fuelmassflowrateswerecalculatedashedifferencebetweenthemeasurednozzleexitandoxidizermassflowrates,

fuel total oxm m m= − (7)

Forallrunstheregulatorpressureandinjectorportdiameterwerepre-set to choke the injector flow and ensure a constant oxidizermass flow. Choking the injector flow ensured very low run-to-runvariability in the oxidizermassflow rate, and significantly reducedthe risk of incurring injector-feed coupling instabilities duringcombustion.Thenozzleexitmassflow timehistorywascalculatedfrom themeasuredchamberpressure timehistoryP0, nozzle throat area A*,andtheexhaustgasproperties(flametemperatureT0, ratio ofspecificheatγ,molecularweightMw,andspecificgasconstantRg)usingthe1-dimensionalchokingmassflowequation,(Anderson[38],Chapter4).

( )11

00

2*1total

gm A P

R T

γγγ

γ

+−

= ⋅ ⋅ ⋅ ⋅ + (8)

Themeanlongitudinalfuelregressionratewascalculatedfromthefuelmassflowby,

2

fuelL

fuel L

mr

r Lπ ρ=

⋅ ⋅ ⋅

(9)

Integrating Eq. (9) from the initial condition to the burn timeframessolvesforinstantaneousmeanportradius,

( ) 2

00

1 t

L fuelfuel

r t r m dtLπ ρ

= +⋅ ⋅ ∫ (10)





Theterminalcrosssectionalareaofthefuelportis,

( ) 2

0fuel

c burnfuel port

MA t r

Lπ

ρ∆

= ⋅ +⋅

(11)

The mean fuel regression rate over the duration of the burn iscalculatedby

( )0

fuel

fuel port t burnburn

Mr

L r r tπ ρ

∆=

⋅ ⋅ ⋅ + ⋅ (12)

The mean oxidizer mass flux, total mass flux, O/F ratio, andequivalenceratioareestimatedby

( )

( )

( )

( )

0

0

.

.

tburn

ox

oxc burn

fuelox

c burntburn

ox

fuel

stoich

actual

m t dtG

A tM

GA t

m t dtO F

M

O FO F

= ∆ = = ∆ Φ =

∫

∫

(13)

For each data point in the burn time history, two-dimensionaltablesof thermodynamicand transportpropertieswere interpolated

tocalculate thegasconstantRg, ratioof specificheatsγ, andflametemperature T0. The table of equilibrium properties of the GOX/ABSexhaustplumeweredevelopedbyRef.3withmeasuredchamberpressure P0, combustion efficiency η*, and mean O/F ratio as independentlookupvariablesforthetables.Reference3usedNASA’sindustrystandardchemicalequilibriumcodeCEAcode24toperformthecalculations.Eachfuelgrainwasburnedmultipletimestoallowinterim fuel mass consumption measurements between burns. Thecorrespondingoxidizermassconsumedwascalculatedbyintegratingtheventurimassflowtimehistoryovertheburnduration.ThemeanO/F ratio over the burn duration was estimated by dividing theconsumedoxidizermassbytheconsumedfuelmass.Byadjustingη* theflametemperaturewasscaled

*2

0 0iactual dealT Tη= ⋅ (14)

Toadjust nozzle-exitmassflowand the resulting consumed fuelmassflow,

( )

0

t

fuel total oxM m m dt∆ = −∫ (15)

Adjusting inputcombustionefficiencyupwardshas theeffectofincreasing the calculated fuel mass consumption, and downwardsdecreases thecalculatedfuelmassconsumption.Thefuelmassflowcalculationstartswithanassumedcombustionefficiencyofη*=0.90.ThecalculationsofEquations(3-10)wereiterated,adjustingη* until thecalculatedfuelmassequalsthemeasuredmassandtotalconsumedpropellant O/F (ΔMox/ΔMfuel) within a prescribed level of accuracy(0.5%).

Table 1 Motor geometry and parameter specifications

Parameter Injector Single Port, 0.127 cm (0.05 in.) Diameter

Fuel Grain Diameter: 3.168 cm (1.246 in.)

Length: 6.850 cm (2.70 in.)

Initial Weight: 50.0 gInitial Port Diameter: 0.625 cm (0.246 in.)

Print Density: 0.955 g/cm3

Motor CaseDiameter: 3.8 cm (1.50 in.)

Length: 13.8 cm (5.43 in.) Wall Thickness: 1.5 mm (0.059 in.)

Conical Graphite Nozzle

Initial Throat Diameter: 0.375 cm (0.148 in.)

Exit Diameter: 0.577 cm (0.277 in.)

Ambient Tests Initial Expansion-ratio: 2.07:1

Nozzle Exit Angle: 5.0 deg.

