Liquid Fuels and Propellants for Aerospace Propulsion.pdf

7/21/2019 Liquid Fuels and Propellants for Aerospace Propulsion.pdf

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JOURNAL OFPROPULSION ANDPOWERVol. 19, No. 6, NovemberDecember 2003

Liquid Fuels and Propellants for Aerospace Propulsion:19032003

Tim Edwards

U.S. Air Force Research Laboratory, WrightPatterson Air Force Base, Ohio 45433-7103

Introduction

M AJOR increases in liquid-fueled propulsion performancehaveoccurredin thepast 100years.The Wrightbrothersrstew on 14 December 1903 with an engine that generated slightlymore than 130 lb of thrust for ights that ranged from 120 to 850 ft(Ref. 1). Contrast that with currentaircraft such as the Boeing 777,which can y 8000 n miles, equipped with engines from the GE90and PW4000 series that generate over 100,000 lb of thrust. TheWright brothers probably consumed less than a gallon of gasolinein that rst day of ight tests.In 1997,airlines consumedan averageof 177 million gal/day of jet fuel worldwide.2 Robert Goddardsrst ight of a liquidrocketon 16 March 1926 reachedan altitudeof41 ft and landed184ft fromthe launchpoint(seeRef. 3). Again,thecontrast with current vehicles such as the space shuttle, which canlift over 50,000 lb to 200C miles into low Earth orbit,4 is dramatic.The intent of this paper is to describe the evolution of liquid fuelsfor aircraft and rockets as the engines and vehicles they fuel haveundergone these signicant increases in performance.

Engine Figures of Merit

To understand the evolution of fuels and propellants, it is help-ful to introduce the fuel properties that drive engine performance.For reciprocating, for example, Otto cycle, engines, the key en-gine parameters are horsepower/weight ratio (primarily driven by

compression ratio) and specic fuel consumption (fuel mass owdivided by engine thrust or horsepower). As described later, themaximum compression ratio is typically limited by the tendencyof aviation gasoline to autoignite prematurely, as characterized byoctane number. For gas turbine (Brayton cycle) engines, the mostimportantengineparametersare thrust/weight ratioand specic fuelconsumption. The fuel parameters most relevant to gas turbine en-gine and vehicle performanceare the heat of combustion on a massbasis in air and the density. Through the well-known Breguet rangeequation, these two parameters directly affect the range of the air-craft. For aircrafttoday, the cost of the fuel is also a key parameter.There are many other fuel parameters that are important to reliableengineoperationrelatingto fuel composition,volatility,combustionperformance, stability, and contaminants. The role of these proper-

ties in gas turbine engine performance is a key part of the historicaldiscussion to follow.

Tim Edwards is the Senior Scientist of the Fuels Branch in the U.S. Air Force Research Laboratorys PropulsionDirectorate/Turbine Engine Division. He has 20 years of experience in research in fuels and propellants, workingat both of the Propulsion Directorates sites at Edwards Air Force Base (no relation) and WrightPatterson

Air Force Base (WPAFB). His interests include properties and applications of hydrocarbon fuels/propellants andendothermicfuels, in supportof advancedpropulsionprogramsfor the Departmentof Defense and NASA. Edwards

manages the in-house basic research program in advanced fuels, sponsored by the Air Force Ofce of ScienticResearch and assists in coordinating the on-site research occurring at the National Aerospace Fuels ResearchComplex at WPAFB for the Versatile Affordable Advanced Turbine Engine program. His early work was focused

on spectroscopicmeasurementsof solidpropellantames at pressure. He could be considered a native of WPAFB,having been born in the baseh ospital while his father worked in the Fuels and Propellants Branch there in the late1950s. He received a Ph.D. in chemical engineering from the University of California at Berkeley in 19 83. He is an

Associate Fellow of AIAA.

Received 17 April 2003;revisionreceived 15 August 2003;accepted for publication29 August 2003.This material is declared a work of the U.S.Governmentand is not subject to copyright protection in the United States. Copies of this paper may be made for personal or internal use, on condition that the copierpay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 0748-4658/03 $10.00 incorrespondence with the CCC.Data also available at http://www.chevron.com/prodserv/fuels/bulletin/.

For liquid propellant rockets, the key propellant-related param-eters are specic impulse Isp (thrust divided by propellant massow rate) and propellant density. Of course the key difference be-tween aircraft enginesand rocket engines is the rocket requirementto carryan oxidizeronboardthe vehicle. The propellant-drivenpartsof the Isp parameter are the propellant(fuel plus oxidizer) heats offormation and the propellant stoichiometry.5;6 In general,Isp is pro-portional to the square root of the ratio of the ame temperature tothe combustion-productmolecular weight, so thatIsp is maximizedby high ame temperatures and low-molecular weight combustionproducts. A key parameter for both rockets and airbreathing mis-

siles is the propellant density. For a given application, the gure ofmerit for the propellants is Isp density

a , where a is an exponentthat variesfor differentapplications,but is typicallybetween0.2 and1 (Ref. 7). In general, density is most importantfor the rst stage ofmultistage vehicles. Although the cost of the propellants is a smallfraction of the launch cost, the operationscost and ease of handlingof propellantsis a key concern, particularly for cryogenic and toxicstorable propellants. For ballistic missile applications (where therocket must remain fueled for long-periodsof time), noncryogenic(storable) propellants are used to avoid the difculties of long-termstorage of cryogenic propellants.

Brief Note on Terminology

Fuels based on petroleum distillates have incorporated a host ofsomewhatnebuloustermsbasedintheterminologyoftheearlyren-ing industry. Early petroleum reneries were primarily distilleries,with products (fractions) differentiated by boiling point. Thus, areader may come across unfamiliar terms such as kerosene, naph-tha, gas oil, heavy or light fraction, sweet or sour crude, straight-run, hydrotreated, etc. Table 1 is an attempt to list the commonterms for petroleum fractions.2;811 There is considerable overlapbetween categories,as well as considerable disagreement about thedenition of a particular generic category. For example, one refer-ence gives a denition of kerosene as a rened petroleum distillatethat has a ash point of 25 C (77 F), 11 in contrast to the boilingrange denitions in Table 1. In general, as the boiling temperatureincreases, the molecular weight and density increase and the va-

por pressure decreases. Thus, one might describe the upper boilingrange of a fraction as the heavy or residual end of the fraction.

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Table 1 Petroleum distillate terminology 2;811

Approximate Typical average Most prevalentFractio ns boil in g ran ge, C carbon number carbon number

Generic termsGasoline 275

Specic products

Avgas 45145 C7 C8Motor gasoline 30200 C7 C5

(mogas)Auto diesel 200350 C16 C16Jet fuel 150265 C11 C11

Fig. 1 Carbon numberdistributionfor varioushydrocarbonfuelsand

propellants.

Low- (mercaptan) sulfur-content products are described as sweet

(in contrastto high-sulfurproducts being described as sour). A fuelthat is created primarily by distillation is straight run, whereas afuel subjected to the two most common treatments to reduce sulfurand other impurities would be called either hydrotreated or Meroxtreated. The original product of the petroleumreningindustry waslamp fuel; hence, Whittles use of illuminating kerosene in his gasturbine engine.Aviationgasolineobviouslyfallsin the generic gaso-line fraction, whereas most jet fuels and liquid hydrocarbonpropel-lants can be described genericallyas kerosene.These denitions areshown in Fig. 1, where the carbon number distribution for aviationgasoline (avgas), jet fuel (Jet A), and rocket fuel (RP-1) is illus-trated. The carbon number distribution in Fig. 1 is interpreted as,for example, avgas consisting of over 50% (by weight) moleculesthat contain 8 carbon atoms. Auto diesel fuel falls mostly into the

kerosene range, whereas some heavier, for example, marine, dieselfuels can be describedas gas/fuel oil or residual fuel.Currently, theUnited States consumes roughly twice as much gasoline as dieseland jet fuel combined,whereas other parts of theworld have a moreeven distribution.

Airbreathing Aviation Fuels

This discussion on aviation fuels relies very heavily on severalsignicant references.1216 The interestedreader is stronglyencour-agedto seek out these excellentdocuments,and the referencescitedtherein,for additionaldetailsthatare beyond the scopeof thispaper.

Aviation gasoline (avgas)

In 1903, the Wright brothers ew, quite naturally, on motor gaso-

line. They built their own engine, one of many innovations in theiraircraft. Their four-cylinder water-cooled engine weighed 180 lb(83 kg) and developed 12 hp (9 kW). They used several cans ofStandard Oil motorgasolinefrom a nearbyboatyard.12 This enginewas quite an advance over the steam engines used by some of theircompetitors.Assubsequentaircraftenginesevolved,the tendencyof

Fig. 2 Classes and examples of hydrocarbons found in hydrocar-bon fuels; research octane numbers: isooctane= 100, n-heptane= 0,

1-pentene = 77, methyl cyclohexane (MCH) = 71, and benzene= 123.

the gasoline to autoignite ahead of the spark (preignition) or beforethe arrival of the ame front (knocking, pinging) was a key limiting

factor in aircraft engineperformance.This is the same knockingthatoccursin automobile engines today, but it is much more destructiveand dangerous in aircraft engines. During the World War I period,various researchers(notablyRicardos group in England) identiedthis autoignition phenomenon and began to identify the fuels char-acteristics that controlled it. The hydrocarbon constituentsof fuelscan be divided into ve classes of compounds (Fig. 2): parafns,isoparafns, cycloparafns (naphthenes), olens, and aromatics. Itwas found thataromaticssuch as benzene(benzolin older literaturewas a crude mixture of aromatics consisting primarily of benzene)were less prone to knocking and thus, avgas produced from crudeoilsrich in aromaticswere betterperformersthan fuelsthat were pri-marily parafnic. These differences became evident as the UnitedStates entered World War I and began supplying gasolines from

quite different sources than typically encountered in Europe. Themore-parafnic U.S. gasolines performed relatively poorly in theadvanced European engines such as the French Gnome. A key de-velopment in understanding avgas performance came in the 1920swith the standardizationof test methods and enginesto determine athe tendency of a fuel to knock. This standardizationwork was ledby the Cooperative Fuel Research (CFR) committee in the UnitedStates, composed of representativesfrom the engineand petroleumrening industries. The tendency of a fuel to knock was character-ized by its octane rating, developed by Edgar, which compared thetendencyof a givengasolineto knockto that of a referencefuel com-posed of a mixture of n-heptane(octanenumberD 0) and isooctane(octane numberD 100). On this scale, the Wright brothers 1903fuel has been estimated to be 38 octane,13;17 with their 1910 fuel

estimated as 58 octane (Ref. 18). By 1932, industry had generallyadopted a standardized method for knock rating, using an enginedesigned by H. L. Horning of the Waukesha Motor Company forthe CFR committee,and Edgars fuelstandards.Notethat the octanescale ranks fuels quite differently than the cetane scale for fuels fordieselengine applications.In a compression-ignitionengine suchasa diesel, fuels that igniteeasily are desirable. Thus, cetane (hexade-cane, C16 H34/has a cetane number of 100, whereas the isoparafnheptamethylnonane has a cetane rating of 15. Note the inversionof the scales: Aromatics and isoparafns rank high on the octanescale and low on the cetane scale, whereas the reverse is true ofnparafns.

