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NASA Technical Memorandum 81584
NASA - TM - 81584 19810001559
C o mparis o ns o f F o ur Alternative
P o werplant Types for Future
General Aviation Aircraft
T . J. Wickenheiser , G . Knip , R. M. Plencner , an d W. C. Strack
Lewis Research Center
Cleveland , Ohio
October 1980
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ABSTRACT Recently completed NASA-sponsored conceptual studies have culminated in the identification of promising new technologies for future spark ignition, diesel, rotary, and turbine engines. This paper reports the results of a NASAin-house preliminary assessment study that compares these four powerplant types in several general aviation applications. The evaluation consisted of installing each powerplant type in"rubberized"aircraft which are sized to accomplish fixed missions. The primary evaluation criteria include projected aircraft cost, total ownership cost, and mission fuel.
INTRODUCTION Four recent NASA-sponsored conceptual studies have identified several promising engine technologies applicable to the 9eneral aviation field. Advanced light- weight diesels (ref. I), stratified-charge rotaries (ref. 2), inexpensive gas turbines (ref. 3-7), and spark-ignited reciprocating engines (ref. 8) were defined and then compared individually to current-technology reciprocating engines for several general aviation applications. Each study concluded that major improvements were possible. But because only a few applications were considered in each study, and the mission definitions varied from one study to the others, it is difficult to compare these alternative powerplant types with each other on a consistent basis. The results presented herein address that issue and are based principally on engine characteristics supplied by the competing engine companies. NASAassumptions were used only when required information was otherwise unavailable.
The results presented in this paper are preliminary in the sense that NASA has not yet attempted to resolve seemingly inconsistent or controversial assumptions, particularly in the level of technology advancement represented.
The contractors assumed different levels of technology within engine types, thus, defining comparable technology and relative risk between engine types is further complicated. No attempt is made to define the inter-engine relative technology levels at this time. The results, then, represent only the first phase in a continuing assessment process. As time progresses, further defini- tion of the improvements and risks involved in each engine type will require re-evaluation of the engine assumptions. In fact, this study was undertaken primarily to generate parametric sensitivity data to facilitate later modifi- cations--either by NASAor others--under differing scenarios. Hence, at this stage, no conclusions are offered concerning the relative attractiveness of the competing engine types. Though overall trends are sometimes apparent, the engine ranking is subject to considerable shifting due to the preliminary nature of the engine performance, weight, and cost assumptions.
The engines were evaluated for eleven fixed-wing and two rotary-wing applica- tions. This evaluation consisted of installing each engine type in a "rub- berized" aircraft which is sized to accomplish each fixed mission. Three
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figures of merit were used to compareengine performance: aircraft acquisi- tion cost, mission fuel, and five-yeartotal ownershipcost (TOC). TOC is based on 500 hr / yearutilizationand is defined in Table I.
SYMBOLS AC acquisition cost, 1977 $ BSFC(SFC) brake specific fuel consumption, Ib / HP-hr C d cooling drag, % aircraft drag Ce engine cost, 1977 $ / HP Cf fuel cost, 1977 S / gal.
L / D lift-to-drag ratio MC maintenance cost, $ / flight-hr OEM original equipment manufacturer TBO time between overhaul, flight-hr SHP shaft horsepower, HP SIR spark ignited reciprocating engine W e engine weight, Ib / HP Wf mission fuel weight, Ib ANALYSlS Engine Technology Levels Three terms will be used to differentiate different versions within each engine type: current, advanced and very advanced technology. Current technology is defined as circa 1980 production engines, except for the diesel and rotary where no aircraft production engines are available (prototypes were used in- stead). Since no relative technology level assessment has been completed between engine types, the terms advanced and very advanced technology define relative technology levels within particular engine types, but do not neces- sarily represent comparable technology levels between types. Aggressive research programs would be required to enable production versions of these engines by the early 1990's.
Engine Design Assumptions Spark-lgnited Reciprocating Engines - The engine design features for the Spark- Ignited Reciprocating (SIR) engine (ref. 8) are defined in figure I. The current technology version serves as the baseline engine from which the improvements of all of the other engines are measured. This current technology engine is natu- rally aspirated below I0,000 foot cruise altitude and turbocharged above I0,000 foot. The advanced and very advanced versions are both turbocompounded. Since a turbocharger is necessary to achieve superior performance, no naturally- aspirated version was considered.
The current and advanced SIR engines use a homogeneouscharge of aviation gasoline while the very advanced version uses a stratified-charge, which provides multifuel capability. This capability becomes increasingly important as aviation gasoline becomes scarcer and more expensive.
The cooling drag penalty for the current technology SIR was assumed to increase the aircraft drag by I0 percent. Reference 9 indicates cooling drag to be 5-20 percent of aircraft drag. A 50 percent reduction in SIR cooling drag is predicted in reference I0 due to better cooling techniques and better nacelle-cooling system integration. This 50 percent reduction was assumed for the advanced technology version, while a 65 percent reduc- tion was assumed for the very advanced version relative to a current SIR.
Rotary Engines - The engine design features for future rotary engines (ref. 2) are shown in figure 2. The major emphasis of the rotary study was directed at an advanced technology version, though a lower-level effort did postulate a more advanced version, herein labeled very advanced. Both versions used a stratified-charged fuel injection system that provides multifuel capability.
The advanced version is supercharged, while the very advanced version is pres- sure compounded. Both versions require a turbocharger to achieve high perform- ance and both are liquid cooled. The advanced version is assumed to have 65 percent less cooling drag than a current SIR, while the very advanced version is assumed to have negligible cooling drag.
One of the design features of the very advanced rotary is the retracting apex seal. This retracting seal allows higher engine RPMwithout excessive seal wear. Higher RPMpermits higher airflows and thus higher horsepower from the same size engine.
Diesel Engine - The engine design features for the lightweight diesel (ref. I-) are shown in figure 3. Both the advanced version and the very advanced version are two-stroke radial designs having the same cooling requirements as comparable technology SIR versions.