Once the total mass flow and combustion chamber propertieswerecalculatedasdescribedabove,the1-dimensionaldeLavalflowequations(Anderson[38],Chapter4)wereusedtocalculatetheexitplane Mach number, pressure, effective exhaust velocity, thrust,thrust coefficient, specific impulse, and characteristic velocity.ThefollowingflowsequencewasusedforthedeLavalflowmodel

a. NumericalSolutionforExitPlaneMachNumber,

( )1

2 121 2 11* 1 2

exitexit

exit

A MA M

γγγ

γ

+− − = + +

(16)

b. ExitPlaneStaticPressure,

0

12112

exit

exit

Pp

Mγγγ −

=− +

(17)

c. EffectiveExhaustVelocity,

( )0

2112

g exit exite exit exit

totalexit

R T p p AC M

mM

γλ γ

∞⋅ ⋅ − ⋅= ⋅ ⋅ +

−+

(18)

d. ThrustandThrustCoefficient,

total eF m C= ⋅ (19)

0 *FFC

P A=

⋅ (20)

e. Finally,SpecificImpulse,CharacteristicVelocity,andDensitySpecificImpulse

0sp eI g C= ⋅

(21)

0 **total

P Acm⋅

=

(22)

0* sp g spI S g Iρ = ⋅ ⋅

(23)





InEq.(18) ( )( )1 2 1 cosexit exitλ θ= + , where exitλ isthemomentum

thrust correction factor and exitθ is theconicalnozzleexitangle.InEq.(21)g0 isthenormalaccelerationofgravityatsealevel,9.8067 m/s2.Fortheρ*Isp calculation, sgisthemeaneffectivespecificgravityofthepropellants, and is calculated as

1

g gox fuelg

OS SFS OF

⋅ +=

+ (24)

In Eq. (24) the parameter refers to the storage specific gravityof theoxidizer andnot thedownstreamspecificgravity.The thrustcoefficientCFandspecificimpulse Isp werealsocalculatedfromthethrustvaluessensedbytheteststandloadcell.Thevaluescalculated

by Eqs. (20) and (21) provide redundant measures, and will bepresentedlaterinordertosupporttheverisimilitudeofthecollectedtestdata.

Results and discussionThe results of the previously described testing campaign are

presented in this section. The properties of the resulting Nytroxsolution are described first, followed by the test results from the2-PhaseNytroxventuricalibration.Finally, theresultsfromsixteensuccessful tests are presented.Of those tests, 6 hot-fire burns usedGOXastheoxidizer,and10hotfireburnsswappedoutNytroxforGOX.TheGOXresultsestablishedthesystembaseline.ResultsfromtheGOX andNytrox burn testswill be presented individually andthencompared.Table2summarizestheNytroxprocessingresults.

Table 2 Nytrox mix batch specifications

Nytrox batch No.

NOS tank tare weight

NOS tank internal volume Final tank fill pressure Final tank fill

temperature Final filled tank weight

1 7.657 kg 6.846 liters 1250 psig (8704.4 kPa abs) 0ºC 10.524 kg

2 7.657 kg 6.846 liters 1270 psig (8842.3 kPa abs) 0ºC 10.578 kg

Oxidizer added

Mean oxidizer density N2O added to tank O2 added to

tank Total O2 mass fraction, %

1 2.867 kg 0.419 g/cm3 2.234kg (4.924 lb) 0.633 kg 22.10%

2 2.921 kg 0.427 g/cm3 2.267kg(4.996 lb) 0.654 kg 22.40%

Liquid density (Peng-Robinson)

Vapor density (Peng-Robinson) Liquid O2 mass fraction Vapor O2

mass fraction Liquid Mol. Wght. Vapor Mol. Wght.

1 0.785 g/cm3 0.241 g/cm3 12.63% 37.14% 42.02 g/mol. 38.63 g/mol.

2 0.782 g/cm3 0.247 g/cm3 12.63% 37.14% 41.97 g/mol. 38.62 g/mol.

Mass of liquid in tank

Mass of vapor in tank Volume of liquid in tank

Ullage volume in tank

Mix quality in tank

1 1.757 kg 1.110 kg 2.238 liters 4.608 liters 0.387

2 1.799 1.122 2.302 liters 4.544 liters 0.384

Making the Nytrox mixture

TheproceduresdescribedbySectionHwerefollowedtogeneratetheNytrox batches used for this testing campaign. For each batchthe NOS run tankwas filledwith 5 lb. (2.27 kg) of nitrous oxide, and allowed to chill in the ice bath.A5 lb. fill is 1/2 of the ratedfillcapacityforNOSruntank.Oncethetanktemperaturestabilizedat 0ºC, the O2needlevalvewasopenedandoxygenwasallowedtopercolatethroughthesystem.Onceconnectedwiththeregulatorsetat1250psig, theprocess takesabout2hours to reachequilibrium.Atthetimeofthispublication,twocompletebatchesofNytroxhavebeenprocessed.Thebatchcomparisonsareremarkablysimilar.Batch1reachedanequilibriumpressureof1250psig,withbatch2reachinganequilibriumpressureof1270psig.Batch2endedupwithslightlymoreoxidizerinthefill2.921kgascomparedto2.891forbatch1.However,theintensivefluidpropertieswereremarkablysimilar,withthe effectivefill densities for batch 1 and batch 2 being0.419 and 0.427 g/cm3respectively.Thisdifferenceisoflessthan1%.Thus,theestablishedfillprocedureswerequitesuccessfulandworkedaswellasplanned.

Nytrox venturi calibration

Asdescribed in theprevioussection, theflowpf the two-phase,binaryN2O/O2fluidmixturethroughtheventuriflowmeterisquitecomplexandafirst-principleflowmodehasnotyetbeendeveloped.Thus, a simple calibration procedure was performed in order tomeasurethedischargecoefficientCdwithsufficientaccuracytoobtainreasonableNytroxmassflowresults.ThesetestswereperformedusingtheBatch2Nytroxmixture.Atotalof10cold-flowcalibrationrunswereperformed,withthefirst5batchesflowingfor2secondseachandthelast5flowingfor10secondseach.Theregulatorpressurewasset to310psig,andtheoxidizer in theNOSrun tankwasweighedbeforeandaftereachburn.Thecoldflowtestapparatuswasidenticalto the previously- described hot flow setup, except that the thrustchamberandfuelgrainwereremovedandtheignitionsparkwasnotinitiated.

