Nasa Aprop1

David J. Bents, Ted Mockler, and Jaime MaldonadoLewis Research Center, Cleveland, Ohio

James L. Harp, Jr. and Joseph F. KingThermoMechanical Systems Inc., Canoga Park, California

Paul C. SchmitzPower Computing Solutions Inc., Avon, Ohio

Propulsion System for Very High AltitudeSubsonic Unmanned Aircraft

NASA/TM—1998-206636

April 1998

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David J. Bents, Ted Mockler, and Jaime MaldonadoLewis Research Center, Cleveland, Ohio

James L. Harp, Jr. and Joseph F. KingThermoMechanical Systems Inc., Canoga Park, California

Paul C. SchmitzPower Computing Solutions Inc., Avon, Ohio

Propulsion System for Very High AltitudeSubsonic Unmanned Aircraft

NASA/TM—1998-206636

April 1998

National Aeronautics andSpace Administration

Lewis Research Center

Prepared for theAerospace Power Systems `98sponsored by the Society of Automotive EngineersWilliamsburg, Virginia, April 21–23, 1998

Available from

NASA Center for Aerospace Information800 Elkridge Landing RoadLinthicum Heights, MD 21090-2934Price Code: A03

National Technical Information Service5287 Port Royal RoadSpringfield, VA 22100

Price Code: A03

NASA/TM—1998-206636 1

Propulsion System for Very High Altitude Subsonic UnmannedAircraft

David J. Bents, Ted Mockler, and Jaime MaldonadoNASA Lewis Research Center

James L. Harp Jr. and Joseph F. KingThermoMechanical Systems Inc.

Paul C. SchmitzPower Computing Solutions Inc.

ABSTRACT

This paper explains why a spark ignited gasolineengine, intake pressurized with three cascaded stagesof turbocharging, was selected to power NASA’scontemplated next generation of high altitudeatmospheric science aircraft. Beginning with the mosturgent science needs (the atmospheric samplingmission) and tracing through the mission requirementswhich dictate the unique flight regime in which thisaircraft has to operate (subsonic flight @ >80 kft) webriefly explore the physical problems and constraints,the available technology options and the cost driversassociated with developing a viable propulsion systemfor this highly specialized aircraft. The paper presentsthe two available options (the turbojet and theturbocharged spark ignited engine) which arediscussed and compared in the context of the flightregime. We then show how the unique nature of thesampling mission, coupled with the economicconsiderations pursuant to aero engine development,point to the spark ignited engine as the only costeffective solution available. Surprisingly, this solutioncompares favorably with the turbojet in the flightregime of interest. Finally, some remarks are madeabout NASA’s present state of development, andfuture plans to flight demonstrate the three stageturbocharged powerplant.

INTRODUCTION

Because of the increasing influence of man-madepollutants and their potential impact on Earth’satmosphere, the science community is expendingconsiderable effort to gain a better understanding ofits detailed chemistry and dynamics. Much of the workinvolves the development of more sophisticatedcomputer models of the atmosphere. These arevalidated through correlations with observed data,which includes both remote sensing and in situmeasurements. At present, the highest prioritymeasurements are in situ measurements at altitudesabove 73 kft to over 80 kft, especially within 12° of

the Equator. The in situ measurements are hardest toobtain since they involve physical samples taken byairborne instruments. Aircraft are the preferredinstrument platform because of the length anddirectedness of the flightpath, which allows largenumbers of samples to be obtained at specifiedlocations in the atmosphere, at the specific timesdictated by science opportunity.

The most urgent need is for an aircraft that can flylong distances at altitudes significantly above 80 kft.The aircraft presently used for sampling, even thehigh altitude ER-2, are not capable of flying muchhigher than 73 kft. While balloon borne instrumentscan reach altitudes as high as 130 kft, the undirectednature of balloon flight limits the geographic coverageand spatial resolution that is needed (the coverageachieved by a single airplane flight is the equivalent of10—100 simultaneous balloon flights). There are alimited number of supersonic aircraft capable of flyingover 80 kft, but these aircraft achieve high altitude byflying supersonically. The aerodynamic heating andshock associated with supersonic flight causechanges to the air sample which negate themeasurements being made. The airplane whichperforms the sampling mission must be subsonic.This presents a contradiction—the most straight-forward way to achieve high altitudes is to fly fast, butthis airplane must fly high and slow—a very difficultthing to achieve. Because of the exponential lapse ofair density with altitude, a subsonic aircraft flying at80 kft altitudes cannot generate much lift (see Fig. 1).Even at reasonably high speeds (M = 0.5) thedynamic pressure available limits wing loadings to only7 to 12 psf; more like a sailplane than a poweredaircraft.