One of the unique design features of these diesels is the independent turbo- charger loop. Along with the turbocharger compressor and turbine, a burner is also provided with the necessary ducting to allow the airflow to bypass the diesel and thereby permit the turbocharger to act as an independent auxiliary power unit. While this feature adds cost and complexity to the engine, it also offers some significant design improvements. This indepen- dent loop permits the turbocharger to provide auxiliary power while on the ground without the necessity of starting the entire powerplant. Many of . the starting problems (hot, cold, and restart) associated with diesel engines are eliminated by first starting the turbocharger loop to preheat the air.
Since high diesel compression ratios are normally required only for accept- able starting performance, the independent turbocharger loop allows the engine to be designed for much lower compression ratios. The lower compression ratio results in lower stresses and lighter recip-related component weights.
The turbocharger is a key component in the diesel design. It is required to operate well beyond current turbocharger capability. With the high turbo- charger pressure ratio, the engine exhaust air does not contain sufficient energy to power the turbochargerabove 17,000 foot cruise altitudes. To eliminateexcessivethrust lapse at high altitudesit is necessaryto burn fuel in the turbochargerloop to add additionalenergy to the turbine inlet air. A correspondingBSFC penaltywas charged to the engine as shown in figure 4.
Turbine Engines - The engine design featuresfor the inexpensivegas turbine are shown in figure 5. The turbine engines used in the present study!are representative of those proposed by the General Aviation Turbine Engine (GATE) contractors (refs. 4-7). These studies addressed only one level of technology, herein termed advanced technology, with the major effort aimed at engine cost reduction, while still achieving some performance improvement. The turbine engine is assumed to take advantage of the 40 percent cost reduction projected for GATElow-cost manufacturing techniques with another 25 percent cost reduc- tion predicted for large production rates (around I0,000 engines / year per manufacturer}.
An additional cost reduction method considered in the GATEstudies was the common-core engine family. The core engine has a single-stage, 9:1 pressure •ratio centrifugal compressor driven by a single-stage radial turbine. As the horsepower requirements increase a single-stage axial compressor, a single-stage axial turbine as well as an identical set of gearbox components are added to obtain a 12:1 pressure ratio machine with about 70 percent higher horsepower.
Only one level of technology is presented for the turbine engine. Though a more advanced version can be configured, its relationship to the low-cost GATEtheme is difficult to determine. An expensive turbine engine could be cost prohibitive in light general aviation airplanes.
Installation Penalties Engine installation penalties are shown in Table 2. The SIR and rotary engines are assumed to obtain the air conditioning and pressurization requirements from the turbocharger and no penalty was assessed to the engine. The diesel, with its high turbocharger design pressure ratio, did not have sufficient energy in the engine exhaust gas to supply the air conditioning and pressurization re- quirements for the entire flight regime, thus a 4 percent horsepower penalty was charged to the diesel. The turbine engine installation penalties are assessed differently from its positive displacement counterparts. While the power requirements for the alternator, fuel pump and oil pump are similar, the turboprop air conditioning and cabin pressurization requirements are not obtained from a turbocharger or as a horsepower penalty from the shaft.
Instead, the turbine engine meets these requirements using compressor dis- charge air which causes significant BSFCpenalties. This pressurization require- ment need not take this form; instead, an auxiliary shaft-driven compressor or midstage bleed could be utilized to reduce the BSFCpenal tymarkedly. However, it is assumed herein that compressor discharge bleed will be used to satisify the air conditioning and pressurization requirements (I to 1½ Ib / min / pass com- pressor discharged bleed was assumed), Engine P erformance Assumptions BSFC Assumptions- Projectedcruise BSFC for each engine type, along with the corresponding percent improvement relative to their current technologyver- sions are shown in figure 6,for 350 thermodynamic SHP at sea level static standardconditions. A 20 percent BSFC improvement over current state-of-the- art is projectedfor the advanced turboprop. This improvementis primarilydue - to improvedcomponentefficiencies and more efficientthermodynamic cycle (made possible by material improvements and lower cooling requirements). The BSFC for the advanced SIR is projectedto improve20 percentas a result of higher design brake mean effectivepressure (BMEP),more efficientcombustion, and turbocompounding.A further 5 percent BSFC improvementis projectedfor the very advancedversiondue to charge stratification.A 20 and 30 percentBSFC improvementis projectedfor the advancedand very advanced rotary engines.
In the advancedversion, this improvementis due to better combustion(stratified charge), higher BMEP and apex seal improvements. The BSFC improvements for the very advancedversion are a result of the above with the additionaladvantage of pressurecompounding. The advanceddiesel's 8 percent improvement is a result of more efficientcombustionand reducedfriction losses,while the 20 percent improvementprojectedfor the very advanced version is a result of the added benefitsof higher BMEP and reduced coolinglosses. The diesel BSFC accounts for fuel burned in the turbochargerloop at high altitudes. Below 17,000 feet this fuel burning is not necessaryand the diesel BSFC is lower (figure4). The assumed effectsof engine size on BSFC are shown in figure 7 as a functionof rated power (for turbine enginesthis is assumedto be iden- tical to "thermodynamic horsepower", that is, maximum deliverablepower without gearbox or any other limitations). All positivedisplacementenginesare assumed to follow the same BSFC trends with horsepower. The turbopropBSFC is much more sensitiveto size effects than positivedisplacement engines in this horsepower range due to limitationson end wall clearancesand manufacturing tolerances, and the dominanceof the boundry layer in the flow path. TurbopropBSFC is also dependenton speed and altitude. The turbopropBSFC curve is regressed throughthe BSFC points used in the study as calculatedat each aircraftcruise condition(usinga NASA in-house thermodynamic design point code which accounts for speed, altitudeand size effects). The break in the turbopropcurve at 300 horsepoweris due to a change from a 9:1 overallpressure ratio below 300 horse- power to 12:l above in accordancewith the common core concept referredto earlier.