Figure 15 plots the cold-flow test results. Plotted are the testdata, a linear least squares curve fit, and the curve uncertaintiesboundaries plotted at the 95% confidence level.The abscissa plots





thetotalintegratedmassflowoverthecold-flowrunaspredictedbythecompressibleventurimodelfromEqs.(1)and(2),assumingthatCd=1.0.Theordinateplotstheactualflowedoxidizermassmeasuredfromthepre-andpost-testweightsoftheNOSruntank.Thecurvefit

coefficientsarealsonotedonthisfit.Generally,thefitisrathergoodwithonlyslightbiasofabout2.6grams,whichislikelyduetoatanktareweighterror.

Figure 15 Nytrox Venturi calibration data.

Interestingly, the curve fit shows a slope (corresponding to theeffectivedischargecoefficient)valueatapproximately1.389;avaluewhichistheoreticallylargerthanthemaximumpossiblevalueof1.0foranidealgasflow.Clearly,thereweresomeun-modeledtwo-phaseeffectshappeningduring thisflow.With thismethod themolecularweight,gasconstant,and ratiospecificheatswerecalculatedbasedonthetankullagevaporcompositionasshownbyTable2,withbeingcalculated from the averages of the two Nytrox batches. Pressingforward to the Nytrox Hot-fire testing campaign the followingparameterswereusedfortheventuriflowcalibration,

• MW=38.625g/mol

• γ=1.3399

• Rg=215.26J/kg−K

• Cd=1.38915

Baseline GOX hot fire test summary

As stated previously, a series of hot fire test with GOX as theoxidizerwereperformedinordertoestablishabaselineforthesmallthruster system. Figure 16 summarizes the baseline test results.Plotted are Isp, CF, c*, and themeanABS fuel regression rate.ThefuelregressionrateisplottedasafunctionofbothoxidizerGox and totalmassfluxGtotal.BecausetheachievedO/FratiofortheGOX/ABSthrusterisratherlow,Figure18(d)demonstratesclearlythattheablatedfuelmassfluxhasaconsiderableeffectontheoverallfuelregressionrate.The specific impulse and thrust coefficient curves plot values

calculated using both the sensed thrust from the load cell, and thethrustcalculatedfromchamberpressureusingthemethoddescribedintheprevioussection.TheexpectedvaluescalculatedfromCEA24 assuming100%combustionefficiencyandfrozenflowatthenozzlethroat.arealsooverlaidontheareIsp, CF and c*plots.Theplotteddataaregenerallysupportedbythetheoreticalcalculations.Additionally,thevaluesshownbyFigure16agreewithresultspreviouslypublishedby refs. Refs,7,8 and these results support the hypothesis that the reassembled test article and test standwas returned to its previousstateofperformance,forwhichthereisanextensivedatabase.

Nytrox87 hot fire test summary

Figure17summarizestheresultsofthe10Nytroxhotfiretests.TheNytroxmixtures used for these tests are described previously,with themixtureproperties listedbyTable2.Aswith thepreviousplotofthebaselinedata,Figure17plotsIsp, CF, c*,andthemeanABSfuelregressionrateplottedasafunctionofbothGox and Gtotal.Notethat,duetothehigherO/Fratio,theablatedfuelmassfluxhasaloweroverall influence upon the fuel regression rate. The correspondingCEA curves assuming aNytrox 87 (87%N2O) liquid compositionare also plotted. Here there is significantly more scatter exhibitedbythedata,alikelyresultofthemassflowuncertaintyascalculatedby the venturi flow meter, and the variability of the Nytrox fluidcomposition as the tank empties.As expected from the theoreticalcomparisonsofFigure9themeanIsp and c*valuesareapproximately10% lower, due to the reduced flame temperature associated withNytroxcombustion.Alsonotethat,whenusingNytroxasa“dropin”





replacementforGOX,themotortendstorunslightlyricherthantheO/Fvaluerequiredforoptimalperformance.SincethethrusterfuelgrainhadbeenpreviouslyoptimizedforbestO/FratiobasedonGOXas the oxidizer, this result is also not surprising.ThemeanNytrox

combustion efficiency, calculated for each burn as the ratio of themeasuredc*tothetheoreticalvalueaspredictedbyCEAwas97.25%.

Figure 16 Summary of the GOX/ABS baseline test results.

Data comparisons

This section presents a series of bar charts that compare themeanpropertiesofthethrusterasderivedfromthe6baselineand10Nytrox-evaluationhotfiretests.Thecomparisonsincludethenominalperformance parameters Isp, c*, CF, and ρ*Isp. Chamber pressure,mass flow mean oxidizer-to fuel-ratios and equivalence ratios arealsocompared.Theireffectsonfuelregressionratewillbeassessed.Finally,therequiredignitionenergy,ignitionreliability,andignitionlatencieswillbeassessedandcompared.