THE PROPULSION CHALLENGE

It is widely acknowledged that the propulsionsystem is the most difficult technical challenge.Whether it is manned or unmanned, an aircraftdesigned to fly subsonically >80 kft for >4 hrs will

NASA/TM—1998-206636 2

require a propulsion system that is quite different fromexisting systems. Because of range and flightduration, air breathing propulsion is required. In thisflight regime, however, air breathing propulsion isdifficult to achieve. The difficulty arises from theexponential lapse of air density and pressure withaltitude. At 80 kft the ambient air density and pressureare about 1/30th of sea level values. As Fig. 1 shows,subsonic forward speeds do not generate much inletpressurization (at 80 kft, M = 0.5, less than 10 psf isavailable). Therefore, turbomachinery is needed tosupply most of the intake pressurization required tocompress ambient air into a powerplant working fluidof reasonable density.

To pressurize the intake to 1 atm at 80 kft, anoverall pressure ratio (OPR) better than 36:1 isrequired. Several turbomachinery stages are neededfor intake pressurization. For example, at least 3centrifugal compressor stages are required to providean OPR of 30 to 40; more if an axial compressor isused. Because of low inlet density, theturbomachinery is large in size, especially the firststage. Volume flow (corrected flow) requirementsincrease with altitude, which translates to largerturbomachinery diameters. Pressure ratiorequirements also increase with altitude, whichtranslates to more turbomachinery stages. Sincepower is proportional to airflow for any air breathingengine, the machinery size required to process airflowfor a given rated power will grow with altitude as OPRand corrected flow are increased. Figure 2 illustrateshow machine diameter and length must change asOPR and corrected flow are increased to compensatefor altitude. Two turbo-machines are shown. One is anaxial flow unit typical of a turbojet engine, pressurizinga 40 psia combustor flowing a constant 100 lbm/sec.The other is a centrifugal unit common toreciprocating engine turbochargers, maintaining aconstant 30 in HgA manifold pressure and flowing0.2 lbm/sec. Both machines are compressing fromaltitude ambient conditions at the inlet (US Std.Atmosphere). Interstage cooling is assumed forboth, and consistent tip speed limits and inlet flowvelocities are observed for each machine type. Notehow machine dimensions must increase as the inletconditions are changed from sea level to 90,000 ft.For either unit, length increases more than 5 timeswhile diameter increases more than sevenfold. Weightis proportional to the cube of linear dimensions.

The atmospheric density lapse causes the needfor very large heat exchangers. The density lapsefrom sea level to 80 kft produces a five fold decreasein Reynolds number (Re) and more than tenfolddecrease in convective heat transfer. Heat exchangersizing is driven upwards by three factors: (a) reducedconvective heat transfer available, (b) the lowerdensity ambient air has less heat capacity, and (c)

the need for more heat rejection, due to the increasedcompression heat loads associated with higher OPR’s For example, Fig. 3 shows how the weight and frontalarea of a typical aircraft engine coolant heatexchanger must increase to reject the same heat loadat altitude versus a sea level unit (including the effectsof reduced air temperature with altitude ).

Flying subsonically in low density air also createsthe requirement for larger thruster “capture area” or“actuator disk” areas in order to achieve reasonablepropulsive efficiencies that are needed to reduce fuelconsumption and provide range. Figure 4 shows therelationship between the minimum thruster “capturearea” needed to maintain 100 lbf thrust versus altitude(US Std. Atmosphere), for a given propulsiveefficiency. Two representative airspeeds areillustrated. Mach 0.8 represents the approximateupper speed limit for a subsonic jet aircraft, whileMach 0.4 represents a slower aircraft speed thatreduces fuel burn. At 80 kft and M = 0.8, a 50 percentpropulsive efficiency requires at least 1 ft2 capturearea, which is equivalent to the nozzle exit of a smallturbojet. If capture area can be raised to 8 ft2 ,however (equivalent to a turbofan engine), propulsiveefficiency improves to 90 pct. At the slower speed ofM = 0.4, the same 8 ft2 capture area would produceonly 70 pct efficiency—to regain 90 pct efficiency thecapture area has to be increased to at least 30 ft2 (i.e.a propeller—note that actual capture areas will besomewhat larger than this idealized minimum). Subsonic flight at altitudes >80 kft will favor thrusterswith relatively large actuator areas (i.e. a propellerdriven aircraft), and the actuator size (i.e. propellerdiameter) will be 2 to 3 times larger than what iscommon for a conventional aircraft.

Because of the increased size and weight of theair handling, thermal management and thrust deliverycomponents, a propulsion system designed for highaltitudes is significantly larger and heavier than its lowaltitude counterpart. Further complicating matters, thehigh altitude aircraft will need more power to stay aloftbecause of the faster flight speeds necessary (tomaintain dynamic pressure and support its weight inlow density air). The propulsion system grows in bothrating and in specific weight, which tends to claimgreater and greater fractions of the airplane’s grossweight. This of course runs counter to the airplane’sability to carry the weight.