Engine Weight Assumptions - Engine installed specific wei g ht comparisons are shown in figure 8. The already lightweight turboprop is predicted to improve a relatively modest 20 percent due mainly to reduced airflow (higher specific power) resulting from higher cycle temperatures. The use of higher compressor and turbine loadings which permit fewer stages also contributes to lower weight.
The 17 and 30 percent projected weight reductions for the advanced and very advanced SIRaredue to turbocompounding, higher BMEPand lighter materials (titanium parts). The I0 and 40 percent weight reduction predicted for the advanced and very advanced rotary engines, respectively, are the result of higher BMEP, turbosupercharging, higher RPMand pressure compounding. Com- posite materials are also being considered for the very advanced rotary. The large 35 and 45 percent weight reductions for the advanced and very advanced diesel are a result of turbosupercharging, higher RPM, lighter materials, higher BMEPand elimination of the scavenging blower.
Currenttechnologyturbopropshave significantly lower specificweights than their positivedisplacementcounterparts. However,with the largeweight reductionsprojectedfor the rotary and diesel engines,the turbopropweight advantageis drasticallyreduced. The effect of engine size on specific weight is shown in figure 9. These curves are based on existing weight trends for all engine types except the diesel. The diesel sizing effectswere supplied by the authors of referenceI.
Engine Cost Assumptions- Although engine originalequipmentmanufacturer (OEM)cost estimateswere provided to NASA by the contractors, they were essentiallyrough estimatesbased on controversial assumptions. To avoid the appearanceof accuracywhere none exists, none of their projectedengine costs were utilized herein. Instead,as shown in figure lO, each engine was assumed to cost $40 / SHP (above250 SHP) except the naturally-aspirated cur- rent technologySIR which was set at $30 / SHP. Higher costs were assumedfor smallerengine sizes in conformance with existing cost trends. Interestingly, most of the contractors'estimateswere in the neighborhood of $40 / SHP including those for the turboprop (whichrepresentsa 65 percent decreasedue to advanced manufacturingtechniquesand larger productionrates). Differentcost assump- tions can be handledthrough the use of sensitivitycoefficients presentedin appendixA where examples are also provided.
Engine TBO and MaintenanceCost Assumptions- Time between overhaul (TBO) and engine maintenancecost assumptionsare shown in figures II and 12, respectively.
Since turbineengines have historically had longer TBO and lower maintenance cost than positivedisplacementengines,the majority of the GATE contractors did not specificallyinvestigatepossible improvements in these areas. The diesel contractordid investigate and projectimprovements in TBO and mainten- ance cost. The rotary contractordid not supply TBO and maintenanceinformation (exceptfor the TBO of the advanced version);NASA arbitrarilyassumed it to have the same requirementsas the diesel. Thus, when comparingthe projected TBO and maintenancecost of the advanced turbineenginesto the advanceddiesel and rotary engines, the tubropropadvantageis significantly lower than com- paring actual TBO and maintenancecost of current turbineengines to current positivedisplacementengines.
Severalother engine assumptionsare shown in Table 3. The diesel, rotary, and SIR engines have nominal criticalaltitudesof 17,000,20,000,and 21,000 feet respectively. The diesel and the SIR are assumedto cruise at 70 percentmaxi- mum rated power, while the rotary cruises at 75 percent power. Gas turbines are assumed to cruise at 2140OF (typcialGATE value). Engineswhich do not have multifuelcapabilitywere assumed to use aviation gasolineat $1.10 / gallon ($1977)while engines having multifuelcapabilitywere assumedto use Jet A fuel at $1.O0 / gallon. Forthe fixed wing missions,an 87 percent propellerefficiency was assumedwith SIR and diesel engines,while slightlyhigher propellereffi- ciencieswere assumed with rotary and turbine enginesdue to less engine vibra- tion.
Aircraft and Missions A plot of design cruise altitude and speed for the 13 missions is given in figure 13. Each mission is assigneda number for labelingpurposes. These missions are furtherdefined in Table 4 . Two groups of aircraftare presented: (1) Current--resembling today's aircraftand missions,but slightlybetter airframe technology,and (2) Futuristic--similar to Malcolm Harned'sspecula- tion (ref.ll) with 15 percent reductionin aircraftempty weight, 15 percent reductionin aircraftzero-liftdrag, engines submergedin the fuselage,higher wing loadingand aspect ratios, full-spanFowler flaps, and higher cruise speeds and altitudes.
The current altitude limit shown in figure 13 is the approximatelimit imposed on current positivedisplacementenginesdue to turbocharging capability.
Projectedimprovementsin turbochargingtechnologywill raise this limit. The high altitudemissions 7 and II were includedspecifically to investigatepos- sible shifts in rankingdue to (1) the criticalaltitude assumptionsfor the positivedisplacementengines and _2) the lapse rate and cabin pressurization losses of the turboprop.
RESULTS Figures14-16 presenttrend curves for the fixed-wingmissions and are used to illustratethe major results. More specific informationfor each mission is presentedalong with sensitivityinformationin appendixA. These results were obtained using the GeneralAviation SynthesisProgram (GASP)described in reference12.
The missions in these figures are grouped in order of increasingcruise alti- tude and, within each altitude group, in order of increasingcruise speed.
The projectedfuel savings for the 25,000 foot altitudemissionsare 25-35 percent for the advanced technologyversions and 35-50 percent for the very advanced versions (figure14). About 5 percent lower fuel savingsoccur for the lO,O00 foot altitude missions. This is a consequenceof selectinga naturally-aspirated SIR (with its 7 percent better BSFC) for the lower altitude baselinebut a turbochargedSIR for the higher altitudeswhile assumingthat all future engineswould require the turbocharger to meet engine weight and BSFC projectionseven at low altitudes.