Performance comparisons

Thesebarchartscomparethemeanresultsfromthehotfiretestdatapresentedintheprevioussection.Figure18comparesthemeanperformanceparametersincludingIsp, c*, and CF,alongwiththemeanoperatingchamberpressureP0 andmassflowsof the thrusterusingthetwopropellantclasses.Thecomparisonsalsoshowtheerrorbarscalculatedassumingastudent-tdistributionandan95%confidencelevel.AsshownbyFigures18(a)&18(b)theGOX/ABSdataexhibithigher Isp and c* values, 224.4s and 1740m/s, respectively; thando

theNytrox/ABSdatawithIsp and c* values of 203.1s and 1547m/s, respectively.Thisresult,aspredictedbyFigure9(c),wasexpected.AsshownbyFigure18(c)bothpropellantcombinationexhibitnearlyidenticalthrustcoefficientvalues,indicatingthatthenozzlegeometryiswelloptimizedforbothtestseries.

Figure18(d)comparesthemeancombustionchamberpressures.The GOX/ABS thruster system was originally tuned to achieve adesignchamberpressurewasapproximately830kPa(120psi).TheNytrox/ABSthrusterwastunedtogivethesamethrustandchamberpressurelevelastheGOX/ABSbymanuallyadjustingtheupstreamregular’smanualoutputsettingduringtheinitialtestfiring,andbeforetheevaluationburnseries.Thus,boththrustersystemsachievedthedesignpressurelevelandthrust(10N)ratherprecisely,indicatinglittlenozzle erosionduring the testing campaign.This chamber pressureequivalenceisalsoreflectedbytheCFbar-chartsofFigure18(c).

AsshownbyFigures18(c),18(d)&18(e), it isalso interestingto note that although both systems obtained nearly equivalentmechanical performance levels in terms of thrust and chamberpressure,theoxidizerandtotalmassflowlevelsareentirelydifferent.





The nytrox flow through the injector is approximately 25% higherthan theGOXflow for the same pressure level;whereas, the totalmassflowfortheNytrox/ABSisonly10%higher.Thisproportionallylowertotalmassflow,andoverallthrustlevelisprimarilyduetothe

lower combustion temperature, and the associated lower c* of the Nytrox/ABSpropellants,ascomparedtoGOX/ABS.ThisresultwaspreviouslypredictedbytheCEAdatapresentedinFigure9(b),andisreflectedbythebar-graphsofFigure20(b).

Figure 17 Summary of the Nytrox/ABS hot fire test results.

Figure 19 compares the mean density specific impulse values.In contrast to the mass-based Isp comparisons, the Nytrox/ABS propellantsexhibitahigherdensityspecificimpulse,approximately1080 N-s/litercomparedtothe1036 N-s/liter for GOX/ABSbaseline.This calculation is based upon the oxidizer storage density, andnot the downstream flow density. For this calculation the GOX isassumedtobestoredat2000psig,andat288ºC,withtheresultingstorage density of approximately 0.185g/cm3. The Nytrox storagedensityistakenfromtheaverageofthetwobatchesfromTable2orapproximately0.423g/cm3.Also,notethatthisrelativelylowNytroxstoragedensityresultsfromtheNOSruntankonlybeinghalf-filled(5lbs.)withnitrousoxideduringprocessing.Ifthetankwerefilledto3/4thcapacityof7.5lbs.(3.4kg),thentheresultingstoragedensityclimbs to approximately,0.562 g/cm3, and the ρ*Isp value jumps to1300N-s/liter,and improvementofmorenearly23%.Figure19(b)showsthisresult.

Oxidizer-to-fuel and equivalence ratio comparisons

Generally,ABS burned as a hybrid rocket fuel tends to have ahigher overall performancewhenburned at an equivalence ratioF,thestoichiometricO/FdividedbytheactualO/F,whichliesbetween1.5 and 2.0. Burning at these fuel-rich equivalence ratios also has

beneficialeffectofreducingtheflametemperature,loweringnozzlethroaterosion,andproducingaplumewithalowermolecularweightcomposition.

The bar charts of Figure 20 compare themeanGOX/ABS andNytrox87/ABSO/FandEquivalence Ratios fromthehot-firetestingcampaign.Figure20(a)plotsthemeanO/FratioforbothGOX/ABSandNytrox/ABS.Figure20(a)alsoplotsthetheoreticalstoichiometricO/F ratios. Figure 20(b) plots the equivalence ratios. Both fuelsburn slightly rich,with theNytrox/ABS burning richest.Note thatalthoughtheO/FratioincreasessignificantlywiththeintroductionofthehigherdensityNytroxastheoxidizer,theequivalenceratiosforthe two propellant combinations remain nearly constant,with onlyaslightincreaseinΦ forNytrox/ABS.TheGOX/ABSthrusterwasoptimizedtoburnatanequivalenceratioofapproximately1.75,andthe resultingO/F ratio tends to lie righton topof theO/F ratio formaximumc*.Figure20(b)showsthisresult.WhenGOXisswapped-outforNYtrox87;however,theresultingO/Fratiodropstoavalueslightlylowerthantheoptimallevel.Figure20(b)showstheresult.WhenburnedwithNytrox,enlargingtheinitialportdiameteroftheABSgrainwouldshifttheresultingO/Fratiobacktoneartheoptimalvalue, and presents a simplemethod for optimizing the fuel graingeometry.





Figure 18 Comparing the performance of the test thruster using GOX/ABS and Nytrox87/ABS.

Figure 19 O/F and equivalence ratio comparisons.





Figure 20 Comparing mean O/F and equivalence ratios for GOX/ABS and Nytrox87/ABS tests.