In a reciprocating spark-ignition engine, the fuel and air aremixed, injected into a cylinder, compressed, and then ignited. The

development of higher octane avgas was driven by the engine per-formance increasesob tainablevia turbocharging/superchargingandincreasedcompressionratios.2;4 Turbochargingis accomplished bycompressing engine inlet air using a turbine driven by hot engineexhaust gas. Supercharging is the same air compression accom-plished by a shaft-mounted turbine. Turbocharging/supercharging


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dramatically increase the maximum speed and altitude that can beachieved, when compared to an engine where the intake air is atambient pressure. The compression ratio is the ratio of the com-bustionchamber volume at bottom dead center (maximum) and topdead center (minimum). Higher compression ratios enable smallerengines for a given power output, in other words, increasedenginepower/volume ratios. The maximum compression ratio and maxi-mum amount of inlet air pressurization are both limited by knock-ing/fuel autoignition. A tremendous advance in fuel performance

came from the investigations of Midgeley and Boyd, working forCharles Kettering in the Dayton EngineeringLaboratory (Delco) inDayton, Ohio. From an exhaustive study of knock promoters andpreventers emerged the remarkable additive tetraethyl lead (TEL)in December 1921. This additive enabled the production of largequantities of higher octane avgas. It was found that the sensitivityof avgas to lead varied widely, with aromatic avgas less sensitiveto knock retardation by lead (as compared to more parafnic gaso-lines). The problem of deposition of solid lead oxides on exhaustvalves was solved by the additionof ethylenedibromide(EDB) withthe TEL. TheEDB reactedwith theleadin thecombustionchamber,forming volatile lead bromide compounds.

Fuelreners andenginemanufacturersengagedin a gameof tech-nological leap frog in the late 1920s and 1930s, as improvements

in engines led to requirements for improved fuels (and vice versa).The rst avgas specication that included octane number perfor-mance requirements was issued by the U.S. Army for 87 octane in1930.12;14 A signicant milestone came in 1934, when the Power-plant Branch of the U.S. Army Air Corps at McCook(later Wright)Field in Dayton issued the specication X3575 for 100/130 octaneavgas.This specicationcame aboutfrom the work at McCook Field(Dayton, Ohio) led by Heron, which demonstrated the signicantgains in performancefrom higheroctane fuels. The rst number inthe specication is the lean octane rating and the second is the rich(high-power) rating. Octane ratings above 100 were characterizedby performance numbers, with the reference fuels being mixturesof isooctane and TEL. For example, 3cm3 TEL/gal of isooctaneyields a rich performance number (PN) of 145. The PN scale is

roughlylinearly related to maximum engine output power; in otherwords, a fuel with a 145 PN will yield 45% more knock-limitedoutput power than a 100 PN (D100 octane) fuel. Alternatively, anengine rated at 1500 hp on 145 PN avgas may only develop 600 hpon 80 octaneavgas.16 The 100/130 octanespecication (AN-F-28 in1942) allowedup to 3-cm3 of TEL/gal,but still required high octanebasegasoline.The productionof this fuelchallengedreners for thenext10 years andlead to thedevelopmentof alkylationandcatalyticcracking processes. The availability of 100 octane avgas from theUnitedStates is widely acknowledgedas oneof thekey contributorsto the British victory in the Battle of Britain against a numericallysuperior foe duringWorld War II.The productionof avgasincreasedfrom 54 million gal/yearin 1932 to more than 25 million gal/day atthe end of World War II. The unprecedented industry/government

cooperation needed to surmount the challenges posed by this enor-mousexpansionin rening capacityhas beendescribedby Heron.14

Thequestfor more powerfulenginesled to 108/135and 115/145 av-gas specications in the 1940s. The rather chaotic pace of gasolineand engine development led to the creation of 12 separate gradesof gasoline by 1945, often distinguished by dye.19 In that year, theAmerican Society for Testing and Materials (ASTM) codied thegrades of avgas in specication D 615: 80/87 (red) with 0.5-cm3

TEL/gal,91/96 (blue,4.6-cm3/gal), 100/130 (green ,4-cm3/gal), and115/145 (purple, 4.6-cm3/gal).

As gas turbine-powered aircraft became predominant followingWorld War II, the 100/130 and 80/87 octane specications becamethe primary ones used. Civilian airlines had generally resisted theuse of 100 octane fuel before the war because of cost factors. How-

ever, improvements in fuels and engines enabled greater altitudesof operation,which had many operationalbenets. The availabilityof improved fuels has been credited as the biggest contribution ofWorld War II to the commercial airline industry.20 Once the air-line industry adopted the gas turbine engine in the late 1950s,21 theuse of aviation gasoline was conned to the relatively small general

Fig. 3 Trends in U.S. aviation fuel consumption.2;22

aviationmarket. An illustrationof thesemarket changescan be seenin the fuel consumption numbersin Fig. 3.2;22

By the 1980s, various forces had reduced the available gradesofavgas to essentially one: 100LL (low lead), currently specied un-der (ASTM) specication D 910 as containing a maximum of 0.56g/l TEL maximum. Type certication can be obtained from the air-craft manufacturer for the use of unleadedmotor gasoline.The use

of (unleaded) automotive gasoline in aircraft is controversial.

22;23

Aside from the lead issue, there are differences in volatility, addi-tives and blending components (especiallyoxygenates),and qualitycontrol.

Thus, the developmentof high-performanceavgas in many wayspaced the development of reciprocating aircraft engines. The rela-tionship between jet fuels and gas turbine engineswas not as close,however. Even in 1949, it was anticipated that improvements ingas turbine engine performance would be much less sensitive toimprovements in fuel.14

Turbine Engine Fuels

Theearly pioneersin gas turbinedevelopment,Whittlein EnglandandVonOhainin Germany, faceda wide varietyof optionsinchoos-

ing a fuel for gas turbines. Whittle had considered diesel fuel, butended up choosing illuminating kerosene because of an expectedrequirement for a lower freeze point than that available with diesel(see Ref. 15). In contrast, Von Ohain originally demonstrated histurbine engine with hydrogen, but vehicle considerations led to aswitchto liquidfuel (see Refs. 12 and21).The worlds rst turbojet-powered ight was made on 27 August 1939 in a Heinkel 178 air-craft burning avgas. The rst ight of the Whittle engine occurredon 15 May 1941 in a Gloster Meteor aircraft using kerosene asthe fuel. Despite their head start in turbojet engine development,Germany did not decide until 1943 to produce jet-poweredaircraft.One of the arguments for development at that time was Germanysshortage of high octane fuel and that the jet engine could run ondiesel fuel.12

Most of the jet engines developedbefore the end of World War IIutilized conventional kerosene as a fuel. The rst jet fuel speci-cation was directorateof engine research and development(DERD2482),publishedinEnglandin1947.19 Asenginesandspecicationsdeveloped,it became apparentthat several fuel propertieswere keyto bounding the envelope of jet fuel characteristics. High-altitudeoperation meant fuel freeze point required attention. However, thelower the freeze point, the lower the fraction of crude oil that wassuitable, so that freezepoint had to be balanced against availability.Higher fuel volatility/vapor pressure aided vaporization -control ledengineperformancerequirementssuch as altituderelight,whichhadto be traded against boiloff and entrainment losses from fuel tanksat altitude (as well as safety concerns from explosive mixtures intank vapor spaces).24 In the UnitedStates,JP-1, JP-2, andJP-3 were

unsuccessful attempts to balance the conicting requirements ofvolatility, freeze point, a nd availability/cost.15 Two fuels emergedin the late 1940s and early 1950s from this chaotic situation: awide-cut naphtha/kerosene mixture called JP-4 in the United States(MIL-F-5624 in 1950) and a kerosene fuel with a 50C (58F)freezepoint(DERD-2494 in Englandand Jet A-1 in ASTM D-1655


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in the United States). This freeze pointwas arrived at through a sig-nicant research effort. ASTM D-1655 also specied Jet A with a40C (40F) freeze point. The Jet A-1 freeze point was changedto 47C in the late 1970s to increase availability.19 Civil avia-tion currently uses Jet A-1 (or its equivalent)throughoutthe world,except for domestic carriers in the United States, who use Jet A.Military aircraft used JP-4 until converting to JP-8 in the 1980s. JP-8 (MIL-T-83133)is essentiallyJet A-1 with threemilitary-speciedadditives (as described later). The conversion to JP-8 occurred pri-

marily to improve the safety of aircraft, although the single fuel forthe battleeld concept (and the similarity of jet fuel to diesel fuel)is centered on the use of aviation kerosene in all U.S. Air Force andU.S. Army aircraftand ground vehicles. A similar process is occur-ring in the U.S. Navy, where the large variety of liquid fuels havecompressed down to just two, JP-5 (for aircraft) and F-76 diesel forall other liquid fuel requirements.

The history of the evolution of conventional,widely availablejetfuels from the late 1950s to the present is mainly the story of theevolutionof test methods and fuel additives to maintain the integrityof the jet fuel supply and to improve safety and correct operationalproblems.Becausetheirimportance,specications/testmethodsandadditives are discussed separatelylater.