Except for the turboprop,not much sensitivityis shown in the resultsto missiondefinition. This reflects the moderatedifferencesin BSFC and engine weight among engine types as displayedin figures 7 and 9. Due to its light- ness, the turbopropcompetes relativelywell at 25,000 feet, but for the lower and slower missions (l, 2, 3, 5), it does not, because its efficiencydrops for small sizes and power loading is low. Overall, the advanced positive- displacementenginescompete so closelywith each other that no clear-cut ranking is apparent. Even for the very advanced versions,which show more spread, the differencesdisplayedcould easily shift with differing (but equally plausible)assumptions. The very advanced rotary is shown to be the most fuel efficienttype due to its low weight, low BSFC, zero coolingdrag, and high cruise power rating (75 percent). However, it is also judged to entail the most technological risk and is probably further into the future than other types.
The relative shift in the diesel's positionbetween the 25,000 foot mission and the lower altitudemissions stems from a necessityto burn supplementary fuel (figure4) in the turbochargerloop above the 17,000 foot critical alti- tude to prevent a seriouspower lapse. Finally, it should be noted that the fuel savings are shown in terms of fuel weight rather than fuel cost. The actual cost savingsare about 15 percentgreaterdue to the fuel density and price advantageof Jet A and diesel fuel compared to aviation gasoline. The turbopropand diesel already utilizethe lower grade fuels while the rotary and SIR types currentlydo not (bothadvanced rotary versions plus the very advancedSIR would too, but not the advanced SIR).
Aircraft acquisitioncost reductionpotentialis shown in figure 15. Very little cost reductionpotentialis shown for the lO,O00 foot cruise altitude missions because the higher engine cost necessaryfor the turbocharged advanced engine (a 25 percent cost reductionwas assumedfor the non-turbocharged SIR base) offsets the structuralcost savings broughtabout by lower engine and fuel weights. The higher altitudemissions,with a turbocharged base, show a substantialacquisitioncost reductionfor the advanced technologyengines.
This indicatesthat less expensive,naturally-aspirated versions of these enginesmight offer substantialcost reductionpotentialfor the low altitude mission, provided large BSFC and engine weight penaltiesdo not occur.
The turbopropsuffersa downward shift in relative rankinggoing from the lO,O00 to the 25,000 foot missions. This is a result of the turboproplapse rate and pressurization requirementsnecessitating larger sea level horsepower.
This higher horsepowerrequirementincurs both engine cost and weight penal- ties to the airplane.
Five-yearTotal OwnershipCost (TOC)reductionpotentialis shown in figure 16 Ten to 35 percent improvements are shown for the lower altitudesand 20-45 per- cent gains at the high altitudes. For this criterion,the turboprop'sranking improvesconsiderablydue to its low maintenancecost. While the spread in ranking is fairly broad, with the rotary still on top and the SIR at the bottom, the overall interpretation is that all of these alternatives offer substantial improvementpotentialand, therefore,all should be retainedas competitive candidates.
The maximum rated horsepowerforecastedfor the lO,O00,16,000,and 25,000 foot cruise altitudemissions are shown in figure 17. The range covered by these missions was I00-500 horsepower / engine.For the positivedisplacementengine, the very advanced rotary requiredthe least horsepowerdue to its high cruise power rating (75 percent),zero coolingdrag, light weight, and high propeller efficiency. The turbopropwith zero cooling drag, lightestweight, and higher propellerefficiencywould be expected to have the lower horsepowerrequire- ments. At the lO,O00 foot cruise altitudemission, this is the case; however, at the higher altitudes,the turboproplapse rate resultsin high sea level Dower ratings. Also note that as the cruise speed is increasedat a given altitude, the turboproprelativehorsepoweris slightly reduced. This is due to forward velocityeffects as well as a secondaryeffect from favorableengine sizing effects. Missions 7 and II, which are not shown in this figure,have horse- power requirementsbetween450 and lO00 horsepower, with the turboproprequiring significantly more horsepowerthan its positivedisplacement counterparts.
The effect of the futuristicairframes,definedearlier, is to reduce the drag and power loadingof the airplane. For a constantmission this somewhat lessens the importanceof potentialpowerplantimprovements. In this study, however, the futuristicairframeswere combinedwith futuristicmissions (havinghigher cruise altitudesand speeds). The higher cruise altitudesand high wing loadings produce higher L / D, thus decreasingthe power loading,however,the higher cruise speeds increasethe power loading. Thus the combinationof futuristicaircraft and missions resulted in only minor differencesin power loadingand projected aircraft improvements.
. Figure 18 shows the percent improvementin the three figuresof merit for the two helicoptermissions. The resultsare similarto the fixed-wingmissions.
The turboshaftfuel reductionpotentialis considerably lower than the other engine types, but the acquisitionand maintenancecost advantagesare suffi- cient to make it competitiveto the other engine types on a vehicle acquisition cost or TOC basis.
HighAltitude Missions cruising altitudesas high as 30,000 to 45,000 feet have been suggestedfor some future general aviationairplanes. While this is a controversial issue, missions 7 and II were, nevertheless,includedto explore the ramifications of such extremealtitudeson powerplantselection. Current productionturbo- charged SIR engines are not normallycapableof flying these high altitude missions due to turbochargerlimitations. However, it is felt that an oppor- tunity may exist to substantiallyimproveturbocharger technologyand thereby permit operationat higher cruise altitudes. If so, then advancedpositive displacementengineswould have the advantageof avoiding the high power laspe suffered by turboprops,providedthe complexityand penaltiesassociatedwith high altitude turbochargers(yet to be determined)do not overshadowthe laspe rate advantage. As altitude increases,lower atmosphericpressurerequires higher turbochargerpressure ratios to achievethe same manifold air density, thus the same horsepower. The diesel and rotary,with their high pressure ratios required for turbosupercharging are more susceptibleto having insuf- ficientexhaust gas energy to power the turbocharger than the SIR. In fact, above 17,000 feet, the proposeddiesel must supply supplementary energy by burningfuel in the turbochargerloop combustor. Therefore,the very advanced SIR was chosen to comparewith the advanced turbopropto determinethe effect of high altitudeson powerplantselection.