PROPULSION SYSTEM OPTIONS

The weight penalty associated with air handlingand thermal management becomes a majordiscriminator when chosing propulsion for the highaltitude aircraft. There are two powerplant candidatesto consider: turbine engines (i.e. a turbojet) andturbocharged reciprocating engines. The turbojet

NASA/TM—1998-206636 3

engine would be associated with a relatively highspeed (0.7<M<0.8) aircraft of conventional size withwing loadings in the range of sailplanes and lightaircraft (15 to 25 psf); the turbocharged reciprocatingengine would power a very large very lightlyconstructed propeller driven aircraft (M = 0.4, wingloading 4 to 7 psf) that maximizes range andendurance. Figure 5 shows a represent-ative turbineengine (turbojet); Fig. 6 shows a representativeturbocharged reciprocating engine. Both are shownconfigured for the >80 kft flight regime. As designedfor >80 kft flight, the distinctions between the twobecome somewhat blurred since the turbochargedreciprocating engine could be considered as a variantof the turbine engine, where a reciprocating enginehas replaced the combustor core. Both require largeamounts of air, but there are major differencesbetween them concerning airflow usage. The turbineengine ingests large amounts of air whose heat ofcompression is retained in the cycle, while theturbocharged reciprocating engine compresses only asmall fraction of the air it ingests, and uses theremainder only for heat rejection. Heat rejectionincludes what is lost to engine oil and coolant, plusthe induction air stream’s heat of compression. Theturbine engine rejects almost all of its waste heat inthe cycle exhaust (which is used for for thrust), whilethe turbocharged reciprocating engine rejects asignificant fraction of its waste heat to the engine oiland coolant. Large HX’s are required to removewaste heat from engine coolant and oil, andcompression heating from the induction air (but thecooling air flowing through them does not have to bemechanically compressed).

At low altitudes, the turbine engine can generate 2to 5 times higher power density than the reciprocatingengine as long as inlet air mass flow is adequate. Mass flow is easily obtained at low altitudes where airdensities are high, and is achievable at higheraltitudes by flying at faster speeds using inletprecompression. Historically the turbine engine wasthe key to high altitude flight, since it was the firstpowerplant with high enough specific power to pushlevel flight into the supersonic range. The turbineengine’s higher fuel consumption, 1.5 to 2 times thatof the reciprocating engine, is not a disadvantage formost aircraft applications because of the higherspecific power.

The acknowledged altitude records for subsonicflight, (shown in Table I), are dominated by turbinepowered aircraft. The highest is held by the Viet Namera AQM91 Compass Arrow spyplane, which achievedbetter than 80 kft more than 25 years ago [1].Powered by a special design turbojet engine (shownin Fig. 7) this aircraft achieved its record altitude flying

at M = 0.83, a speed which was just enough to givethe inlet precompression needed to keep itscombustor lit. The turbine engine exhibits a specificpower that varies roughly proportional to ingested airdensity since machine size is fixed. As density dropsoff at higher altitudes, the machine ingests less airmass, resulting in reduced power and reduced thrust. Combustor pressure is correspondingly reduced;eventually to the point where combustion of hydro-carbon fuel can no longer be supported. Figure 7shows the J97’s thrust lapse curve, whose behavior istypical of all turbine engines. As an example, theCompass Arrow’s J97 turbojet which was capable of>4,000 lbf thrust at sea level, would produce only184 lbf at 80 kft (M = 0.85) and is operating on theverge of flameout.

The high altitude J97 engine never went intoproduction because Compass Arrow was ultimatelycanceled. There remain twenty-four (24) J-97 pre-production prototype units (the engine was never fullyqualified) which were surplussed to NASA followingthe Air Force’s decision not to pursue systemacquisition. These are in storage at NASA AmesResearch Center.

It would be possible to design a new turbojetengine for flying beyond the J97’s ~80 kft limits usingpresent day materials and turbine technology. Thisengine would incorporate a high pressure ratiocompressor (25 to 35 to 1) with wide chord first stageblades (to minimize Reynolds number effects) and astabilized pilot flame combustor (using a secondaryfuel such as hydrogen) to prevent flameout at highaltitudes. Figure 8 shows the preliminary conceptlayout for such an engine. Compared to the J97, thisnew engine has larger flow areas (larger diameter),more turbo-machinery stages, higher overall pressureratio, higher flameout altitudes, and more thrust. AsFig. 8 shows, this engine would be capable ofaltitudes up to ~90 kft at subsonic speeds. Becausethese adaptations for the >80 kft flight regime make itlarger and heavier than the J97, it would not becompetitive with other jet engines for most missions.

Unfortunately, costs preclude development of anew turbine engine. Compass Arrow’s J97, which wasa Cold War scaling/adaptation from the larger GE-1,cost approximately $60M to develop during the mid1960’s. This is a sum roughly equal to $300M today. The market anticipated for atmospheric scienceaircraft consists of no more than a dozen units. As aresult, the only turbine engine that would be availablefor a new atmospheric science aircraft would be a J97unit rebuilt from the remaining inventory of prototypehardware that never became a manufacturersupported product.