Fuel regression rate comparisons

Figure21comparesthemeanregressionratesforGOX/ABSandNytrox87/ABS.The Bar chart of Figure 21(a) compares themeanregressionratescalculatedbyfittingthemeanregressionratedataofFigures16(e)&17(e)withanexponentialcurvefitoftheform,

noxr a G= ⋅ (25)

andintegratingovertheoperatingmassfluxrange,inthiscase0-12 g/cm2-s,

max

0max 1

Goxn nox ox ox

ox

a ar G dG GG n

= ⋅ = ⋅+∫

(26)

Note that the GOX/ABS data exhibits a significantly higherregression rate then does Nytrox/ABS. Because essentially nodatabaseexistsfortheregressionratesoffuelsburnedwithNytrox,thedataofFigures16(e)17(e)arealsooverlaidwithregressionratepreviouslycollectedbytheAuthorforaprevioustestcampaign(Ref.[3]) at the PRL using Nitrous Oxide/HTPB and Nitrous/ABS aspropellants.Alsoplottedarepreviously-derived44 exponential curve fits for a) Liquid oxygen (LOX)/Paraffin, b) LOX/HTPB, c) N2O/Escorez-HTPB, and d) LOX/High-density-polyethylene (HDPE).Table 3 summarizes the exponential fit parameters for the plottedcurves on Figure 21(b).Although the regression rates achieved byGOX/ABSandLOX/Paraffinareclearlyhigher than the remainderof the propellant combinations; theNytrox87/ABS regression ratesare still clearly superior than N2O as a stand-alone oxidizer, and is afactorthatwillleadtoasmallerrequiredfuelgraindiameter.Thisvolumetric reductionwill partially offset the loss in overall systemdensitythatderivesfromtheswitchtoNytroxfromN2O.

Ignition reliability, latency, and required energy

As discussed previously, the arc-ignition system using GOX/ABSaspropellantshasbeensufficientlymaturedtoapower-efficient

system that be started, stopped and restartedwith a highdegreeofreliability.Oneof thekeyobjectivesof thisresearchcampaignwasto demonstrate that Nytrox, can be “dropped in” to the system tosignificantly increase the volumetric efficiency of the propellants,while still allowing for a high degree of ignition reliability, withgoodenergyefficiency.Clearly,thearc-ignitionsystemwaseffectiveinignitingtheNytrox/ABSpropellants;however,basedonthedatafrom the early testing campaign, the conclusions in thismatter aresomewhatmixed.

First,whenavirginfuelgrainisfirstburned,theignitionreliabilityis less than 50% and pre-lead gaseous oxygen into the pipingupstreamoftherunvalvewasrequiredtoensureignition.Oncethefirstignitionisachieved,thenthesystemreliablyigniteswithnoGOXpre-lead,evenwithadead-coldsystem.Thereasonsforthisbehaviorare still unclear at thispoint in thedevelopment campaign,but theauthors conjecture that the Nytrox expansion into the combustionchambersuper-chillsthefuelgraintothepointthattheABSmaterialimpedancerisestoapointwheretheHVPScannotprovidesufficientpowertopyrolizethefuelgrain.Onceaconductionpathissetintothefuelmaterialaftertheinitialburn,thenthisissuegoesaway.Cleary,furtherresearchisrequiredinordertoclearlydeterminethesourceofthisbehavior,andtoestablishproceduresormethodsthatmitigatethisdifficultyforoperationalsystems.

Second, theNytrox/ABSpropellants exhibit significantlyhigherignitionlatenciescomparedtoGOX/ABS.Figure22comparestimehistoriesdemonstrating the typical ignitionbehavior fora2-secondpulseofthethrustersystemusingfirstGOX/ABS,andthenNytrox87/ABS.Plottedare(a)Thrustmeasuredbythecalibratedloadcell,(b)ChamberPressure,(c)OxidizermassflowmeasuredbytheVenturiFlowmeter,and(d)theOutputignitionpowerfromtheHVPS.Notethat forbothruns the ignitionpowerstarts1 full-secondbefore theoxidizer run valve opens, and the ignition power profiles for bothconditions are quite similar. After opening the oxidizer flow isimmediate,andtheGOX/ABSmotorlightsandreachesfulloperating





pressure within 200 milliseconds.However, theNytrox/ABSMotorexhibits a more significant latency, with an additional latency ofapproximately250 ms required toreachfullchamberpressure.Theauthors believe that this latency is a result of super-chilled nytroxenteringthecombustionchamber,causedbytherapidexpansionand

phase change of the entering fluid.Not until the chamber pressurebuilduptomorethantwoatmospheresdoesfullignitionoccurs.Thistwo-atmosphere prerequisite for ignition was previously noted byRef.1

Figure 21 Regression rate comparisons.

Figure 22 Comparisons of typical ignition response time histories for GOX/ABS and Nytrox87/ABS thrusters.

Table 3 Exponential fit parameters for GOX/ABS, Nytrox87/BS and other selected propellant regression rates

Propellant GOX/ABS Nytrox87/ABS N2O/ABS N2O/HTPB

Paraffin/LOX

LOX/HTPB

Lox/HTPB-Escorez LOX/HDPE

a coefficient 0.0522 0.04206 0.00742 0.00795 0.0488 0.0146 0.0099 0.0098

n exponent 0.491 0.387 0.799 0.773 0.491 0.681 0.68 0.62





The latency exhibited by the Nitrox/ABS thruster occurred fortheentiretestruns,andappearstobeendemictotheuseofNytroxasahybridoxidizer.Thislatencycanmostlikelyonlybeovercomebytheintroductionofacatalystmaterialtojumpstartdecompositionprocessandwarmtheflowupstreamoftheinjector.Whitmoreetal.,45 havepreviouslydemonstratedtheabilitytosignificantlyreducearc-ignition combustion latencieswith anupstreamcatalyst using90%hydrogenperoxide andABSpropellants.Nytrox response latencieshavenotfullycharacterizedatthispointbeen,andarerecommendedasapointofemphasisforfurtherstudy.