Specialty fuels were developedfor various applicationsthrough-

out the second-half of the 20th century. In the early 1950s, JP-5(includedin MIL-F-5624) was developed.JP-5 is a high-ash-point(60C/140F) aviation kerosene used onboard U.S. Navy ships toenhance safety. The development of higher Mach aircraft led toseveral specialty fuels. As ight velocity increases, aerodynamicheating leads to larger amounts of heat being rejected to the fuel,both in the tanks and in the engine, leading to vapor pressure andthermalstability concerns. The cutoff point between the use of con-ventionalJet A-1/JP-8 fuelsand speciallyproducedfuelsis betweenMach2.2 and3. Thus,the Mach 2.2ConcordeusesJet A-1, whereasthe Mach 2-3 XB-70 and SR-71 used specialty fuels. JP-6 (MIL-F-25656) was a low-volatility kerosene developed for the Mach 2CXB-70.25 The Mach 3 SR-71 required JP-7 (MIL-T-38219), a low-volatility/high thermal stability, highly processed (low sulfur and

aromatics) kerosene.15;24

The U-2 high-altitudereconnaissanceair-craftrequiredbothimprovedthermalstabilityandlowerfreezepointin its fuel (JP-TS, MIL-T-25524) because of its high-altitude, long-duration cruise.These specialty fuels gavehigher performancethanconventional aviation kerosenes, at the expense of higher fuel andlogistical costs (JP-7 and JP-TS are roughly three times the cost ofJP-8 and Jet A-1). The accepted operational temperature limits ofthese various fuels are approximately 163C (325F) for Jet A/JetA-1/JP-8/JP-5, 219C (425F) for JP-TS, and 288C (550F) forJP-7 (Ref. 26).

Russian jet fuels underwent a parallel evolution throughout thisperiod.27;28 In most areas, current Russian fuels TS-1 and RT andRussian specications (GOST 10227) are interchangeable with JetA-1/JP-8. The main difference between fuels TS-1 and RT is in the

area of thermal stability: TS-1 is a straight-run fuel, whereas RT ishydrotreated. By comparison with Jet A-1/JP-8, TS-1 and RT arelighter (have a lower initial boiling point and 10% recovery pointin distillation) and have a correspondingly lower ash point andfreeze point. Thus, worldwide there are three major specicationsin civil use: ASTM D 1655, British Defence Standard (Def Stan)91-91 (successor to DERD 2494), and GOST 10227. Internationaloil companies have created the Joint Check List to standardize jetfuel deliveriesworldwide under Jet A-1/Def Stan 91-91. The Inter-nationalAir TransportAssociationhas alsoissuedguidancematerialfor its members codifying the Jet A/Jet A-1/TS-1 specications.

Pustyrevdiscussestwo specialtyRussianfuels specied in GOST12308:T-8V, a higherdensity/higherash-pointkeroseneand T-6, ahigh-densitykerosene(specic gravity 0.84 vs 0.8 for Jet A-1/JP-8),

which has no commercial or military counterpart in Europe or theUnited States.28 U.S. Air Force programs in the 1980sdemonstratedthe production of fuels similar to T-6 (Refs. 29 and 30), but nospecication was published in the absence of user requirements.

Beyond Mach 5, ight speeds are considered hypersonic (vs su-personic). The high-heat loads encountered by vehicles in hyper-

sonic ight leads inevitablyto regenerativecooling of (at least) thecombustor. The heat loads and fuel ows are such that high lev-els of fuel heat sink are required. This heat sink can be obtainedfrom sensible heating of high-heatcapacityfuels such as liquid hy-drogen or the use of endothermic (hydrocarbon) fuels. For manyapplications, hydrocarbon fuels are preferred due to their greaterdensity31 andease of handling.Endothermic fuels achieveheat sinkby deliberatereactions of the bulk fuel, such as dehydrogenationorcracking.3235 Engineconceptsfor hypersonicaircraftare currently

under development in several programs. Potential future develop-ments in this area will be discussed in the Future Trends section.The petroleum shortages of the 1970s led to the search for do-

mestic sources of liquid transportation fuels. Large United Statesreserves of coal and oil shale (and Canadian reserves of tar sands)spurred the development of conversion processes to produce fuelsfrom thesenon-petroleumsources.In the 1980s,programs were ini-tiated to demonstratethe suitability of fuel derived from shale,3638

coal,39 and tar sands.40 Engine testing and ight demonstrationsofshale-derived JP-4 indicated no deleterious effects resulting fromthe use of shale-derivedfuel. The successof this program indicatedthat the JP-4 specication was restrictive enough to provide ade-quate fuel, regardlessof the hydrocarbonsource.Jet fuels producedfrom synthesis gas (COCH2/via FischerTropsch (FT) technol-

ogy are currently being studied for their suitability for aircraft (seeRefs. 41 and 42). The synthesis gas canbe produced from coal, nat-ural gas, or other carbon-containingmaterials. A 50/50 mixture ofpetroleum-derivedJet A-1 and isoparafnic kerosene derived fromcoal is being deliveredto aircraftin South Africa. A thoroughstudyby Southwest Research Institute demonstrated that the 50/50 mix-ture properties fell well within the Jet A-1 specication range andshould have no impact on engine operation.41;42 The compositiondifferences between the two fuels are shown in Figs. 4 and 5. Lu-bricity and elastomer compatibility issues for use of the pure FTfuel are currently being addressed.

World-widejetfuelconsumptionwas177milliongal/dayin1997,with about 40% of the consumption in the United States.2 U.S. jetfuel consumptionis predominately(90%) by commercialaircraft.

Fig. 4 Composition distribution for South African petroleum-derivedJet A-1 (Ref. 41).

Fig. 5 Composition distribution for isoparafnic kerosene produced

from coal-derived synthesis gas by FT process.41


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Fuel Specication Requirements and Test Methods

Fuel specications and testing have been an integral part of thedevelopment of aviation fuels. Fuel specications are performancespecications in general and are written to allow any particularcombinationof hydrocarbonsthat meets the specied performance.Properties are allowed to vary to the extent that the variations willhave no deleteriouseffecton the performanceof the fuel in aircraft.The tighter (smaller variations in properties allowed) the speci-cation is written, the higher is the cost of the fuel in general. The

rst U.S. Government specication was issued in 1907 for the pur-chase of gasoline (U.S. Navy specication 24 G.5), and requiredonly that the fuel be of a certain specic gravity (density)and that itnot be contaminated with materials that would leave a residue afterfuel evaporation.12 As fuel demand grew and problems in servicebecame apparent, limitations on chemical composition and moreperformance requirements were added. By 1922 there were stan-dard tests established for color (Saybolt number), corrosion andgums (copper dish test), and sulfur content (Doctor test), as wellas others.15;19 As already described, octane number became an im-portantspecication requirementin 1930.Higher altitudeoperationled to a requirement for maximum freeze point. ASTM has been akey force in the development and standardization of test methodsfor fuel specication testing.

Current aviation fuel specication requirements can be brokeninto three categories: chemical composition, physical properties,andmiscellaneousrequirements.Compositionrequirementsincludetests to determine the hydrocarbon types in the fuel, to evaluatethe sulfur content, and to quantify the organic and inorganic acidsin the fuel. Physical properties specied include density (or spe-cic gravity), volatility, vapor pressure, ash point, viscosity, andfreeze point. Miscellaneous requirementsinclude determination ofheat of combustion, burning quality, corrosivity, cleanliness, par-ticulates, thermal stability, color, existent gum, a nd electrical con-ductivity. The requirementstable from the JP-8 specication is pre-sented in Table 2. The Jet A/Jet A-1 specication (ASTM D1655,Def Stan 91-91) is very similar. Note the signicant number oftest methods employed in certifying JP-8, most driven by opera-

tional problems.Moredetaileddiscussionscan be found in Refs. 15and 19.

A summary of jet fuel properties is given in Table 3.20;4346 Anexample of the variations in delivered JP-8 aromatic concentrationover several years is shown in Fig. 6, illustrating the variability incomposition permitted within the specication.

Additives15;19

Fuel additives are chemicals that are added to fuels to impartspecic properties, or to counteractthe effects of fuel contaminantsthat are not easily removed. The selection of additives is subjectto many constraints. Additives intended to enhance performanceof the fuel in one aspect can have deleterious effects on other fuelproperties. A major concern is the compatibility of additives with

thematerialsin thefuelsystemandin theenginehot section.A largenumberof fuel additiveshave been developedand investigatedoverthe years. The following paragraphs contain brief discussions ofthe various additives currently in use. Many of these additives areshown in Fig. 7.

Fig. 6 Distribution of JP-8 aromatics for 19901996 fuel purchases.44

Fig. 7 Jet fuel additives.

In 1954, the U.S. Governmentbegan using commercialpipelinesto transport aircraft fuel to U.S. Air Force facilities. To combat theexcessive corrosion in the ground fuel systems, and to reduce thecarryoverof corrosion products into aircraft fuel systems, corrosioninhibitoradditiverequirementswere addedto specicationsfor both

avgas and turbine engine fuels. A specication (MIL-I-25017) wasissuedin 1954for corrosioninhibitoradditives.Theseadditivesalsoactto improvefuel lubricity(lubricatingcapability)andare typicallyadded at approximately 20 ppm.

Water contamination in aviation fuels has always been a seriousproblem. In liquid form, water can cause temporary ameout in theengine, but in solid form (ice), it can block lters and fuel linesand completely stop the ow of fuel to the engine. In the 1940sand 1950s, free, undissolved water in fuel was suspected as thecause of many in-ight incidents and accidents. A major researchand development program was initiated to solve the water-in-fuelproblem as a result of a B-52 crash in 1958. One objective wasthe development of a fuel system icing inhibitor (FSII). The FSIIwas to be added to the fuel, but would preferentially migrate to

any free water present and act as an antifreeze. The current icinginhibitorused is diethyleneglycol monomethylether at a maximumconcentrationof 0.15 vol%. The FSII also acts as a biocide.

Because of their low electrical conductivity, aviation fuels canbuild up a static electrical charge, especially during fueling. Dis-charge of thisbuiltupcharge in areas where ammablefuel/air mix-tures exist, for example, fuel tanks, has been a problem. A solutionis use of static dissipator additives (SDA) in the fue l. Octel Stadis450 is the only currently approved SDA for turbine engine fuels.15

Typical concentrations of 0.52.0 mg/l increase the fuel conductiv-ityto between200 and600 pS/m. As shownin Table 2, theminimumconductivity allowed in JP-8 is 150 pS/m. The main difference be-tween commercialJetA-1 fuelandmilitaryJP-8 fuelis thespeciedpresence of corrosion inhibitor, FSII, and SDA.

Antioxidantadditives are added to turbine engine fuelsand otherpetroleum products to prevent the formationof gums and peroxidesduringstorageby reducing the formationof freeradicalsin thefuel.Peroxidesare deleteriousto thermal oxidativestability, possibly be-ingprecursorstotheformationofdeposits.Peroxidesalsoattackfueltankpolysulde sealantsand otherfuel systemelastomers.The mostcommon antioxidantsare hindered phenols,exemplied by 2,6ditertbutyl4methylphenol[butylatedhydroxytoluene(BHT), alsoused in food]. Normal antioxidant concentrations in turbine enginefuels range up to 14 mg/l.