NASA/TM—1998-206636 4

PROPULSION SYSTEM SELECTION FORATMOSPHERIC SCIENCE AIRCRAFT

These considerations leave the turbochargedreciprocating engine as the only remaining candidate,since it is the only low cost candidate. The turbo-charged engine is low cost, because of the existingtechnology base of mass produced automotive andgeneral aviation hardware that can be adapted tobuild such an engine. The technology is widely avail-able and well supported. Recent trends in automobilemanufacture to reduce weight (improve fuel economy)have rendered this technology base more applicableto aircraft propulsion, to the extent that many generalaviation home builders have developed automotivepowerplant conversions that are weight competitivewith certified aero engine installations. There is amarketplace for turbocharged engines that alreadyincludes a number of small business developers whomainly modify and assemble hardware manufacturedby others (for auto racing, experimental and homebuiltaircraft etc), some of whom might find a nicheparticipating in the development/manufacture/service/support of specialized turbocharged powerplants forhigh altitude unmanned aircraft (potentially aprofitable niche for commercial HALE platforms, a verysmall niche for atmospheric science aircraft).

The turbocharged powerplant is cheaper becauseit is built up from mass produced components fromother applications which have been adapted. Thedevelopment costs are low because of theconsiderable design heritage that survives from pre-jetage aviation development, and experience gainedfrom earlier attempts to develop turbocharged highaltitude powerplants. Several multiple stageturbocharged systems have already beendemonstrated either in high altitude flight or altitudetest chambers. Table II (previous high altitudeturbocharged IC engines) summarizes the flightcapability/performance that has been achieved todate.

The turbocharged reciprocating engine istechnically quite competitive with the turbine engine athigh altitudes. Although the power density of a turbineengine is higher than the reciprocating engine atnormal altitudes, the reciprocating engine begins tocompare favorably with turbine engine at altitudesabove 80 kft, especially at the lower speeds whereinlet precompression is not available. While it must beaugmented with multiple stages of turbocharging andintercooling for atmosphere pressure/density compen-sation, the reciprocating engine’s weight growth withaltitude is not as rapid as the pure turbine engine’s;mainly because the turbomachinery needed to raiseOPR is confined to the induction air, which is less than1/10th the overall air consumption.

The highest altitude potential accrues to the sparkignited gasoline engine since it burns a nearlystochiometric fuel air mixture, thus maximizinginduction air utilization. Furthermore, its exhaust gasesare hot enough (1400 to 1600 °F) to yield excessenthalpy which is needed to provide the turbochargercompressor work. Intake pressurization isaccomplished by multiple stage units arranged incascade, so that as altitude increases and ambientpressure decreases, the increasing pressure ratioacross the turbine increases enthalpy extraction,roughly balancing the increased compressor loading.

The power density of the turbochargedreciprocating engine is limited by core enginedetonation limits not the turbomachinery, so theresulting curve of performance with altitude (shown inFig. 9) is “flat” extending from sea level to criticalaltitude where the turbo-machinery was sized (todeliver rated intake airflow and pressure). Abovecritical altitude the turbo-machinery can no longersustain these airflows, so the performance curveexhibits a lapse behavior similar to the turbojet. As adesign parameter, critical altitude shouldapproximately coincide with the aircraft’s designaltitude and not exceed it, since the high altitudepower generation capability so dearly paid for inpropulsion weight is wasted beyond that point.

Figure 10 shows specific air consumption of fiveair breathing engine types. Of these, the spark ignitedgasoline engine has the lowest specific air consump-tion. Owing to its near stochiometric combustion, itutilizes all the air which is processed—a majoradvantage where air processing makes up most ofthe powerplant. In addition, the turbocharged reci-procating engine retains lower fuel consumption.Table III compares the specific weight and thrustspecific fuel consumption of representative turbojetand turbocharged powerplant installations at 80 and90 kft. Specific weights apply to the entire propulsionunit (including drivetrains, propellers, heat exchangers,etc.) and flight speeds are chosen to provide eachpowerplant with its inherent competitive advantage. Comparison shows that at 80 kft the turbochargedpropeller unit will be slightly heavier per lbf thrust thanthe turbojet, but it will have lower thrust specific fuelconsumption (TSFC). If the comparsion is repeated at90 kft, the turbocharged unit is somewhat lighter thanthe turbojet per lbf thrust and has lower TSFC. Thesesame trends were also observed when turbochargedreciprocating engines were compared with specializedturboprop units designed for 90 kft [2].

These performance advantages may be exploitedto a limited extent, but due to the increasing size andweight of ancillaries required to maintain altitudeperformance, the powerplant eventually becomes tooheavy to be supported by the wing loading available.

NASA/TM—1998-206636 5

In order to illustrate how size and weight areincreasing, Table IV presents a weight breakdown fora triply turbocharged >80 kft altitude powerplantbased on a Rotax engine core. This breakdownshould be compared to an equivalent weight break-down for the single stage turbocharged Rotax enginethat is designed for 15 kft. NASA airframe/propulsion studies [3] indicate that for subsonicairframes using modern composite materials, themaximum altitude potential for subsonic flight will belimited by powerplant growth to about 90 kft (seeFig. 11, the no fly zone).