Another issue is the increasingly more difficultprovisionfor cabin pressuri- zation at higher altitudes. Piston engine advocatesargue that this requirement penalizesthe turbopropmore than the piston engines since turbochargers can supply cabin air at negligiblepenalty. While this is true at low altitudes, it is not clear it will remain so at higher altitudesdue to insufficient exhaust energy to power the turbocharger. Also, cabin pressurization for turbopropsneed not take the form of compressordischargebleed; an auxiliarylow-pressureshaft- driven compressoror midstage bleed could be used to reduce the turboproppenalty markedly. Nevertheless,it is assumed herein that compressordischargebleed would be used to pressurizeturbopropaircraftand that no penalty for pressuri- zation would be incurredby the SIR.
I0 The resultsof varying the cruise altitudewith these assumptionsare displayed in figure 19. These resultsare for mission II, an 8-placeexecutivetwin with a nominal45,000 foot cruise altitude. Similarresultswere obtained for mission 7, a 6-placehigh performancesingle. The solid SIR curve represents a SIR engine with an ideal turbocharger--one capableof providing70 percent engine rated power regardlessof altitudeand without any weight or BSFC penal- ties. The projectedairplane performanceimprovescontinuously with higher cruise altitude,reflectingbetter aircraftL / D (wing loadingwas not optimized for maximum L / D as altitudevaried). The dashed SIR curve representsa SIR engine using a turbochargerwith a criticalaltitude of 21,000 feet. (The laspe rate above this critical altitudewas providedby the contractor.) This engine is able to provide 70 percent power to about 35,000 feet, and then decreases to 56 percent at 45,000 feet.
The turboproppoweredairplane performanceis not only affected by improving L / D, but also by the turboproppower lapse and cabin pressurization penalty.
The power lapse must be offset by an increase in engine size--theimproving BSFC associatedwith the larger engine is offset by the accompaningweight and cost penalties. As can be seen in the figure,the idealizedSIR is about lO percentbetter than the turbopropin terms of fuel and TOC and, about 25 percent better in terms of acquisitioncost at 45,000 feet. The 21,000 foot criticalaltitude SIR is comparableto the turbopropin terms of fuel and TOC but has about a 20 percent advantagein acquisitioncost. Up to about 35,000 feet, the turbopropis superior. At higher altitudes,the SIR may be better if very high criticalaltitudescan be successfully achievedwithout signifi- cant penalties.
SensitivityTo Engine Parameters The precedingcomparisonswere based on preliminaryengine performancecalcu- lationswhich incorporatedspeculativetechnologylevels, inconsistent assump- tions betweenmultiple contractors, etc. Hence, the specificcomparisons between engine types are subject to modification as more credibleengine data becomes availablein the future. An extensivebody of parametricvariations is presentedin appendixA so that the precedingresultsmay be correctedfor variationsin BSFC, engine weight, O.E._enginecost, engine maintenancecost, cooling drag and fuel cost. To illustratethe relative importanceof each engine assumption,graphicalexamples are given in figures20-22,which show the effect of varying the BSFC, engine weight,O.E.M. enginecostand engine maintenancecost on projectedTOC and mission fuel. These examplesare for mission lO, a 6-place business twin. Figure 20 shows the effect that varying the engine BSFC has on TOC and mission fuel. Similarly,figure 21 shows the effect of varying engine weight. Both BSFC and engine weight can be predicted with reasonableaccuracy. A lO percentengine weight variation,for any one engine type, causes its projectedimprovements to shift only slightlywith respect to the other engine types, while a lO percent BSFC change causes sig- nificant shifting in the relative engine ranking. The nominalvalues assumed previouslyare denoted with symbolson each line.
The effects of variationsin engine O.E.M. cost and engine maintenancecost on aircraft TOC is shown in figure 22. The maintenancecost shown here includesoverhaul cost reserves. A lO percent variationin engine O.E.M.
cost for one engine type does not cause significantshiftingin engine ranking while a lO percent variationon maintenancecost will producesome relative shifting betweenengine types. Engine O.E.M. and maintenancecosts for these advanced engines cannot be predictedwith any degree of accuracy at this time.
II Largevariations in theseparameters for any engineare possible and can cause significant shifting in relative engineranking.
Sensitivity to MissionParameters The effects of cruisespeed,cruiserangeand aircraft utilization on projected aircraft improvements for missionlO are shownin figures23-25. In these figures, the veryadvanced technology versions of eachpositive displacement engineare compared with eachotherand withthe advanced turboprop as well.
As cruisespeedis increased, all enginetypesshowincreasing relative improve- ment in all threefigures of merit(figure 23}. Thoughactualfuelburned, acquisition costand TOC worsenwith speedfor all advanced enginetypes,the currenttechnology SIR engine, whichtheyare compared against, worsensat a fasterrate. As the speedincreases the higheraircraft powerloading offers the lighter, more efficient engines morepotential for improvement. The turbo- propwiththe lowestspecific weight,favorable performance scalingtrends, and forwardvelocity effects has a significantly largerrateof improvement thanthe otherenginetypes. In thisparticular case,the turboprop-powered aircraft wouldappearat leastcomparable to all the veryadvanced positive displacement enginesin all threefiguresof meritif the cruisespeedwereraisedabove350 knots.
The effect of cruise range on the three figuresof merit is shown in figure 24. All engine types produced slightlygreater improvements as range increased, however, no significantshift in engine rankingoccurred over a wide band of cruise range. Hence, for the short missions flown by generalaviation,range is not an improtantparameterin choosing an alternativepowerplant.
The effect of aircraft utilizationon TOC is shown in figure 25. Some shifting in relative ranking occurs below 500 hours / yearutilizationdue to the effect of acquisitioncost. Above 500 hours / yearoperatingcost becomesmore important and the utilizationhas little effect on the choice of an alternativepowerplant.