Metal deactivatoradditives (MDAs) were initiallyadded to gaso-lines that had been treated using the copper sweetening process (amethodto convertmercaptansulfur compounds to less noxioussul-fur compounds). Copper is known to catalyze oxidation reactions

that form gums, and so MDA was used to deactivate any traces ofcopper left in the fuel. MDAs function by forming a chelate withthe metal. The chelate effectively isolates the metal from the fuel.MDAs are optionaladditivesin militaryturbineengine fuels. MDAsare approved for use in military turbine engine fuels at concentra-tions up to 6 mg/l.


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Table 2 JP-8 specication chemical and physical requirements, MIL-DTL-83133E

Test methodsProperty Minimum Maximum ASTM standards

Color, Saybolt a D156b or D6045Total acid number, mg KOH/gm 0.015 D3242Aromatics, vol% 25.0 D1319Sulfur, total, mass % 0.30 D129, D1266, D2622, D3120,

D4294b or D5453Sulfur mercaptan, mass % or 0.002 D3227

Doctor test Negative D4952Distillation temperature, Cc D86,b D2887

(D2887 limits given in parentheses) a

Initial boiling point 205 (186)10% recovered a

20% recovered a

50% recovered a

90% recovered a

Endpoint 300 (330)Residue, vol% 1.5Loss, vol% 1.5

Flash point, C 38 d;e D56, D93b or D3828d

Density or gravityDensity, kg/L at 15C or 0.775 0.840 D1298 or D4052b

Gravity, API at 60F 37.0 51.0 D1298

Freezing point,

C 47 D 2386,

b

D 5901, or D5972Viscosity, at 20C, mm2/s 8.0 D445Net heat of combustion, MJ/kg 42.8 D3338c or D4809b

Hydrogen content, mass % 13.4 D3701,b D3343Smoke point, mm, or 25.0 D1322Smoke point, mm, and 19.0 D1322

Naphthalene, vol % 3.0 D1840Calculated cetane index a D976f

Copper strip corrosion, 2 h at100C (212F) No. 1 D130

Thermal stability D3241g

Change in pressure drop, mm Hg 25Heater tube deposit, visual rating


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Table 3 Typical aviation fuel properties15;17;20;35;4446

Property Avgas JP-4 JP-5 JP-7 JP-8 (Jet A/A-1) RP-1

Approximate formulaa C7 H15 C8:5 H17 C12 H22 C12 H25 C11 H21 C12 H23H/C ratio 2.09 2.00 1.92 2.07 1.91 1.95Boiling range, F (C) 115295 140460 360495 370480 330510 350525

(46145) (60240) (180260) (190250) (165265) (175275)Freeze point, F (C)b 80 (62) 57 (49) 47 (44) JP-8/Jet A-1:60 (51); 55 (48)

Jet A: 50 (45)Flash point, F (C) 10 (23) 147 (64) 140 (60) 127 (53) 134 (57)

Net heating value, 18,700 18,500 18,875 18,550 18,650Btu/lb (kJ/kg) (43,490) (43,025) (43,895) (43,140) (43,370)

Specic gravity, 16C (60F) 0.72 0.76 0.81 0.79 0.81 0.81Critical temperature, F (C) 620 (325) 750 (400) 750 (400) 770 (410) 770 (410)Critical pressure, psia (atm) 450 (30.5) 290 (19.5) 305 (20.5) 340 (23) 315 (21.5)Average composition

Aromatics, vol% 25 10 19 3 18 3Naphthenes 29 34 32 35 58Parafns 59 45 65 45 39Olens 10 2 2 2Sulfur, ppm 370 470 2 490 20

aFor illustration of average carbon number, not designed to give accurate H/C ratios.bTypical.

inventoryaircraft. The JP-8C100 program led to a number of otheradvancements, including new test techniques, improved fuel/waterseparation technology, and a better understanding of fuel thermal-oxidative stability. The initial JP-8+100 additive package is basedon a detergent/dispersant, packaged with an anti-oxidant and MDA(as shown in Fig. 7) at a concentrationof 256 mg/l. The JP-8C100additive has been shown to produce 5095% reductions in depositsin more than 10 thermal stability test rigs with a wide variety ofJP-8/Jet A fuels.4957 Several base-level trials have quantied sig-nicant reductionsin fuel-relatedengine maintenanceduringuse ofthe additive. Further discussionon the topic of thermal stability ap-pears in the Future Trends section. Over 300 additive combinationswere screened in this program.

Missile Fuels

Missiles imposequite differentconstraintsfrom reusable aircrafton their fuels. Because of the expendable nature of missiles (and,thus, their relatively short liferequirementand small fuel consump-tion), the cost of missile fuels is usually not a signicant driver formissile system cost. Missiles are typically volume limited, ratherthan weight limited, as are most aircraft, so that fuel volumetricenergy is a key parameter. Storage stability is key because mis-siles may be stored fueled for periods on the order of 10 years.Low-temperature and low-pressure performance are key drivers forair-launched missiles, which may be cold-soaked at very low ambi-ent temperatures.For example,the U.S. AirForcerequiresa 54C(65F) freeze point fo r missile fuels, signicantly lower than the

freeze point of the fuel for the aircraft carrying the missile.Before the 1960s, the only fuels available for missiles were JP-4

and JP-5. RJ-4 was developed for the ramjet-powered Talos mis-sile in the early 1960s.15;58 As shown in Fig. 8, RJ-4 is a mix-ture of the isomers of exo-tetrahydrodi(methyl cyclopentadiene).Also known as TH dimer and H-MCPD, the need for a non-chemical name is obvious. Joint U.S. Air Force/Navy research inhigher density missile fuels for turbine-engine-poweredcruise mis-siles culminated in the creation of JP-10 in the mid-1970s.15;58

JP-10 is ex o-tetrahydrodicyclopentadiene (Fig. 8) and is the o nlyairbreathing-missile fuel in operational use by the United Statesat the present time. JP-10 has the impressively low freeze point of79C (110F).Higherdensitymissilefuelshavebeen developed,for example, RJ-5, perhydrodi(norbornadiene), but cost and freeze-

pointlimitationsprevented elduse. The potentialfor extendingtherange of missiles beyond what is possible with conventional high-density missile fuels prompted interest in slurry fuels that containadditives, such as boron, carbon, or aluminum in suspension in agelled form. Such slurryfuelscan providea very highheating valueper unit volume. The result of the many years of slurry fuel work

Fig. 8 Missile fuel structures.

demonstratedthat formulation of stabilized slurry fuels is possible,but to make such fuels a viable option for t urbine-engine-poweredcruise missiles, a substantial amount of additional fuel system and

propulsion engineeringwould be required.

Liquid Rocket Propellants

This discussionon rocketpropellantscan only give a avorof theinnumerable propellant combinations that have been tested/ownsince Goddards rst ight. Interestedreaders can consultthe manyreferences that follow, for example, Forbes and Van Splinter,59 orconsult the many web-sites devoted to space travel, for example,www.astronautix.com. A brief introduction to propellant develop-ment from the perspectiveof notable rockets is followed by a moredetailed discussion of the various classes of propellants.

Brief History of Liquid Rockets and Their Propellants

Early pioneers in the study and assessment of liquid rock-

ets include Konstantin Tsiolkovsky, Hermann Oberth, and RobertGoddard (see Ref. 3). In the early 1900s they pioneered the analy-sis of propulsionconcepts to get into space, often inspired by JulesVerne and other ction writers concepts for visiting other planets.These analyses concludedthat only liquidpropellants(vs solidpro-pellants or gases) had the necessarypropertiesto achieve access tospace. Tsiolkovsky pointed out the advantages of liquid hydrogen(LH2 ) and liquid oxygen (LOX) very early. In fact, 2003 could bedescribed as the centennialof rocket regenerativecooling,rst pos-tulated by Tsiolkovskyin 1903(see Ref.60). Manyprivate clubsandgovernmentorganizationssupported early tests of rocket concepts.Robert Goddard is credited with the rst ight of a liquid propel-lant rocket in March 1926. His choice of propellants was gasolineas the fuel and LOX as the oxidizer. LOX and noncryogenic nitric

acid were common oxidizers in this early period, but many fuelswere used. Throughoutthe 1930s, many groups in Europe and theUnited States tested larger and larger rockets, although none cameclose to getting to space. The advent of World War II acceleratedthe development of rockets, especially in Germany, with the mostnoticeable development being the A-4 rocket (popularly known as


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Table 4 Major operational liquid propellant families

Family Example Stoichiometry Isp , sa (Ref. 63) Attributes

Cryogenic LH2/LOX 2H2 CO2 ! 2H2 O 391 High performanceKerosene (cryogenic) RP-1/LOX CH2 C1.5O2 !CO2 CH2 O 300 Higher density, cheaper handlingStorable Nitrogen tetroxide/hydrazine 2N2H4 CN2O4 !3N2 C 4H2O 292 Storable, hypergolicStorable nontoxic Kerosene/peroxide CH2 C3H2O2 !CO2 C 4H2 O 273 Storable, relatively nontoxic

Mono-propellant Hydrazine 2N2 H4 !2NH3 CH2 CN2 23066 Decomposition over catalyst eliminates

need for separate fuel plus oxidizer

aEquilibrium 1000-psia (1-atm) expansion.

the V-2). The A-4 ight on 3 October 1942 was the rst rocketight that demonstrated the potential of achieving orbit, returningto Earth 120 miles from the launch site. The propellantsused on theA-4 were a 75% ethyl alcohol/water mixture as the fuel and LOXas the oxidizer. At rst glance, the use of an alcohol/water mixtureseems ratherodd becausemany other fuels give higherperformance(as described later). However, the use of the alcohol/water mixturehad the advantage of a lower ame temperature, which helped theA-4s fuel-cooledcombustorto survivethe relativelylongburn time.Other authors credit the scarcity of gasoline and diesel fuel for thechoice of ethyl alcohol as the V-2 fuel. Many other rockets weredeveloped for other military applications during the 1940s, such as

rocket-assisted takeoff and liquid-fueled missiles. These types ofapplicationsrequired storable,that is, noncryogenic,propellantsforease of use and readiness.61 The most common storable oxidizerin this period was nitric acid. Various hydrocarbons were used asstorable fuels (gasoline, jet fuel, etc.), but rockets using the hy-drocarbon/nitric acid combination were plagued with destructivehard starts and other combustion problems. These problems werelargely solved by the use of fuels that were hypergolic with nitricacid, such as aniline and hydrazine. Hypergolic propellants igniteon contact, which simplies the ignition process dramatically andeliminates most of the combustion problems such as hard starts. Inthe 1950s, nitric acid was largely supplanted by nitrogen tetroxideas the storable oxidizer of choice, usually paired with one of thevarious types of hydrazine. Storable propellant development dur-

ing the 1940s and 1950s was a very large research effort, devotedto optimizing propellant properties, material compatibility, storagestability, and combustion performance. After World War II, long-range rocketdevelopmentin the UnitedStates and the Soviet Unionwas the most vigorous, based in large part on the German A-4 (andits developers). The Soviet Union achieved the most notable suc-cessesin thelate 1950s,orbitingthe rstsatelliteand the rsthuman.The rocket used was the multistageR-7, with kerosene and LOX asthe propellants. In the early 1960s, the X-15 explored the limits ofwinged ight, reaching Mach 6.7 in 1967. The X-15 was poweredby a liquid ammonia/LOX rocket.