Reliability of the turbocharged reciprocating enginewill be lower than a turbine engine since it is physicallymore complicated. Operational reliability of a systemconsisting of so many interconnected elements is asignificant issue. Historically, turbocharged piston aeroengines have required regular maintainance overoperating intervals measured in tens of hours, andcomplete overhauls after hundreds of hoursoperation. This contrasts with modern turbine engineswhich require maintainance after hundreds of hours ofoperation, and thousands of hours between over-hauls. A powerplant developed for the atmosphericscience mission will be very limited production; soreliability issues associated with inherent systemcharacteristics are more likely to be overshadowed bythe teething problems associated with few-of-a-kindsystems. Since a remotely piloted aircraft gives theoperator only limited ability to detect problems in flightand even less ability to respond to them, low reliabilitymeans a higher likelihood of loss for both mission andaircraft. These risks may be tolerable if humanoperators are no longer in harm’s way.

CURRENT PROPULSION DEVELOPMENTS

Since the ERAST program was inaugurated in1994, NASA has pursued development of a threestage turbocharged gasoline engine as the propulsionoption for its very high altitude atmospheric scienceaircraft. NASA is currently developing a subscaleprototype propulsion unit intended for an unmannedflight demonstration known as Alliance One. Unlikethe German Grob/DLR “Strato 2C” atmosphericscience aircraft ([4 and 5] see Table I), Alliance One isintended to be a technology demonstrator addressing>80 kft subsonic flight, not a mission demonstration.

Present development is focussed on a three stageturbocharged powerplant using the four cylinder Rotax912 engine core. Popular with remotely piloted aircraft(RPA) and “homebuilt” experimental aircraftdevelopers, this engine is in volume production andhas factory technical support available. Its versatiledesign accommodates both air and liquid coolingsuch that it can be all liquid cooled for high altitudes or

fully enclosed installations. The three stage turbo-charger system is being developed by Thermo-Mechanical Systems (TMS) of Canoga Park,California, the original developers of the TEAL RAINthree stage turbocharged engine ([6], see Table II). To save money, much of the original TEAL RAINhardware has been re-utilized. Figure 12 is a photo-graph of the triply turbocharged Rotax engine in theTMS chamber, Fig. 13 shows the power levels atvarious altitudes which have been demonstrated bythis engine to date (exceeding the original programgoals). These demonstrations are an importantmilestone, but will not lead immediately to a highaltitude flight since it is only a brassboard demon-stration of critical hardware, not the entire propulsionunit which has yet to be developed. So far, ratedhorsepower has been produced in a dynamometeraltitude chamber under quasi-steady state conditions,controlled by hand. Critical subsystems still missingfrom the demonstration include thermal management,transmission, drivetrain and propeller, powerplant andpropulsion unit automatic controls.

Work is already underway on many of thesesubsystems, with thermal management receiving themost attention. High altitude low Reynold numberheat exchangers are presently being researched byNASA Lewis and a consortium of four heat exchangermanufacturers led by the Ohio State UniversityResearch Foundation. Nacelle and inlet aerodynamicsare being researched by groups at NASA and OldDominion University. Definition of the propeller hasalso begun, led by Alliance partner Aurora FlightSciences. Drivetrain and propeller development isconsidered a unique challenge since the propeller,despite its being a variable pitch unit, will be subjectto greater than 2:1 speed variation when traversingfrom sea level to >80 kft. Aurora Flight Sciences, Inc.is developing a two speed transmission and drivetrainto complete the propulsion unit. Automatic controlsdevelopment, a natural part of the integrationprocess, is proceeding with turbocharged engine/heatexchanger integration. TMS has developed automaticpowerplant controls for the throttle, intercoolers andturbocharger wastegate using the engine controlcomputer’s expansion capability. The control actuatorsthemselves are developed by General Atomics.Aurora is developing a supervisory propulsion controlsystem that will interface the powerplant, transmissionand propeller with the aircraft flight controls.

After all this propulsion hardware has beendeveloped and extensively ground tested to ensure it“works as advertised,” it will be integrated into theAlliance One airframe leading to the ultimate objectiveof ERAST propulsion development: flightdemonstration of affordable technology to meet thescience mission requirements.

NASA/TM—1998-206636 6

SUMMARY AND CONCLUSIONS

The atmospheric science mission requirementsdictate a very high altitude subsonic aircraft. Thisaircraft requires a propulsion system which presentlydoes not exist, but can be developed from airbreathing engine technology. There are twocandidates: turbojets and propeller/turbochargedspark ignited reciprocating engines. Although theturbojet would appear to be the leading candidate, asit was for high altitude missions historically, there is noturbine engine available for the atmospheric scienceflight regime, and it is unlikely that a suitable turbineengine will ever be developed, due to the limitedmarketplace associated with atmospheric scienceapplications.

Therefore, the propeller/turbocharged spark ignitedgasoline engine is the only option available. Fortunately, the spark ignited powerplant enjoys somesubtle physical advantages over the turbojet for highalitude low speed flight, arising from its near-stochiometric combustion. Its low specific airconsumption reduces the amount and weight ofturbomachinery required to generate power at >80 kft,which apparently results in lower weight and lowerTSFC than a turbojet in this flight regime, despite thelarge heat exchangers.