CONCLUDING REMARKS This study has compared four proposedalternativeengine types for future gen- eral aviation. Each of these enginesoffers substantialaircraft efficiency and economic improvementsin terms of aircraft total ownershipcost, acquisition cost and mission fuel. The rotary appears very promisingdue to its very low weight, BSFC and coolingdrag. However, its advantagescould easily be lost if it cannot achieve these goals--especially since differencesin engine cost and technologyrisk were not taken into account. The diesel offers large advan- tages due to its good BSFC and relativelylight weight, but the presentdesign suffers large BSFC penaltiesat high altitudes. The turboprop,despite poor BSFC,iscompetitivedue to its low weight, zero cooling drag, and low maintenance cost, especiallyfor missions with high cruise speeds, but this competitiveness depends on meeting the forecastedlarge engine cost reductions. The SIR has similar performanceto the diesel, but loses economic competitiveness due to its large maintenancecost. Maintenancecost forecastingis difficult,and projectionsmay differ greatly from the actual cost. The presentedsensitivity data permitscomparisonof engine types using differentassumptionsfrom the nominalones used in thisreportand also permits evaluationof the benefitof incrementalchanges in engine parameters.
L To compare these alternativepowerplantsglobally involvesconsiderations of many more characteristics than the three figuresof merit presentedherein.
Engine vibration,emission,noise, reliabilityand other factorsmust all be taken into account. With the relativeclosenessof the projectedimprove- ments and the uncertaintyof engine assumptions,it is prematureto draw firm conclusionsregardingthe relativeattractiveness of the alternativeengines.
APPENDIX A ENGINE PARAMETER SENSITIVITIES One of the primary purposes of this study was to present sensitivity infor- mation to adjust the foregoing results for changes in the assumed engine characteristics. Sensitivity coefficients for total ownership cost, acqui- sition cost, and mission fuel with changes in cruise BSFC, installed engine specific weight, specific O.E.M. engine cost, fuel cost, and cooling drag are given in the bottom halves ofTables AI-AI3. The base values of the three figures of merit as well as the assumed engine parameters are given in the top halves of the Tables.
To illustrate the use of this sensitivity data and as a check on its accuracy, consider the problem of estimating potential aircraft improvements due to very advanced technology given the improvements for advanced technology. To do this for mission I0, we note from Table AIO that the baseline improvements for the advanced SIR, for example, are: mission fuel, Wf - 30.6% A / C acquisition cost, AC - 7.8% total ownership cost, TOC - 19.8% These improvements stem from the following engine assumptions (Table 7): specific fuel consumption, SFC - 0.361 Ib / hp-hr specific fuel weight, We - 1.42 Ib / hp specific engine cost, Ce - 40.0 $ / hp (OEMI engine maintenance cost, MC - 16.57 $ / fl-hr engine cooling drag, CD - 5.0 % fuel cost, Cf - I.I $ / gal For the very advanced SIR, the engine assumptions are (Table 7): SFC - .334 Ib / hp-hr We - 1.20 Ib / hp Ce - 40.0 $ / hp (OEM) MC - 15.56 $ / fl-hr CD - 3.5 % Cf - 1.0 $ / gal Using these values plus the sensitivitydata in the lower portionof Table AIO yields the estimatedaircraft improvements for a very advanced SIR: %AWf = %AWf+ _Wf (SFC- SFC)+ BWf (_ee - We) + _Wf (_-_ _ CD) @SFC _ BCD = 30.6 + 214 (.361 - .334)+ 6.5 (I.42 - 1.20) + 0.73 (5.0 - 3.5) = 38.9 (Actual valueis 39.2,fromTableAl4) @AC IS-#-CSFC)+ @AC_ @AC , 7- BAC (_-_ CD) %AAC = %AAC + _ - BTe e'We) +BTe_ ' e-Ce) +_ - = 7.8+29 (.361 - .334) + 4.6 (1.42- 1.20) + 0.39 140.0-40.0) + 0.25(5.0-3.5) = lO.O (Actualvalue is I0.8) •@TOCI_- SFC)+BTOC -- .@TOC,-_- Ce)+ @TOC(_-_MC) %ATOC= %ATOC*_ @-Te (We-We).B-_-e_e - @MC - + BTOC (__Cf). BTOC BCf . _ (_- CD) = 19.8+76 (.361 - .334)+ 5.210.42- 1.20)+ 0.33 (40.0- 40.0)+ 0.82 06.57 - 15.56) + I07 (.183-.149) + 0.48(5.0-3.5) = 28.2 (Actualvalue is 26.1) In the above, bars over values denote baselinevalues and the cost of fuel is convertedto a $ / Ib basis by dividingby the fuel density (i.e.,for AV gas, Cf = l.lO / 6.0 = 0.183 $ / Ib and for Jet A, Cf = 1.0 / 6.7 - 0.149 $ / Ib).
Results for the SIR, diesel, and rotary are displayedin Table Al4 and indicate good agreementbetweenestimatedand actual values.
REFERENCES (I) Brouwer, A. P.: 150 and 300 kW Lightweight Diesel Aircraft Engine Design Study. NASACR-3260, 1980.
(2) Jones, C.; and Berkowitz, M.: Multifuel Rotary Aircraft Engine.
AIAA Paper 80-1237, June 1980.
(3) Strack, W. C.: New Opportunities for Future Small Civil Turbine Engines: Overviewing the GATEStudies. NASATM 79073, 1979.
(4) Baerst, C. F.; and Furst, D. G.: General Aviation Turbine Engine (GATE) Study. (AiResearch 21-2997, AiResearch Manufacturing Companyof Arizona; NASAContract NAS3-20755.) NASACR-159482, 1979.
(5) Smith, R.; and Benstein, E. H.: Advanced General Aviation Turbine Engine (GATE) Study. (TELEDYNE-CAE-1600,Teledyne CAE; NASA Contract NAS3-20757.) NASACR-159624, 1979.
(6) Lays, E. J.; and Murray, D. L.: Advanced General Aviation Turbine Engine (GATE) Concepts. (WRC-78-113-15, Williams Research Corp.; NASAContract NAS3-20758.) NASACR-159603, 1979.