ThebarchartofFigure23comparestherequiredignitionenergiesforGOX/ABSandNytrox 87/ABS.Thesedatawerecalculatedfromthepreviouslydescribeddatapresented inFigures16and17.Bothpropellantexhibitaverylowrequiredignitionenergy,andinspiteofthepreviouslydescribedignitionlatency,theNytrox/ABSpropellantsdo not appear to need a statistically significantly great energy forstartup.Bothsystemshaveameanstartupenergylessthan3.5joules,and to a 95% confidence level, neither requiremore than 5 joulesfor ignition. This energy level is contrasted to the ECAPS Prismaspacecraft46whichusedtheADN-basedLMP-103sgreenpropellant.Forfirstignitionthissystemrequireda10wattpreheatforaslongas20,consumingmorethan12,000joulesofenergy.

Figure 23 Comparing the required ignition energies for GOX/ABS and Nytrox87/ABS.

Effect of storage pressure and temperature on ignition latency

Inatypicaltest-procedurethetankwaschilledtozeroCinanicebathwhilefilling,andthengraduallywarmedtoambienttemperatureduring testing.An interesting result of this test procedure is that aclearcorrelationoftheignitionlatencytothetankstoragetemperatureand pressure appeared. Figure 24 shows this result by plotting the

ignition latency, calculated as the 63.2% first-order response risetime,plottedagainsttheinternaltanktemperature.Thetrendisverydistinctivewiththelatencygrowingfromonly50msecattheoriginalpressureandtemperature,togreaterthan1secondasthetankwarmedtoroomtemperature.

Itappears that,as the tankwarmsand thepressure levelgrows,thenthesaturatedoxygeninsolutionbeginstoprecipitateout.Asthetankwarms, the liquidNytroxstatehas lesserand lesseroxygen insolution,makingignitionincreasinglymoredifficulttoignite.Thus,maintaining a cold tank temperature duringfiring appears to be anessentialelementtoreducingignitionlatency

TheresultsplottedbyFigure25supportthisassertion.Figureplotstheoxygenmass-fractioninthenitroxsolutionusingthepreviouslydiscussedPeng-RobinsonModel.35Thiscalculationassumesthatthetankisfilledat0ºCand900psigaugepressure,withO2 percolated to fullsaturation.Thetankissubsequentlywarmed,allowingtheinternalpressuretorise,andalsochangingthemassproportionsofO2 in the vapor and liquid phases.Theoxygen content in the liquid solutiondrops precipitously until at around 1100 psig internal pressure, theliquidphasebecome100%nitrousoxide.Thisresultwasdiscoveredshortly before the publicationof this paper, andhas not been fullyexplored.Clearly,adedicatedexperimentwithadirectmeasurementof the internal tankfluid temperature is required to reduce thedatascatter.Widerpressureandtemperatureranges,andinitialsaturationlevelsshouldalsobeexplored.Whenfullycomprehended,itappearsthat thismethodmay result in engineering practices that reduce orfully eliminate the observed ignition latencies using the Nytroxoxidizer.

Extrapolating the specific impulse to vacuum conditions for Nytrox 87/ABS

Recallthatthespecificimpulseanddensity-impulsevaluesplottedon Figure 18were derived from data collected under ambient testconditions at approximately 4700 ft. (1,430 meters) altitude, theelevationofthePRLtestfacilityinLoganUtah.The2.07expansion-rationozzlewasroughlydesigntogiveoptimalperformanceatthisaltitude.Clearly,whenmatchedwithahighexpansion-rationozzle,the vacuum performancewill be significantly better.This data canbe extrapolated to altitude by using the previously presented 1-DdeLavalflowequationsfromEqs. (16)-(24).Using thismodel, thespecificimpulseunderoptimalconditionscanbewrittenintermsoftheoptimalthrustcoefficientandthenozzleexit-to-chamberpressureratio.Theresultforthrustcoefficientis

( )11

1

*0 0

2 2 11 1

exit exit exitFvac test

p A p pCP PA

γγγγ

γγ γ

−+− ∞ −

= ⋅ ⋅ − + − +

( )

111

0

2 2 11 1Fvac opt

pCP

γγγγ

γγ γ

−+− ∞

= ⋅ ⋅ − − + (27)

UsingCFtoscalespecificimpulse,

( )( )

( )( )

( )

( ) ( )

111

*0 00

* 1101

0 *0 0

2 2 11 1

2 2 11 1

vac

Foptsp opt

sp test Ftest exitvac exit test

test

pP A PCI g mP AI C p pp Ag m

P PA

γγγγ

γγγγ

γγ γ

γγ γ

−+− ∞

−+− ∞∞

⋅ ⋅ − ⋅ − + ⋅ = =⋅

− ⋅ ⋅ ⋅ − + ⋅ − +

(28)





Figure 24 Thruster ignition latency correlated with internal Nytrox tank pressure.