Borrowing from the technology heritage ofautomotive and homebuilt/experimental aircraftpowerplants and Cold War era high altitude militaryreconaissance development programs, NASA is nowdeveloping a variant of the 3 stage turbochargedpowerplant for application to the atmospheric sciencemission. To date, only the engine and turbochargerperformance has been demonstrated, under carefullycontrolled conditions. Work is ongoing to develop theremaining required components, and build thedemonstration hardware into a propulsion systemwhich will lead to an aircraft technology flight feasibilitydemonstration.

REFERENCES

1. W. Wagner and W. Sloan, “Fireflies and OtherUAV’s,” Teledyne Ryan Aeronautical Co.,San Diego CA, June 1993.

2. J.L.Bettner, C.S. Blandford and B.J. Rezy,“Propulsion Assessment for Very High AltitudeUAV Under ERAST,” NASA Report 95N27866,Allison Engine Company, June 1995.

3. Report of the Environmental Research Aircraft andSensor Technology (ERAST) Program Leader-ship Team, “A Review of Remotely PilotedAircraft (RPA) Technology Required for HighAltitude Civil Science Missions,” NationalAeronautics and Space Administration,Washington DC, March 1996.

4. H. Tonksotter, “The Strato 2C Propulsion System;A Low Cost Approach for a High Altitude LongEndurance Aircraft,” Industrieanlagen-Betriebsgesellschaft mbH, March 1994.

5. Anon, “Strato 2C Technical Description”, DeutscheForschungsanstalt fur Luft und Raumfarht (DLR) , November 1993.

6. J. Harp, “Turbocharger System Development andPropulsion System Testing:, TMS Report No.SR-36, prepared for Developmental SciencesInc. under Contract No. DSI-80-TR-SC-05-A,ThermoMechanical Systems Inc., Canoga ParkCA, May 1982 (declassified Mar 1994).

NASA/TM—1998-206636 7

TABLE I.—PREVIOUS HIGH ALTITUDE SUBSONIC AIRCRAFTAircraft designation Original purpose, and

year flownAltituderecord

Propulsion systemused

Science platformavailability

WB–57 High altitude strategicbomber (1949)

65,876 ft8/29/55

Bristol“Olympus” turbojet

NCAR atmosphericscience

ER–2 High altitudereconnaissance (1955)

73,200 ft8/4/95

GE-F118turbofan

NASA atmosphericscience

AQM91MCompass Arrow

High altitudereconnaissance (1969)

>81,000 ftSept. 1969

GE-J97-3turbojet

military only, nolonger exists

Grob Egrett High altitude scienceaircraft (1988)

53,055 ft9/1/88

Garrett TPE331turboprop

DoE atmosphericscience

Boeing Condor High altitude ELINTreconnaissance (1989)

67,028 ft2/15/89

2 stage turbochargedspark ignition engine

military only, nolonger exists

Grob Strato 2C Atmospheric science(1995)

60,867 ft8/4/95

3 stage turbochargedspark ignition engine

DLR (Germany)program was

canceled

TABLE II.—TURBOCHARGED RECIPROCATING ENGINESDeveloper Core engine

usedNumber of

stages/turbomanufactor

Rated hp demoat rated alt.

Highest recordedground test

performance

Highest altitudeachieved in flight

TEAL RAIN ThermoMechanical

Systems(TMS)

3 cylinderDrake 36.6 cid

3 stages/TMS 70 hpat 65 kft

Feb. 1982

47 hpat 90 kft

Mar. 1982

ground demoonly

Condor BoeingTeledyne

ContinentialMotors

6 cylinderContinental

350 cid

2 stages/Teledyne

continental

182 hpat 67 kft

Feb. 1989

data notavailable

67,028 ftFeb. 1989

Strato 2C Grob/IABG/DLR

6 cylinderContinential

550 cid

3 stages/IABG/P+W/

Garrett

400 hpat 78 kft

Dec. 1994

308 hpat 82 kft

Apr. 1995

60,876 ftAug. 1995

Raptor D2 ScaledComposites/

TMS

4 cylinderROTAX 74 cid

2 stages/TMS 103 hpat 54 kft

Jan. 1996

47 hpat 70 kft

Jan. 1996

Not flown yet

Perseus B/Theseus Aurora FlightSciences

4 cylinderROTAX 74 cid

3 stages/Garrett 73 hpat 59 kft

May 1994

73 hpat 59 kft

May 1994

20,000 ftMar. 1996

Altus GeneralAtomics

AeroSystems/TMS

4 cylinderROTAX 74 cid

2 stages/TMS 103 hpat 54 kft

Jan. 1996

47 hpat 70 kft

Jan. 1996

Not flown yet

ERAST Alliance IProof of Concept(P.O.C)

ThermoMechanical

Systems(TMS)