(7) Gill, J.; et al.: Study of an Advanced General Aviation Turbine Engine (GATE). (DDA-EDR-9528, Detroit Diesel Allison; NASA Contract NAS3-20756.) NASACR-159558, 1979.
(8) Stuckas, K. J.: Advanced Spark-lgnition Aircraft Piston Engine Study. (Teledyne Continental Motors, NASAContract NAS3-21272.)
NASACR-165162, 1980.
(9) Winblade, R. L.; and Westfall, J. A.: NASAGeneral Aviation Research Overview-1976. SAE Paper 760458, Apr. 1976.
(I0) Kempke, E. E.; and Willis, E. A.: Overview of NASAResearch on Positive Displacement Type General Aviation Engines. NASA TM-79254, 1979.
(II) Harned, M. S.: General Aviation Aircraft in the '90s. Astronaut.
Aeronaut., vol. 18, no. 1, Jan. 1980, pp. 46-53.
(12) GASP- General Aviation Synthesis Program. (Aerophysics Research Corp.; NASAContract NAS2-9352) NASACR-152303, Vol. 1-Main, Vol. 2-Geometry, Vol. 3-Aerodynamics, Vol. 4-Propulsion, Vol. 5-Economics, Vol. 6-Mission Analysis, Vol. 7-Weight and Balance, 1978.
T , LE _I . - TOTAL O _ HERSH I P COS T (TOC)ASSUHPTIONS 5 YEAR P ERIOD OFO W NERSHIP TOC- ACQUISITION COST . OPERAT _ G COST . IHTEREST-RESALE VALUE [ i MHERE : Acquisi t ion Cost Is the tot a l retail price of the a ircr a ft Operattn9 Cost ts based on 500 hr / y r utilization a nd tncludes fuel, o11 , Inspection a nd maint e nance, ov e rha u l reserve, Insur a nce , sto r a ge a nd FAAtax Interest ts based on a 5 y ear loan with a 2 0 _ downpa y m ent a nd a 10 _ tnterest rate Resale Valu e t$ a ssumedto be 70 _ of a cquisition cost TA B LE2 . - A CCESSORY PO W ER R E Q U I RE M E N TS ENGI NE A LTER NA TO R FU E L P UM P O I L PU M P A / C & PRE SS U RIZA T I O N TYP E R E QUIREMENTS REQUIREMENTS REQUIREMENTS REQ U IREM E NTS SIR 3. 7 HP 1. 4 HP 0.5 H P 0 Rot ar y 3 . 7 HP 1 .4 HP O . S HP 0 Dies e l 5 . 0 HP 3 . 5 HP 0 . 5 HP l O HP Tur bop r op 5 . 0 HP f o r Sl ngl e 8 . 0 HP f o r Twi n 1 . 0 - 1. 5 Ib l m l n l pa ss Co m p re sso r D i sc har g e Air N o te : -Re qui reme n t s a re quo te d f o r a co mp l etea l r c r a ft a n d a re b as e don a 3 5 0 h o r s e po wer d e slgn . Fo r a 11 e n g ln e t ype s e xc eptthe t u rbo p r o p, t he s e re qu ireme n t s a re 1 | n earl y sc a l e d wit h e ng ine siz e.
T he tu rb op r o p re qu irement s are a ss ume d f | xe d .
TA B LE3 . - E N G IN E AS SU M PT I O N S ENG IN ET Y PE( a ) N O MINAL CRU IS EPO W ER OVE R HA U L CO OL INGDRA G FU EL , P R OPELLE R CRI T ICAL S E I'FIN G C O S T % A / C DRA G T Y PE I COS T EFF ICI E N C Y A LT ITU DE % PO W ER / R I T(°F ) ( b ) ( AI RP LAN E S O N L Y ) ( $ 1 GAL) FT S.I. Reclp Cu rrent(Ba s e l ine) - N at. A spi rat ed 0 7 0 0. 4 3 10 A v Gas / l .lO 0. 8 7
Tu rb och arg ed 1 8 D O 0 i / lO Av G as / l . lO
A dvanced 21 DO0 _i _ 5. 0 A vGas / l.]O VeryA dv a nc e d 2 1 00 0 _ 3. 5 Je t A I I.O0 JL Die s el A dv a nced 17 000 70 0 .43 5 . 0 Oe t A / 1. O0 0 . 8 7 V e ry Adv a nced 1 7 0 0 0 70 0.43 3.5 O e t A / 1. O0 0 .87 Ro t a ry Adv a nced 20 0 00 75 0.43 3.5 O e t A / 1 .O0 0.88 Ve ry Adv a nced 2 0 000 75 0 . 43 0 J e t A l l .DO 0,88 Tu r b i ne En gi nes Adv a nced 21 40(c) 0.376 0 J e t A / I .O0 0.89 (a ) Al l a dvancedp ositive d i spl a c eme n t eng i nes are turb oc har g e d .
(b) F ra c tl o n o f i n itialll s t pri c e (O. E .M. c o s t. 0. 6 ).
( C )Typl c al GAT E r o t o r inlet temperat u re.