Figure 25 Peng-Robinson model prediction of O2 mass faction in Nytrox liquid phase as function of tank temperature and pressure.

Using the motor parameters, thrust coefficient, mean chamberpressure,andtheCEA-derivedThermochemistry

ParametersforNytrox 87/ABS,andassuminga97.3%combustionefficiency (from Section O., Figure 26 plots this extrapolation forbothGOXandNytroxoxidizers.Plotted are (a)optimal expansionratio for the Nytrox 87/ABSmotorasafunction,(b)optimalCFasafunctionofexpansionratio,(c)optimalspecificimpulseasafunctionofexpansionratio,and(4)optimalspecificimpulseasafunctionofaltitude.AlsoplottedastheredsymbolsaretheactualvaluesfortheNytrox 87/ABSmotor.Noteatanexpansionratioof50,correspondingtoanaltitudeof29km(95,000ft.)theoptimalCFexceeds1.8andtheoptimalIspreachesavalueofapproximately295 s.ThisIsp value is more than25%higher thancanbeachievedbyanyof the“green”ionic liquid propellants or mono-propellant hydrazine. Using the295s Isp value to extrapolate the ρ*IspfromFigure21(b),theprojectedvacuumvaluerisestoapproximately2560 N-s/liter.

With more than 10 successful ground test burns, reliable on-demandignitionshavebeendemonstratedusingNytrox 87 as a drop-in replacement for GOX in the USUArc-ignition thruster system.Specific impulsevaluesexceeding200 s under labconditionshavebeendemonstrated.ThisIsp level extrapolates to nearly 300 s under vacuum conditions with an optimized nozzle. This Isp value far exceedsspecificimpulsevaluesachievedbyotheravailablein-spacemonopropellants including hydrazine, LMP-103S, andAFM315-E.Table 4 compares the performance of the peroxide/ABS system tohydrazine,LMP-103SandAFM315-E.47Withtheexceptionofρ*Isp, theNytrox/ABS systemoutperforms the other propellants in everymeasureablecategory.However,eventhis lowerdensityIsp value is misleading.BecauseNytroxhadtheabilitytosafelyself-pressurize,thereisnoneedforanadditionalvolumetricallyinefficienttosafelyself-pressurize, there is no need for an additional volumetricallyinefficient oxidizer pressurization system. Thus, even in termsof volumetric efficiency, Nytrox 87/ABS appears to have a clearadvantage.





Figure 26 Extrapolating test specific impulse and thrust coefficient to optimal high altitude conditions.

Table 4 Comparison of Nytrox/ABS performance characteristics to existing space mono-propellants8

Propellant Hydrazine LMP-103S AF-M315E Nytrox/ABS Hybrid

Flame Temperature 600-750ºC 1600ºC 1900ºC 3000ºC1

Isp, s 220-225252 (theory), 266 (theory) 300 (theory)

235 (delivered) 245 (delivered) 294 (delivered)2

Specific Gravity 1.01 1.24 1.465 0.650 (87% N2O)

Density Impulse, N-s/liter 2270

3125 (theory) 3900(theory) 2560 (vacuum, extrapolated)

2915 (delivered) 3650 (delivered) 1900 (ambient, delivered)

Preheat Temperature 315ºC, cold-start capable 300ºC 370ºC N/A none-required

Required Ignition Input Energy, Joules N/A

12,000 J (10 Watts @ 1200 seconds)

27,000 J (15 Watts @ 1800 seconds

2-5 J (4-10 Watts for 500 msec)

Propellant Freezing Temperature

1-2ºC -7ºC < 0ºC (forms glass, no freezing point) -70ºC

Cost $ $$$ $$$$ $

Availability Readily Available Restricted Access Limited Access Very Widely Available3

NFPA 704 Hazard Class 5

1Dataforhydrazine,LMP-103SandAFM315-EweretakenfromRef.[7].2DuetothehighpyrolysisenergyofABSfuel,3.1MJ/kg,motorsareablativeandself-cooling.3Extrapolatedtovacuumconditionsbasedongroundtestdata.480-90%N2Osolutionseasilymanufactured,asperprocedureinthispaper.5Baseduptheconstituentcomponents,HydroxylAmmoniumNitrate(HAN)and2-Hydroxyethylhydrazine(HEHN)





Summary and conclusionThePropulsionResearchLaboratory(PRL)atUtahStateUniversity

recently developed a promising “green” propulsion alternative thathasthepotentialtoreplacehydrazineformultipleapplications.Theuniquehybridpropulsiontechnologyderivesfromthenovelelectricalbreakdownpropertiesof3-DprintedABSdiscoveredserendipitouslywhile investigating the thermodynamic performance of ABS as ahybrid rocket fuel.This concept has been developed into a power-efficientsystemthatcanbestartedandrestartedwithahighdegreeofreliability.Multipleprototypeground-testunitswiththrustlevelsvaryingfrom4.5Nto900Nhavebeendevelopedandtested.