4 cylinderROTAX 74 cid

3 stages/TMS 100 hp at 76 kft

Feb. 1997

85 hpat 85 kft

Nov. 1997

Not flown yet

NASA/TM—1998-206636 8

TABLE III.—TURBOJET VERSUS TURBOCHARGED IC ENGINEPowerplant/Propulsion system

(inlet recovery = 1.0)Uninstalled weightincluding propeller

Deliveredthrust

Specificweight

Specificconsumption

At 80 kftJ97 turbojet at M = 0.83J97 turbojet at M = 0.8New turbojet at 0.5<M<0.85Strato 2C (3 stage TCSI) at M = 0.580K ERAST 3 stage TCSI at M = 0.4

715 lbm715 lbm920 lbm2457 lbm587 lbm

184 lbfflameout190 lbf360 lbf91 lbf

3.9 lbm/lbf– – – – – –4.8 lbm/lbf6.8 lbm/lbf6.5 lbm/lbf

1.3 lbm/hr per lbf– – – – – –  – – –0.8 lbm/hr per lbf

0.44 lbm/hr per lbf0.44 lbm/hr per lbf

At 90 kftNew turbojet at 0.5<M<0.8590 K ERAST 3 stage TCSI at M = 0.4

920 lbm667 lbm

120 lbf90 lbf

7.7 lbm/lbf7.4 lbm/lbf

0.8 lbm/hr per lbf0.44 lbm/hr per lbf

TABLE IV.—TURBOCHARGED POWERPLANT WEIGHT BREAKDOWN COMPARISON OF “UNINSTALLEDHARDWARE” NO PROPELLER

Rotax 914 3 stage turbocharged ROTAXRated HP at altitude 80 hp at 12,000 ft 80 hp at 85,000 ft

Weight breakdownCore engine w/ignition and fuel systems,induction, exhaust collector andwastegate

weight = 145 lbm weight = 184 lbm

Turbochargers (no of units) 1 TC unit—7 in. od× weight = 15 lbm

LP unit—20 in. odweight = 60 lbmIP unit—13 in. odweight = 30 lbmHP unit—7 in. odweight = 11 lbm

Ducting and insulation weight = 4 lbm weight = 72 lbmIntercoolers 7 × 5 × 2 in.

weight = 4 lbmLP coolers: 12 × 56 × 2-1/2 in.weight = 2 × 7.1 lbmIP coolers: 12 × 69 × 2-1/2 in.weight = 2 × 9.7 lbmHP coolers: 12 × 39 × 2-1/2 in.weight = 2 × 5.8 lbm

Oil cooler 7 × 3 × 1-3/8 in.dry weight = 5 lbm

oil to coolant HX: 9 × 9 × 2 in.dry weight = 6.7 lbm

Engine coolant radiator 8 × 4 × 1 in.dry weight = 7 lbm

42 × 44 × 3/4 in.dry weight = 13 lbm

Gearbox 1 speed, integral w/core engine 2 speeds,55 lbm

Lube system include oil tank andancillaries

dry weight = 15 lbm dry weight = 25 lbm

Liquid coolant weight = 5 lbm weight = 22 lbmOil weight = 4 lbm weight = 18 lbmTotal weight 204 lbm 543 lbm

NASA/TM—1998-206636 9

Figure 1.—Pressure ratio needed for flat-rated output power and relative wing lift versus altitude for constant flight Mach number of 0.50.

NASA/TM—1998-206636 10

Figure 2.—Turbocompressor size variation with altitude.

LAC

Dim

ensi

ons

rela

tive

to s

ea le

vel i

nlet

val

ues

8

7

6

5

4

3

2

1

LCC

3 Stages

2 Stages

1 Stage

DAC,DCC

Inlet altitude (kft)0 10 20 30 40 50 60 70 80 90 100

Sizing for altitude compensation (constant mass flow rate)

Axial Compressor10 - 100 lbm/sec

(pressurizes combustor to 40 psia)

Diameter DAC

Length LAC

Centrifugal Compressor0.1 - 0.5 lbm/sec

(pressurizes engine intakemanifold to 30 in Hga)

Diameter DCC

LengthLAC

NASA/TM—1998-206636 11

Figure 3.—Heat exchanger size variation with altitude.

Figure 4.—Approximate capture area and diameter required for 100 lbf thrust versus altitude, airspeed and desired propulsive efficiency.

NASA/TM—1998-206636 12

Figure 5.—High altitude turbojet.

Ambient air

Compressor

Inlet

LPC

HPC

Combustor

Compressor Turbine

——High Pressure——

——Low Pressure——

Turbine

HPT

LPT Exhaust

Nozzle

Figure 6.—Three Stage Turbocharged Powerplant.

Wastegate

ThrottleFourcylinder

coreengineTrans-

mission

Ambient air

Ambient air

Ambient air

Ambient air

Ambient air

Variable area nozzle(v.a.n.)