The difculty of summarizing rocketpropellantsis illustrated bythe Saturn V. In 1969, the United States reached Earths moon pro-pelled by the Saturn V/Apollo, which utilized engines ranging in

thrust from 100lb (445 N) (attitudecontrol systems) to 1.5 106 lb(6.7 kN) (the rst-stage booster engines). These engines used vedifferent propellants in three combinations.The rst-stage boosterengine propellantswere RP-1 kerosene and LOX. The second- andthird-stage booster engines were LH2/LOX. The Apollo commandmodule and the lunar descent and ascent engines, as well as theattitude control thrusters, utilized nitrogen tetroxide oxidizer andseveral types of hydrazinefuels. Spacecraft during this period gen-erallyusedthis storablepropellantcombinationfor control.An alter-nativepropellantforthese applicationswas hydrazineas a monopro-pellant, utilizinga catalyst to aid in the decomposition.A departurein design, if not in propellants, occurred with the U.S. space shut-tle, rst own in 1981. The space shuttle uses a reusable LH2/LOXmain engine, with a detachable external propellant tank and solid

rocketboosters,to reachorbit.The spaceshuttleorbitalmanueveringengine uses storable propellants(hydrazine/nitrogen tetroxide).Theauxiliarypowerunit useshydrazinemonopropellant.Major expend-able launch systems at the close of the 20th century utilized eitherLOX/LH2, forexample,someArianestages,Titan 4; LOX/kerosene,for example, Sea Launch, Atlas, Delta 3; or storables,for example,

Fig. 9 Equilibrium temperature in space;71 limits of thermal equilib-

rium depend o n insulation, absorptivity/emissivity, etc.

Chinese Long March, some Ariane stages.62 Some properties andattributes of these propellant families are presented in Table 4. Inclosing, note that the propellants currently in use were all in use oractive development in the 1950s (if not earlier); recent increases invehicle performance have come more from improvements in hard-

ware than from improvements in chemical propellants. New pro-pellants have not had a good record of making it into operationalvehicles, especially launch vehicles, in the past several decades.

Summary of Operational Propellants

This section gives some summary details on the various opera-tionalpropellantfamilies.Moredetailmay befoundin Refs.6, 7, 59,61,and6371. Note that some oxidizersandfuelscan alsobe usedasmonopropellants (such as hydrogen peroxide and hydrazine). Notealso that the term storable is confusing unless a storage location isgiven. For example, Fig. 9 shows that LH2would be space storablein orbit around Pluto, whereas LOX is space storable throughoutmuch of the solar system.71 Typically, storable is meant to referto the Earths surface conditions for launch vehicles and strategic

missiles.

Cryogenics (LOX/LH2, LOX/Kerosene)

For the properties of LOX/LH2, LOX/Kerosene, and kerosenepropellants,see Table 5.5;45;59;7274 Oxygen/hydrogen is the highestperformingoperationalpropellant family. Of course,the lowtemper-aturesand explosionhazards of hydrogenimposesignicant launchcostpenalties.Thelowdensityofliquidhydrogenisalsoadrawback.For these reasons, Russian launch vehicles and many U.S. vehicleshave used various types of liquid hydrocarbonsas a fuel. AlthoughGoddards initial ight test used gasoline,operationalhydrocarbon-fueledrocketssince thenhave typicallyused kerosene.As discussedearlier, the term kerosene is used to describe a distillate fractionof petroleum boiling between (roughly) 200 and 300C (390 and

570F). Jet fuels, as described earlier, are kerosenes; naturally, jetfuels such as JP-4 were the initial kerosenes used in rocket tests(although JP-4 does not strictly t the denition of a kerosene be-cause of its low initial boiling point). However, it was found that jetfuel didnot have tight enoughspecicationpropertiesto serve as aneffectiveprop ellant. The Rocketdyne Rocket Engine Advancement


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Table 5 Properties of LOX, LH2, and kerosene propellants5;45;59;7274

Relative density Heat of formation,Propellant Formula (specic gravity) cal/g (298 K) Comments

LOX O2 1.14 (90 K) 95LH2 H2 0.071 (20 K) 950RP-1 Approximate 0.806 (22C) 400a See Table 1;

C11:7H22:8 H/C1.952RG-1 C12:3H23:9 0.832 (22

C) 433 7585% naphthenes;H/C1.946

Syntin C10H16 0.8577

C234

aSignicant variations in Literature:423 (Ref. 73), 389 to 606 (Ref. 74), 400 (calculation from heat of combustion),74

444 (Ref. 59), and 389 (Ref. 5). RG-1 value calculated from heat of combustion using method in (Ref. 74).

Table 6 Properties of storable propellants5;59;63;65;66;67

Oxidizer, fuel, Heat of formation, Freezing Relative densityPropellant Formula and/or monopropellant cal/g point, K (293 K)

WFNA HNO3 O 660 231 1.50Nitrogen tetroxide N2O4 O 45.5

a 262 1.43Hydrazine N2H4 F, M C376 275 1.01MMH CH3 N2 H3 F C285 221 0.88UDMH (CH3/2 N2H2 F C205 216 0.79Aniline C6H5 -NH2 F C79 267 1.01Furfuryl alcohol C4 H3OCH2OH F 674 240 1.13Hydrogen peroxide (98%) H2O2 O, M 1320 273 1.45

aValue from (Ref. 5) other values: 50.8 (Ref. 65),50.9 (Ref. 59),73.9 (Ref. 63), and C25.0 (Ref. 67).

Program is usually creditedwith eliminatingthe shortcomingsof jetfuels as propellants by developingr ocket propellant 1 (RP-1, MIL-P-25576) in the mid-1950s. In comparison to jet fuel, RP-1 has amuch narrower allowable density range and lower limits on fuel

componentsthat were thoughtto causedepositsduring regenerativecooling, such as sulfur, olens, and aromatics.6 Typical propertiesare shown in Table 3. Deposit formation (coking) led to a maxi-mum fuel-wetted surface temperature limit that ranges from 290(550) (Ref. 75), to 450C (850F) (Ref. 76). The variability of thislimit is discussed further later in the Future Trends section. Theamount of coking is very dependent upon the fuel-wetted surface,with copper causing more deposition than steel. Russian kerosene(RG-1, naphthyl) is somewhat different from RP-1, notably with ahigher density/specic gravity (0.832 vs 0.806 at 22C) (Ref. 45).RG-1 has been chilled before use to increase its density (and, thus,the mass of propellant that can be loaded onto the rocket). Engineapplicationsincludethe RD-180, RD-170, and NK-33. Russian air-craft fuel T-6 has been usedin enginetestsin place of RG-1 because

of its similar physical properties.Some upper stages of the RussianProton launchvehiclesused a synthetic (nondistillate)hydrocarbonsyntin (or sintin) to increase payloadover that obtainablewith RG-1 (Ref. 77). These synthetic hydrocarbons produce higherIsp thanRG-1/RP-1 (at the same H/C ratio) by adding strain energy to themolecule and increasing its heat of formation. Syntin productionapparently ceased in 1996 after the break-up of the Soviet Union.78

Thestructureand heatof formationfor syntinin Table 5 areinferredfrom a report of the structure and are unconrmed. The structureis consistent with Russian descriptions of the fuel as a synthetichydrocarbon of the cyclopropane row.77 There is signicant cur-rent interest in cryogenicliquid methane, which has a higherIsp butlower density than kerosene.79

Storables

The properties of storable propellants are given inTable 6.5;59;6367 Earth-storablefuels and oxidizers were often usedin early rockets, and their development accelerated during WorldWar II.61 The mostcommon storable oxidizerwas nitricacid, whichcame in several forms. White fuming nitric acid (WFNA) is essen-

tially pure nitric acid. Red fuming nitric acid co ntains a signicantamount of nitrogen dioxide. Nitric acid caused high corrosion ratesin materials, but the addition of a small amount of hydrouoricacid (HF) acted as a corrosion inhibitor, forming inhibited red fum-

ing nitric acid (IRFNA). Gasoline and nitric acid were an earlystorable combination,but the combinationhad combustioninstabil-ity problems.Anilineand nitricacid were found to ignite on contact(hypergolic),which eliminated the combustion instability problem.Aniline is an example of a nitrogen-containing amine molecule,many of which were found to be hypergolic with nitric acid. Jetfuels could be made hypergolic with nitric acid by the additionof amines, for example, JP-X and the various mixed amine fuels(MAFs). Amines can be classied as primary (R-NH2) (where R isa nitrogen or carbon atom), secondary (R2NH), or tertiary (R3-N).Tertiary amines were the most hypergolic with nitric acid,but otherproperties then came into play in determining the most-desirablefuel. For example, the already-mentioned aniline had a relativelyhigh freezepoint (6C, 21F). A eutectic (roughly 50/50) mixture

of furfuryl alcohol and aniline was found to have a freeze point of42C (44F) and remained hypergolic with nitric acid. N-ethylaniline (C6H5 NHC2H5/had a freeze point of63

C (81F) andalso remained hypergolic.Rocket-assisted-takeoffunits for aircraftand surface-to-air missiles used late in World War II u sed N-ethylaniline and nitric acid.The surface-to-surfaceCorporal missile usedthe nitric acid/furfuryl alcohol/aniline combination.