Thispaper investigates theuseofamedicalgradenitrousoxideandgaseous oxygenfluid blend, “Nytrox,” as an intrinsically saferalternativetobothgaseousoxygenandpurenitrousoxideasahybridrocketoxidizer.Inamannerdirectlyanalogoustothecreationofsoda-water using dissolved carbon oxide,Nytrox is created by bubblinggaseousoxygenunderhighpressureintonitrousoxideuntilthesolutionreachessaturationlevel.Oxygeninmixtureullagedilutesthenitrousoxide vapor, and increases the required decomposition activationenergy of the fluid by several orders of magnitude. Consequently,anyriskofinadvertentthermalorcatalyticdecompositionisvirtuallyeliminated.This paper reports on a preliminary test-and-evaluationcampaignwhere an existing small spacecraft thruster is first testedusing gaseous oxygen and 3-D printed ABS as propellants as abaseline.

Asimplepercolationprocedureformanufacturingan87%Nytroxsolutionusinghigh-purity,medicalgradenitrousoxideandgaseousoxygenwasdevelopedand reported.At this reporting two6.5 literbatchesofNytroxhavebeenpreparedusingthereportedprocedure.Both batches were remarkably consistent, agreeing well withtheoreticalpredictionsfordensityandsaturationpressure.Withmorethan 10 successful ground test burns, reliable on-demand ignitionshave been demonstrated usingNytrox 87 as a drop-in replacementforGOX in theUSUArc-ignition thruster system.Clearly, the arcignitionsystemwaseffectiveinignitingtheNytrox/ABSpropellants;withnostatisticallysignificantincreaseintheoverallignitionenergyrequirements. However, based on the data from the early testingcampaign,theconclusionsinthismatteraresomewhatmixed.

First,whenavirginfuelgrainisfirstburned,theignitionreliabilityis less than 50% and pre-lead gaseous oxygen into the pipingupstreamoftherunvalvewasrequiredtoensureignition.Oncethefirstignitionisachieved,thenthesystemreliablyigniteswithnoGOXpre-lead,evenwithadead-coldsystem.Thereasonsforthisbehaviorare still unclear at thispoint in thedevelopment campaign,but theauthors conjecture that the Nytrox expansion into the combustionchambersuper-chillsthefuelgraintothepointthattheABSmaterialimpedancerisestoapointwheretheHVPScannotprovidesufficientpowertopyrolizethefuelgrain.Onceaconductionpathissetintothefuelmaterialaftertheinitialburn,thenthisissuegoesaway.

Second, theNytrox/ABSpropellants exhibit significantlyhigherignitionlatenciescomparedtoGOX/ABS.However,theNytrox/ABSMotorexhibitsamoresignificantlatency,withanadditionallatencyofapproximately250 msrequiredtoreachfullchamberpressure.Theauthors believe that this latency is a result of super-chilled nytroxenteringthecombustionchamber,causedbytherapidexpansionand

phase change of the entering fluid.Not until the chamber pressurebuild up to more than two atmospheres does full ignition occurs.The latency exhibited by theNitrox/ABS thruster occurred for theentiretestruns,andappearstobeendemictotheuseofNytroxasahybridoxidizer.Thislatencycanmostlikelyonlybeovercomebytheintroductionofacatalystmaterialtojumpstartdecompositionprocessandwarmtheflowupstreamoftheinjector.Nytroxresponselatencieshavenotfullycharacterizedatthispointbeen,andarerecommendedasapointofemphasisforfurtherstudy.

Itappears that,as the tankwarmsand thepressure levelgrows,then the saturated oxygen in solution begins to precipitate out.Asthetankwarms,theliquidNytroxstatehaslesserandlesseroxygenin solution, making ignition increasingly more difficult to ignite.Thus, maintaining a cold tank temperature during firing appearsto be an essential element to reducing ignition latency.This resultwasdiscoveredshortlybeforethepublicationofthispaper,andhasnot been fully explored. This discovery may result in engineeringpracticesthatreduceorfullyeliminatetheobservedignitionlatenciesusingtheNytroxoxidizer.

Finally, specific impulse values exceeding 200 s under labconditions have been demonstrated. This Isp level extrapolates to nearly 300 s under vacuum conditions with an optimized nozzle.ThisIsp valuefarexceedsspecificimpulsevaluesachievedbyotheravailablein-spacemonopropellantsincludinghydrazine,LMP-103S,

andAFM315-E.With theexceptionofdensity specific impulse,theNytrox/ABS systemoutperforms the other propellants in everymeasureablecategory.However,eventhis lowerdensityIsp value is misleading;becauseNytroxhadtheabilitytosafelyself-pressurize,thereisnoneedforanadditionalvolumetricallyinefficientoxidizerpressurizationsystem.Thus,evenintermsofvolumetricefficiency,Nytrox 87/ABS appears to have a clear advantage. When fullydeveloped, a Nytrox-based hybrid system offers high value to thecommercial space industry by offering a high-performing, butinherentlysafe,spacepropulsionoptionforridesharepayloads.

PatentsWhitmore SA. Restartable Ignition Devices, Systems, and

MethodsThereof.USAProvisionalPatentNo.US2015/A0322892A1,Nov.12,2015.

FundingThisworkwaspartiallyfundedwithacooperativeagreementwith

theNASAMarshallSpaceflightCenter,CooperativeAgreementNo.NNM16AA01A.

AcknowledgmentsThe author is especially grateful for the assistance of NASA

MarshallSpaceFlightCenter(MSFC)ER-23,bygraciouslyprovidedaccesstothetestingfacilitiesusedtocollectvacuumchamberdataforthisproject. IdeeplyappreciateMSFCemployeesKevinPedersen,CarlosDiaz,andDanielCavender for their time, technicalsupport,andexpertadvice.

Conflicts of interestTheauthorsdeclarenoconflictofinterest.





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