After-cooler

Inlet/diffuser

HP turbocharger

IP intercooler

Inlet/diffuser

IP turbocharger

IPC

HPC HPT

IPT

LP intercooler

(v.a.n.)Inlet/diffuser

LPC LPT Exhaust

Fixed nozzle

LP turbocharger

Inlet/diffuser

Inlet/diffuser

Enginecoolantradiator

Oil andcoolant

heat loads

Variablearea

nozzle(v.a.n.)

(v.a.n.)

NASA/TM—1998-206636 13

Figure 7.—"Compass arrow" high altitude turbojet (1966-72), thrust versus altitude.

Fn, lbs

80 000

70 000

90 000

60 000

50 000

40 000

30 000

20 000

10 000

5000

WaSLS = 66.2 lbm/sOPRSLS = 11.5FnM = .83 at 80k = 184 lbfSFCM = .83 at 80k = 1.298Lengthbare = 110 in.Weightdry = 694 lb

Performance Size

400030002000100000

Alti

tude

, ft

GE J97-3

(Maximum available thrust)

Classified

Compass arrow

NASA/TM—1998-206636 14

Figure 8.—New high altitude turbojet, thrust versus altitude.

Net thrust, Fn, lbf

Preliminary turbojet

Thrust lapse100 000

80 000

60 000

40 000

20 000

700060005000400030002000100000

Alti

tude

, ft

21"

63"

WaSLS = 86.6 lbm/sBPR = 0OPR = 30Fn90k = 120 lbfWt = 902 lbmSFC90k = .815

NASA/TM—1998-206636 15

Figure 9.—Turbocharged engine performance.

hp

100

90

80

70

60

50

40

30

20

10

0

Power requiredto fly

Delivered propshaft horsepower

Altitude (kft)S.L. 10 20 30 40 50 60 70 80 90 100

Figure 10.—Specific air consumption dictates turbomachinery size.

NASA/TM—1998-206636 16

Figure 11.—Turbocharged powerplant growth trend with altitude.

Figure 12.—TMS triply turbocharged Rotax engine.

NASA/TM—1998-206636 17

Figure 13.—TMS three stage turbocharged engine performance demonstration.

Dyn

amom

eter

hor

sepo

wer

110

100

90

80

70

60

50

40

30

20

10

0

run #58107/03/97

run #15011/20/96

run #27602/18/97

run #27802/18/97

run #28102/18/97

run #28402/18/97

run #28602/18/97 run #290

02/18/97

run #29402/18/97

run #57007/03/97

run #57407/03/97

run #15111/20/96

run #15211/20/96

run #15311/20/96

run #15411/20/96

run #15511/20/96

run #15711/20/96

run #50206/19/97

run #41805/13/97

run #58007/03/97

run #57907/03/97

run #57707/03/97

S.L.Chamber altitude (kft)

10 20 30 40 50 60 70 80 90 100

Operating Envelope

This publication is available from the NASA Center for AeroSpace Information, (301) 621–0390.

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National Aeronautics and Space AdministrationLewis Research CenterCleveland, Ohio 44135–3191

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National Aeronautics and Space AdministrationWashington, DC 20546–0001

April 1998

NASA TM—1998-206636

E–11101

WU–529–10–13–00

23

A03

Propulsion System for Very High Altitude Subsonic Unmanner Aircraft

David J. Bents, Ted Mockler, Jaime Maldonado, James L. Harp, Jr.,Joseph F. King, and Paul C. Schmitz

High altitude; Air breathing propulsion; Turbine engines; Reciprocating engines

Unclassified -UnlimitedSubject Category: 07 Distribution: Nonstandard

Prepared for the Aerospace Power Systems '98 sponsored by the Society of Automotive Engines, Williamsburg, Virginia,April 21–23, 1998. David J. Bents, Ted Mockler, and Jaime Maldonado, NASA Lewis Research Center; James L. Harp, Jr.and Joseph F. King, ThermoMechanical Systems Inc., Canoga Park, California; Paul C. Schmitz, Power ComputingSolutions Inc., Avon, Ohio. Responsible person, David J. Bents, organization code 5440, (216) 433–6135.

This paper explains why a spark ignited gasoline engine, intake pressurized with three cascaded stages of turbocharging,was selected to power NASA’s contemplated next generation of high altitude atmospheric science aircraft. Beginningwith the most urgent science needs (the atmospheric sampling mission) and tracing through the mission requirementswhich dictate the unique flight regime in which this aircraft has to operate (subsonic flight @ >80 kft ) we briefly explorethe physical problems and constraints, the available technology options and the cost drivers associated with developing aviable propulsion system for this highly specialized aircraft. The paper presents the two available options (the turbojetand the turbocharged spark ignited engine) which are discussed and compared in the context of the flight regime. Wethen show how the unique nature of the sampling mission, coupled with the economic considerations pursuant to aeroengine development, point to the spark ignited engine as the only cost effective solution available. Surprisingly, thissolution compares favorably with the turbojet in the flight regime of interest. Finally, some remarks are made aboutNASA’s present state of development, and future plans to flight demonstrate the three stage turbocharged powerplant.