Throughout the late 1940s and early 1950s, a large number ofstorable fuels and oxidizers were tried in various combinations.The most successful were the hydrazines, including monomethylhydrazine (MMH) and unsymmetrical dimethyl hydrazine(UDMH). Hydrazine (N2H4/ has a relatively high freeze point, isonly partially miscible with hydrocarbons, and is shock sensitive.Thus,MMH andUDMH, aloneor in mixtureswithhydrazine,are of-

tenused forimprovedhandlingproperties,despitea somewhatlowerIspthan pure hydrazine.For example, the Titan 2 uses Aerozine 50,a 50/50 hydrazine/UDMH mixture as the fuel. As MMH becamemore available, it was often used in place of Aerozine 50. Russianvehicles typically used UDMH. Addition of hydrazines to hydro-carbonscreated hypergolicmixtures with nitric acid; thus, the Nike


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Ajax missile fuel was 17% UDMH in JP-4. The UDMH/IRFNAmixture was used in the Lance missile and the Bell Agena rocketmotor. For oxidizers,N2O4giveshigherperformancethan nitric acidand is the most common storable oxidizer in current use, but has ahigher freeze point. Strategic missiles red from silos and launchvehicles, (for example, Titan 2) used hydrazine/N2O4 propellants.

Hydrogen peroxide (90%) has a similar freeze point to N2O4(12C, 11F) and somewhat lower performance, but has the ad-vantagesof being less corrosiveand less toxic than eithernitric acid

or N2O4(Refs.80 and81). Peroxide/keroseneenginesare undercon-siderationfor storable applicationswhere the toxicity of the propel-lants is a concern.Interest in peroxide as an oxidizer is increasing,with applications to new engines such as rocket-based combinedcycle.82 The ability of a catalyst to decompose peroxide (key to itsuse as a monopropellant as described later) allows some exibilityin combustionsystem design unavailable with other oxidizers. Onedifcultywithperoxideistheincompatibilitybetweenthedecompo-sitioncatalystand peroxideadditivesused to improve storagelife.81

Monopropellants

Variousmonopropellants(see Table6) have beenstudiedthrough-out the years, including hydrazine, hydrogen peroxide (used as agas generator on the German A-4), and various nitrates. A mono-

propellant is passed over a catalyst to generate hot gas, which canbe used for propulsion, attitude control, or turbomachinery drivegas. The rst monopropellanti n common use was peroxide (H2O2/passed over a catalyst (typically silver was the active metal). Forsafety (detonability) and performance, hydrazine is essentially theonly monopropellantin current use.66 A milestone in monopropel-lant development occurred in 1964, when the rst ambient tem-perature catalyst for hydrazine (Shell 405) was developed.69 Whenpassed over this iridium-based catalyst, hydrazine rapidly and re-producibly breaks down to a mixture of ammonia, nitrogen, andhydrogen. Although theIsp is relatively low compared to bipropel-lants, the simplicity of the single propellant and ease of ignitionand reliable performance have led to widespread use of hydrazineas a monopropellant. However, the toxicity of hydrazine has led to

a renewal of interest in peroxide, despite its lower performance.80Higher performing alternatives are also being examined, such asmixtures based on hydroxyl ammonium nitrate [(HAN), discussedin more detail in the next section].

Future Trends in Fuels and Propellants

Several drivers for changes to fuels and propellants seem appar-ent.Commercialand military kerosene jet fuels are becoming moresimilar and interchangeable.Logistical and cost considerationsleadto the expectation that this trend will continue. Ever-tighter emis-sions regulations seem certain to decrease the sulfur levels in avi-ation fuel in the next few years, probably to levels consistent withdieselandgasoline(lowtensofpartspermillion).Thislevelofsulfurmay create lubricity problems for pumps and other fuel-lubricated

components. As the production of oil begins to decline, the use offuels derived from nonpetroleumsources (coal, natural gas, etc.) isexpected to increase. If these fuels are produced from synthesis gas(FT), then the lack of aromatics in the fuel may cause problemswith seals andpumps,as describedearlier.41;42 In balance,however,low-sulfur/low-aromaticsfuels are expectedto reduce emissionsrel-ative to current fuels. Because of the signicant number of existingengines, it is anticipated that the general kerosene character of jetfuels will continue for many years, whatever the original source ofthe hydrocarbonsin the fuel.

Lowering the cost of access to space is a big driver for all launchcustomers. Improved propellants for both single-stage-to-orbitandtwo-stage-to-orbitvehiclesare beingsought.The use of airbreathingengines for all or parts of these vehicles missions is being consid-

ered. An area of high impact in this search for better propellantsis increasing the density of hydrogen (and kerosene) propellantsthroughvariousmeans, including decreasingtemperature(subcool-ing), gelling, and adding metals.83;84 In contrast, one alternative toachieving lower launch costs has been called the big dumb boosterapproach.4 In this approach, minimum cost is achieved through the

deliberate use of lower cost/lower performance systems, such asthe use of pressure-fed engines and LOX/hydrocarbon or perox-ide/hydrocarbonpropellants.An exampleof thisapproach(althoughnot fully demonstrated) was Beal Aerospaces jet fuel/peroxidebooster, the second stage of which was tested in 2000. One draw-back of this approach is the wide variation in jet fuel density andheat of formation permitted by current specications. The use ofcombined cycle engines raises the interesting option of using thesame fuel for both cycles to save on fuel systemweight and volume.

That raises the question of the most appropriate fuel for the com-bined cycle engine. For example, in a rocket-based combined cycleengine, should you use rocket fuel in the scramjet, or scramjet fuelin the rocket?

Because of the toxicity of hydrazine, alternatives are ac-tively being sought for both bipropellant and monopropellantapplications85;86 (in addition to peroxide as discussed earlier). Cur-rent alternatives being researched include HAN and dimethyl-2-azidoethylamine (DMAZ). One drawback to the alternatives formonopropellant applications is the higher temperatures achievedand the impact of these high temperatures on catalyst materials.87

Another drawback for bipropellant applications is the difcultyachieving the low ignition delay values of the hydrazine-fueledsystems.

Modeling and simulationof complex kerosene fuels has been re-ceivingalotofattention.Becausekerosenefuelsconsistofamixtureof hundreds (if not thousands)of hydrocarbonsthat vary from ren-eryto renery(and dayto dayin a givenrenerydependingoncrudesource), simplications are inevitable. One approach is the use offuel surrogates, where a much smaller set of hydrocarbons is usedto simulate the kerosene fuel, either in modeling or testing.The sizeof the set of surrogate hydrocarbonsdepends on the property beingsimulated.72 For modeling some simple ame behavior, a one- ortwo-componentsurrogate may be adequate. For combustion testingthat focuses on an aspect of fuel combustion that is dependent onmany aspectsof fuel composition(such as boiling range and chem-ical composition), a more complex surrogate might be necessary.46

Two areas of future development with signicant currentinterest

will be discussed in detail: 1) increased heat sink capability of ad-vanced hydrocarbonfuels/propellantsfor high-pressurerocketsandadvanced aircraft and 2) hydrocarbonfuels for airbreathing hyper-sonic vehicles. A common feature of these two areas is the difcultstruggle to increase hydrocarbon fuel thermal performance whileminimizing fuel system deposit formation.

Increased Heat Sink for Advanced Hydrocarbon Fuels and Propellants

The key challenge for hydrocarbon fuels and propellants is thedifculty of extending the cooling capability of the fuel to higherheatuxesand temperatureswithout havingthe vehicles lifelimitedby carbonaceousdepositionformation (coking, fouling) due to fuelthermal instability. For aircraft, higher engine cycle pressures andtemperatures (driven by desired increases in engine performance)

and higher subsystem cooling requirements are increasing the heatload being absorbed by the fuel. This is especially true for super-sonic aircraft, where ram air loses much of its cooling capability.Forrockets,higherchamber pressuresare desirablefor increaseden-gine thrust/weight, but increase the heat ux into the regenerativelycooled structure.

Jet fuel thermal stabilityalready has been briey described in thejet fue l additive section . An excellent review of jet fuel thermal sta-bility has been published.47 As discussed earlier, in current aircrafttherule of thumbis tomaintainJP-8/JP-5/JetA/Jet A-1temperaturesbelow 150163C (300325F) to prevent deposits on fuel systemcontrols, valves, screens, heat exchangers, and other components.As the fuel temperature increases above this level, fuel thermal-oxidativedegradationbeginsto leadto deposits.As shown in Fig. 10

(fora situationwherefuel temperatureincreaseslinearlywhile ow-ing from left-to-right), the roughly 70 ppm (Ref. 57) of dissolvedoxygen in fuel exposed to air reacts with the hydrocarbons to formperoxides and eventually deposits. Oxidative deposition ceases oncomplete consumption of the dissolved oxygen. Further increasesin temperatureleadto thermal cracking of the fueland a reinitiation


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of deposition. An overall outline of the thermal-oxidative deposi-tion process has been presented by Taylor (see Refs. 47 and 173)and is reproduced in Fig. 11. Jones and Balster have quantitativelyassessed the formation soluble and insoluble gums and solids andshown the relationshipbetween fuel oxidationand the formationofdepositprecursors.88 Much of the complexityof this processcomesfrom the involvement of trace heteroatomic species, for example,organic sulfurcompounds,and dissolvedmetals.The averagesulfurcontent is JP-8 is roughly 500 ppm (Refs. 44 and 89). In addition,

the deposits themselves are a trace component of the fuel, account-ing for much less than 1 ppm of the fuel ow througha component.Yet, over the multithousandhour lifeof typical aircraft fuel systems,evenbelow partsper million levelsof depositionmay be intolerable.The JP-8C100 thermal stability additive package discussed earlierappears to reduce deposition by reducing particle agglomerationand subsequent deposition.90 The effect of the additive on the fueloxidation reactions is fairly minor, although a delay in oxidation iscertainly seen.91 A further complication is the use of fuel recircula-tion onboardaircraft,which can acceleratethe depositionprocess.92

Moving beyond the200C (400F) capabilityprovided by a de-tergent/dispersantpackage such as C100 will require control of thefuel autooxidation chain mechanism. Oxygen is consumed rapidlyas temperature increases, due to (roughly) Arrhenius kinetics.47;93

For example, at 150

C (300

F), a typical jet fuel might require aresidence(reaction)time of 1 h to consume the 70 ppm of dissolvedoxygen present in air-saturated fuel. At 260C (500F), completeoxygen consumption requires on the order of 1 s. Under complete

Fig. 10 Fuel deposition regimes.

Fig. 11 Taylor mechanism for fuel thermal-oxidative instability (see Ref. 47).

oxygen consumption conditions, the oxidative deposition (scaledby the fuel ow) is relatively insensitive to ow velocity/residencetime.49;94 However, under conditions of incomplete oxygen con-sumption, the relationshipbetween oxygenconsumptionand depo-sitioncan be complex.95 As shownin Fig. 12,96 thiscomplexprocessoffers severalopportunitiesfor interruptingthe chain.Current mod-elsof theoxidationprocessincludeup to 19 reactionsand more maybe added to account for differences in oxidation behavior betweenthe aliphatic and aromatic fractions of the fuel. The current mod-

els lump he teroatom-containing species into o ne or t wo classes ofcompounds for purposes of modeling the effect of nonhydrocarbonspeciesonoxidationand deposition.97;98 Theseheteroatomicspeciesact as natural antioxidants99 with relatively nonlinear effects on de-positionwhengoodandpoorfuelsaremixed.100 It may benecessaryto separate these deposit precursors into different subclasses basedon a physical property such as polarity101 rather than heteroatomcontent. An alternative is to separate reactive species into differentchemical classes (such as sulfur being split into suldes/disuldesand thiophenes102 /, recognizingthe differingimpacton thermalsta-bility: Thiophenes are relatively unreactive, whereas suldes anddisuldes are recognized for decreasing thermal stability. Then, thechallenge for predicting fuel system life would be to estimate theappropriateheteroatomcontents to representthe averagefuel thatan

engine would encounter over its life, or the appropriatefuel qualitydistribution. It is conceivable that an engines (deposition-limited)life may be dominatedby theoccasionalmarginalfuel it encounters,rather than long periods of operation with typical fuels.

It has long been known that removal of dissolved oxygen in thefuel, typically by sparging with oxygen-free heliumor nitrogen,re-sults in a tremendous increase in thermal stability.47;103 Recently a

Fig. 12 Details of fuel autooxidationmechanism.96


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program has been initiated to develop an onboard fuel deoxygena-tion(degassing)systembasedonmembraneseparation.104 Chemicaldeoxygenationis also under current investigation,105 ;106 but the dif-culties involved with ensuring that the oxygen scavenging occursat the desiredpoint in the fuel system and only scavengesdissolvedoxygen (and not other reactive heteroatoms) have not been over-come. It is desirable that oxygen scavenging occur not in the fueltank (with its unlimitedsupplyof atmosphericoxygenbecausemosttanks are vented) but elsewhere in the fuel system, where the pres-

enceof the dissolvedoxygen could lead to depositformation.Thus,a scavengeradditivethat only becomes activeat temperaturesabove(notionally)93C (200F) would be the most effective. An alterna-tive approach involves the addition of peroxidedecomposers to in-terruptthe chainat theperoxidestep in Fig. 12.Triphenylphosphine(TPP) andsulfurcompoundshave beenstudiedfor thispurpose,withsome successin slowing oxidation and reducing deposition.107 TPPby itself is not as effective as the JP-8C100 additive package in re-ducing depositionfrom JP-8, but the combination together is betterthan either additive separately.

The effect of surfaceson thermal-oxidativedepositionis complexand still poorly understood.47 Smoother surfaces appear to inhibitdeposition, or (more precisely) to lengthen the induction time, theinitial period of low deposition rate often seen on clean surfaces.

Recentresults indicate that inert surfaces based on silicaalso inhibitdepositionin a complexmannerdependingonlocationalongthe fueloxidation reaction pathway.108 ;109 It is to be expected that a surfacecoating would lose its effectivenessafter accumulation of depositseffectivelyisolates the surface from the fuel depositionprocess. Anexampleof the result of usingthe bestavailableadditive(C100),bestavailablesurfacecoating(Silcosteel), and deoxygenationis shownin Fig. 13 for an extended test at 370C (700F) fuel temperatures.This indicates that progress is indeed being made toward the goalof a conventional jet fuel stable throughout the thermal-oxidativeregime. Deposition-resistantheat exchangerdesigns have also beeninvestigated.110

Note that the basis of most of the currentresearchis basedon thepremise that the approach to improved fuels is limited to methods

of improving currentjet fuels, for example, JP-8, Jet A, without in-cluding rening changes. This approach is driven by fuel cost: Thefuel cost is a signicant fraction of aircraft operating cost, so thatchangesin fuel that changethe baseline fuel itself are consideredtobe cost prohibitive.The addedlogisticsburden of handling specialtyfuels is also signicant. There may be niche applicationswhere thisis not the case and specialty fuels are preferable. If one is will-ing to consider higher cost fuels and/or specialty chemicals, thereare a wide variety of fuels that have higher thermal stability thanconventionalkerosenejet fuels, as shown in Table 7 (Refs.49, 111)and Fig. 13 for JP-7. The relatively poor thermal stabilityof methylcyclohexane(MCH) has been observed by several investigatorsandhas yet to be adequatelyexplained.

The desire for increased rocket thrust/weight is driving the pres-

sure of hydrocarbon-fueled thrust chambers higher.84

The current

Fig. 13 Comparisonof deposition reductionapproachesin single-tubeheat exchanger (described in Ref. 109); conditions: 1-gal/h fuel ow

rate, 0.085-in.-i.d. (2.15-mm) tubing, and 750F (400C) fuel outlettemperature.

Table 7 Thermal-oxidativesurface deposition

rates (complete oxygen consumption)49;111

Total surface depositionFuel rate, ppm

JP-8 (Jet A) 0.81.6JP-8C100 0.080.2JPTS 0.12JP-7 0.07MCH 0:5

JP-10 0:1Decalin 0:08

Fig. 14 Effect of chamber pressure on nozzle heat ux.113

issue is how (or if) regenerative cooling limits for hydrocarbonsmight be extended for high-pressure engines. (Cooling limits fornonhydrocarbons are also available.76;112/ The nal developmen-tal F-1 LOX/RP-1 engines for the Saturn 5 approached 70 atm(1000 psia) in chamber pressure in the late 1960s.The current stateof the art in hydrocarbon engines, as exemplied by the RD-180,utilizes chamber pressures in excess of 250 atm (3675 psia). Thisincreasedpressurehelpsenablean increasedIsp(311s (sea level) fortheRD-180 vs 265s forthe F-1).As discussedlater,engineheatuxis roughly proportional to chamber pressure, so that this increasedenginepressure comes at the expenseof increasedcombustion heat

ux delivered to the chamber wall and the fuel. An example ofthis is shown in Fig. 14,113 where the throat heat ux increasesfrom 10 Btu/in.2 s (F-1) to about 70 Btu/in.2 s at 3750 psia. Thisincreased heat ux may require augmentation of the regenerativecooling by lm or transpiration cooling.114 However, lm coolingcan result in performance losses, so that the desired approachis toextendthe regenerativecooling limits of the propellantused to coolthe thrust chamber and nozzle.

At rst glance, it would seem that regeneratively cooled thrustchambers would share many similarities with fuel-cooled aircraftcomponents. However, as touched on in the cryogenic propellantsection, the thermal stability requirements for rockets are muchmore stringentthan thosefor conventionalaircraft,despitethe muchshorter lifetime of the rocket. This is due to the much higher heat

uxes encountered in the rocket. At these high heat uxes, thethermal resistance from deposits in the cooling channel can veryrapidly cause structuraltemperaturesto rise past the point of failure(a burnthrough). The aircraft and rocket conditions are contrastedin Table 8. The rocketthermal stability problem has received muchless attention than the aircraft case. It is instructive to go into a bitmore detailon the rocket case to illustratefuture research directions(and past research head scratchers).

As shown in Fig. 15, regenerative cooling is best expressed as aseries of heat transfer processes.Combustion (hot-side)heat ux iscalculated as

q h g.Tg Twg/

where q is the heat ux for example, in British thermal units per

squre inch per second,h g is the combustion-side heat transfer co-efcient, and the temperatures are approximately as shown.114 ;115

Combustion-sideheat transfer is beyondthe scopeof this paper, buta few comments are relevant.

1) Heat transfer coefcient hg P0:8, so that the trend of increas-

ing chamber heat pressure increases wall heat ux.113


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Table 8 Comparison of aircraft and rocket cooling conditions for notionala dvanced missions

Cooling conditions Aircraft heat exchangers/fuel nozzles Rocket regen cooling channels

Typical maximum heat ux, Btu/in.2 s 1 100Required lifetime 2000 h (minimum) 300 s/missiona 50Materials Superalloys, perhaps coated Cu alloys

Incompatible materials Cu, Cd, Zn, Pb Cu incompatibility with S in fuel75

Effect of dissolved oxygen Oxygen removal below 1 ppm (from typical70 ppm) dramatically reduces deposition47

Little effect?123

Maximum deposition r ates JP-8:JP-7140:1 JP-8:RP-1:JP-71:1:1 (Ref. 117)

Maximum allowable deposit thickness 0.001 in. (10% ow reduction in typical0.020 in.-i.d. passage)

1 106 in. (estimate) to avoid burnthroughin Cu at 100 Btu/in.2 s

Deposition mechanism 150315C (300600F), molecular growth throughhydroperoxide chain mechanism with accelerationby polar heteroatomic impurities;>480C (900F),pyrolytic fuel cracking leads to molecular growththrough radical chain reactions

On Cu surfaces, sulfur-enhanced corrosionof the surface75 ; on non-Cu surfaces,mechanism is unclear.

Specication thermal stability limits? Yes, ASTM D3241 No

aMissionsD 4 h.

Fig. 15 Regenerative cooling schematic.114;115

2) The heat ux absorbed by the propellantalso increasesasP 0:8,but the propellant mass ow increases as P 1:0 , so that the overalltemperature rise in the propellant may decrease with pressure. Inother wordsTwc increaseswith pressure, butTc at the coolant jacketexit may decrease as thrust chamber pressure increases.

3) Carbon (soot) deposition on hot-side (combustion) chamberwalls acts as a thermal resistance and thus reduces heat ux.116

On thecoolantside,heatux is a strongfunctionof propellantve-locity through the Reynolds number (which is directly proportionalto ow velocity),

q D hc.Twc Tc/

whereh c Rea Prb , with the exponent atypically ranging from 0.8

to 0.9 (Refs. 117 and 118).Fuel heat ux capability is increased by increasing ow velocity

and/or careful cooling channel design. Flow velocities can exceed200 ft/s, which can cause uid hydraulic problems.76 High heatuxes can be absorbed by uids undergoingnucleateboiling,119 butcurrent engine designs for hydrocarbons employ pressures well inexcess of the critical pressure [310 psia(21atm) for RP-1], so thatthis type of heat transfer augmentation is not an option. Note thatany thermal resistance on the coolant side of the thrust chamberacts to increase Twcfor a givenheat ux, which canlead to chamberfailure (burnthrough) if material temperature limits are exceeded.Carbon deposition is the limiting factor for hydrocarbonpropellantheat ux capability.113 ;120 As shown in Table 9, the excellent in-sulating properties of coke deposits are evident. The temperature

rise due to coking is proportional to the heat ux, and so rocketscan tolerate much smaller thicknesses of deposition than aircraft.At the high velocities in regenerative cooling channels (which canexceed 100 ft/s), carbon deposition can also create increased fuelsystem pressure drop and result in injection problems. Signicantresearch efforts in the last 20 years or so have been reported by

Table 9 Thermal conductivities of variousmaterials relevant to regenerative cooling

Material k, Btu/h ft F (W/m K)

Copper 210 (360)Alumina 3.5 (6)Superalloy 13 (22.5)Coke deposita 0.07 (0.12)

aData from Hazlett.47

Table 10 RP-1 coking wall temperature limit, Twc

Upper temperatureReference limit, F (C)

Ziebland and Parkinson, 1971 (Ref. 112) 800 (425)Van Huff, 1972 (Ref. 76) 850 (450)Wagner and Shoji, 1975 (Ref. 113) 650700 (340370)

Wheeler, 1977 (Ref. 127) 600 (315)Rosenberg and Gage, 1991 (Ref. 75) 580 (305)Cook and Quentmeyer, 1980 (Ref. 120) 600 (315)Michel, 1983 (Ref. 128) 550 (290)

Aerojet,75;121 United Technologies Research Center (UTRC),122 ;123

Rocketdyne,124 andNASA JohnH. Glenn ResearchCenter at LewisField (GRC).117 ;125;126

Deposition is avoided by keeping the fuel temperature within acoke limit, which is usually specied as a maximum coolant-sidewall temperature (Twc in Fig. 15). However, this limitingTwc is notuniversally agreed on or well characterized in terms of its relation-ship to coolant velocity, system life, etc. As shown in Table 10,

published values for RP-1 vary from 288 to 455

C (550 to 850

F)(Refs. 75,76,112,113,120,127and 128).Typically,propane is foundto havesimilar limiting temperatures,whereas methanes limit is of-ten citedas severalhundreddegreesFahrenheithigher.118 ;120 Thein-teractionof the fuelsulfur with copper(the preferredthrust chambermaterial)is a key driver for hydrocarbonpropellantdeposition.75;121

As described in Table 8, copper is avoided in aircraft fuel systemsbecause below part per million levels of dissolvedcopper in jet fu-els can have deleteriouseffects on thermal stability.47 Under rocketconditions,gold121 andnickel122;123 coatingson coppersurfaceshavebeen found to be very effective in reducing sulfur-relatedcorrosion.

The carbondepositionrate is often approximatedas in Arrheniusform,

deposition rate exp.E=RTwc/

yieldingplotssuchasthatshowninFig.16.75 Suchastrongtempera-ture dependence would be difcult to overcome to increase regen-erative cooling capability of hydrocarbonsby increasing Twc . Notealso the presenceof a velocitydependenceof deposition.As coolantvelocity increases at a constantTwc , the boundary layer next to the


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Fig. 16 Typical deposition (thermal resistance) behavioras a functionof temperature.75

wall thins and the coolant residence time decreases. Note the non-linear behavior shown at lower velocities. This is similar to thebehavior of kerosenefuels in aircraft heat exchangersand fuel noz-zles as described earlier, where deposition peaks at about 260C(500F) and then decreasesas the limited amount of dissolvedoxy-gen in the fuel(the depositprecursor)is consumed.47;129 Depositionafter this point is low, until fuel thermal cracking commences atabout 480C (900F) (as shown in Fig. 10). The very high veloci-

ties (and resulting low residence times) in the rocket lead to somesignicant differences between the aircraft and rocket cases. Forexample, in most rocket situations, the dissolved oxygen is onlypartially consumed, whereas aircraft fuels at the same temperatureswould have completelyconsumed the dissolved oxygen.This high-velocity/short-residence time has produced some incongruous re-sults,whenviewedfromanaircraftperspective.Forexample,UTRCtests of deoxygenated fuels showed little improvement in fuel ther-mal stability123 in dramatic contrast to the results shown in Fig. 13under aircraft conditions. Aerojet tests with pure, very low sulfurhydrocarbons also showed no apparent improvement over RP-1, incontrast to expectations.75 Note that current RP-1 has very low sul-fur levels,on theorderof 20 ppm, much lowerthan thespecicationlimit (500 ppm). It remains to be demonstrated that reducing sul-

fur levels below 20 ppm provides any benet to hydrocarbon fuelregenerative cooling of rockets.

Several types of hydrocarbons are being examined as possibleenhancements/alternativesto RP-1 (Ref. 84). Alternative kerosenestend to focus on higher density, although lower H/C ratios oftenmitigate any net vehicle benets.130 Cryogenic hydrocarbons, forexample, propane, methane, offer increased Isp at the expense ofreduced density.79 High-energy (strained ring) hydrocarbons offerhigh Ispand potentiallyincreaseddensity, but are relativelyunchar-acterized and may have toxicity and cost concerns. Note, however,thatRP-1 costsfor currentexpendablelaunchvehicles(Atlas,Delta)are roughly 0.05% of the $50 million plus launch cost. A key lim-itation of this assessment of alternative propellants is the lack ofunderstanding of the relationship between propellant composition

and deposition levels. An effort is currently underway to collectmore data in this area, using both existing and newly constructedhigh-heat-ux test rigs.131 The need forthis newdata becomes moreevident as one looks at the differences between the behavior of jetfuels under aircraft conditions and rocket propellants under high-heat-ux conditions. For example, recent NASA GRC data showsroughly equivalent thermal stability for JP-8, JP-8C100, RP-1, JP-7, and JP-10 (Ref. 117), as shown in Fig. 17. These results are forstainless-steel surfaces: On copper, the high sulfur JP-8 and JP-8C100 give much larger deposition rates than the other fuels, asexpected. The contrast with the depositionlevels for these same fu-els under aircraft conditions (lower velocity/heat ux) on stainlesssteel,as shownin Fig. 13and Table7, is dramatic.Currenteffortsareassessingthe thermalstabilityunderrocketconditionsforsubcooled

cryogenichydrocarbons,improvedkerosenes,forexample,verylowsulfur, higher density, and prospective high-energy, strained ringhydrocarbons.84

Based on the precedingdiscussion,one might expect that roughly480C (900F) would be an upper (wetted-wall)limit for hydrocar-bon propellants. However, signicantly higher temperatures than

Fig. 17 Average deposition rates of various fuel under rocket regen-erative cooling conditions117 (75 ft/s, 1000 F wall).

Fig. 18 Heat sink required as a function of Mach number33;134137

[turboramjet (TBRJ)].

that are routinely encountered in current hypersonic (scramjet) en-gines. The reason for that lies in the great need for fuel heat sink inscramjet engines.

Hydrocarbon Fuels for Hypersonics

For combined cycle engines used in the rst (airbreathing)stageof a two-stage-to-orbitsystem, as wellas for hypersoniccruisevehi-cles and hypersonicweapons, thescramjetpart of the cycleappearsto be thedriverfor improvedfuels. As describedin referenceson thesubject,3235;132 scramjetoperationat Mach 6 andaboveplaces seri-ousdemandson theheat sink capacityof fuelusedto cool theengineregeneratively (especially for hydrocarbonfuels). As in the rocketrst-stage booster case, the low density of hydrogen makes liquidhydrocarbons competitive as fuels.31 At Mach 8, uncooled scram-

jet combusto r structures can exceed 5000F, well in excess of thetemperaturecapability of known materials, demonstrating the needfor fuel cooling.133 Although the amount of fuel heat sink required

is obviously aircraft and mission dependent, some general trendsare evident in Figs. 18 (Refs. 32, 134137) and 19.138;139 As shownin Fig. 18, heat load is roughly proportional to Mach2, as mightbe expected from the proportionality of air stagnation temperatureto Mach2. As shown in Fig. 19, airframe (external surface) heatingat higher Mach numbers can eliminate many popular lightweightmaterials choices, for example, aluminum, as well as creating sig-nicant fuel tank heating, which can decreasethe ultimate heatsinkcapabilityof the fuel(coolant).High fuel tank temperaturescan alsodrive a fuel volatility limit to avoid fuel boiling in the tanks, as wasrequired with JP-7 in the SR-71. Achieving these heat sink levelscan drive signicant sensibleheating of the fuel,Cp1T. Under typi-cal fuel system conditions,when the fuel temperature exceedsabout480C(900F), thehydrocarbonsin thefuel beginto thermallyreact.

Several of these reactions can absorb signicant amounts of heat;hence,the termendothermicfuel is used whenthe fuel reactionsaredeliberate. This is shown schematically in Fig. 20, which indicatesthe large enhancementof fuel heat sink capabilitypossible with en-dothermic reactions, although the heat sink capability still falls farshort of hydrogenon a gravimetric basis. However,the ratio of fuel


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Table 11 Chemical heat sink values for various endothermic fuelsa

Theoretical chemical Calc. heat of combustionheat sink, unreacted of endothermic products,

Endothermic reaction Reaction types fuel, kJ/kg unreacted fuel, kJ/kg

C7H14 (MCH)!C7 H8 (toluene)C 3H2 Dehydrogenation 2190 45,800C12 H24 (kerosene)!C2 H4 (ideal) Cracking 3560 47,200C12 H24 !CH4 , C2H4 , C2 H6, etc., (actual) Cracking 3500C12 H24 C 6H2 O! 9CH4 C 3CO2 (AJAX stage 1) S team reforming net: (stage 1C 2) Net: (stage 1C 2)CH4 CH2 O!COC3H2 (AJAX stage 2) Steam reforming

Net: C12 H24 C 15H2 O! 9COC3CO2 C27H2 Steam reforming 5490b 21,2402NH3 !N2 C 3H2 Dehydrogenation 2720 19,280CH3OH!COC 2H2 Dehydrogenation 4000 20,420Decalin (C10 H18/! naphthalene (C10 H8/C5H2 Dehydrogenation 2210 40,700

aHeat sink for kerosene calculated using dodecane heat of formation as an approximation.bHeat sink for steam reforming based on total propellant mass (fuel plus water).

Fig. 19 Uncooled surface temperatures as a function of Machnumber.138;139

Fig. 20 Comparisonof heat sink capabilityof LH2and hydrocarbons.

heatsink to fuel heat of combustionis comparableforhydrogenandendothermichydrocarbonfuels.

There are several types of heat-absorbing (endothermic) reac-tions possible during regenerative cooling, as given in Table 11.These endothermic reactions can be purely thermal,

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