T AB LE 4. - MISSION A S S UMP T I O NS T.O.** P ayl o a d Altit ud e, S p e e d , Ra n ge, Di s tance, W I S ,_ Mi ssion Airplane ( I ncl . P i lo t ),lb A R Ft K t s N. Miles ft I b / ft = I S i ng l e2- P L Trainer 455 8 lO000 1 1 0 500 11 00 10 2 Si n g l e4-PL Uti l lty 800 8 1 0000 130 600 16 00 20 m 3 Sing l e4- P L Utility* 800 1 2 1 6000 1 60 800 1 600 35 4 Sing l e4-PL U ti l lty* 800 1 2 25000 2 1 0 1 400 1 600 40 5 Si n gle 6 - PL U tlllty 12 00 8 1 0000 180 600 1 8 0 0 25 6 S i n gle 6 -P L Hl -Perf 1 2 00 8 2 5 00 0 25 0 10 0 0 220 0 35 l S ln g l e 6 -PL Hi-Perf* 12 00 1 2 40000 3 40 16 00 2400 45 8 Twin 4- P L L i ght * BOO 1 0 25000 3 00 1 400 16 00 45 g Twi n 6- P L Me d ium 1 200 8 1 0000 230 1 1 00 17 00 30 1 0 T wi n 6-PL Business * 1200 10 250 0 0 270 1 600 2 2 00 50 m, 1 1 T wi n 8-PL Ex e c utive* 16 00 1 0 4 5 000 38 0 1700 2 500 6 0 12 S lngle 4- P L He licopter 800 -- 2000 110 3 00 6 000*** -- 13 Tw l n 6- P L Hel l c op ter 1200 -- 2000 13 0 500 1 0000*** -- * F uturi s tic Ai r craft& Missi o n s **Overa 3 5 F oot Obstac l e E mptyWeightRe d uce d by 15% F us e l age P re ss ureMaintaine d at 80 0 0 Ft . Air.
Zer o Lift Drag Re d uce d by 1 5% E n g i ne s L o cate d I ns i d eFu s e l age * ** Hover C ei l lngO u t o f G ro u n dE ffect F u ll Span F o wlerF l ap s T ABL E Ai. - S E NSITIVI T Y F O R MISSION1,A 2- P LAC E TRAIN E R W L = 4 55 Ib H • 1 00 0 Oft V = 110k t s R = 500N.M. W / S = I0 Ib / ft 2 { 1 980 Recip TOC* =$50000 AC* =$24000 Wf* = 1901b SHP*= 13 7 "I BENEFITSREL.TO 1980 RECIP. BASELINE ASSUMPTIONS _ATOC , _Z_AC _z_Wf SHP SFC We C_ MC Adv. S.I. Recip 8.9 - 5. 3 22.6 124 .365 I.B6 46.45 3 .62 VeryAdv. S. I. Recip 18.2 0.8 32.6 I1 7 . 3 3 9 1.5 7 46.62 3 .40 Adv. Rotary 18. 3 4.4 24.7 107 .386 1.50 46.57 3. 0 7 VeryA d v. Rotary 33.1 13.2 4 2.6 96 .336 1.00 46. 7 0 1. 9 5 Adv. Die s el 1 7 .6 - 2.2 24. 7 121 .365 1. 6 2 47.2 6 2. 77 VeryA d v. Diesel 26.0 2.6 3 7 .9 I15 .326 1.45 47.60 2. 3 1 Ad v . Tu rb o pr op ( SHP t o s ma ll t o m o d el) ........
SE N SIT I VITY COEFFIC I E N TS (% _ PE RA ) _TOC- _TOC _TOC. _TOC _TOC _TOC _AC _AC _AC _AC _Wf __ aWe bCe 6MC C_ &CO 6SFC _ee _ aCO _SFC S.I. 98. 8 .4 0.47 5.0 148. 0.5 3 41. II.3 0.90 0.60 236. 7 .1 0. 7 7 R o tary 9 7 . 6.4 0. 33 4.9 129. 0.5 7 3 3 . 9 .5 0.81 0.5 9 195. 5.0 0. 7 5 Di es el 8 9 . 7 .I 0 . 35 4 . 9 1 37 . 0 .60 37 . 1 0 . 8 0 .B 7 0 .6 0 • Z14. 4.2 0 _ 70 Turb o prop ..........................
TOC - T o tal Owne r ship Cost, S SFC - Spec i f ic Fu e l C o n sumpti on, Ib / HP-Hr AC - Acqui s ition Cost, $ We - EngineWeight,Ib / HP " Wf - Weight o f F u el, I b Ce - E ngineCo s t, $ / H P Cf - Fuel Cost, $ / Ib MC - E ngineMaintenance Cost, S / FlightHr CD - CoolingDrag; %A / C Drag T A BLE A 2 o - S E N SITI VITY F O R MISSION2, A 4-P LA CE UTILITY W L = 800 l b H • 100 00 ft V = 13 0 kts R • 60 0 N.M. W / S = 20 l b / f t 2 Recip Wf* TOC*= $76000 AC* • $41000 - 270 lb SHP* = 200 BEN E FITSR E L. TO 1980 RECIP. BASELINE ASSUMPTIONS I%Z_T O C _Z_AC %AWf Sli P SFC W e Ce MC A d v. S.I. Recip 6. 9 -5.1 20.4 185 .363 1.5 9 4 3 .2 3 5.44 V ery A dv . S . I . R ec i p 14 .2 - 1 . 6 28 . 9 178 . 337 1 . 35 43 . 5 2 5 .2 3 A d v. R o tary 14. 3 0.4 20.0 1 64 . 3 84 1 .3 7 44. 3 6 4.54 Ver y A d v. R o tary 2 6. 3 5. 9 3 5. 6 1 52 . 33 4 0.914 44.88 3 .01 Adv. Die s el 14 . 7 - 3 .9 2 1 .1 18 3 . 3 6 3 1.45 4 3 . 3 7 4.05 V ery A d v . D ie s el 2 1 . 1 - 1 .1 3 1 . 9 1 77 . 3 2 4 1.31 43 . 6 0 3 . 46 A dv . Tu rb o pr o p 23 . 2 3. 1 7.4 140 . 525 .798 45 . 53 1 . 92 S E N S I T I V ITY COEFF I C I ENTS ( % A PERA) r_ T OC _TOC a TOC _TOC _TOC _TOC _AC _AC _AC 6AC _W£ _ 6co _W e 6W--e--_Ce _MC _ _CD _ _e _Ce aCD &SFC S.I. 82. 5.8 0.4 5- 3.3 144. 0.60 23. 7.7 0.78 0.43 233. 5.8 0.87 Ro tar y 7 3. 4.7 0.33 3.3 132. 0.48 20. 7.1 0.72 0.40 20 5 . 4. 7 0.68 Die se l 77 . 5 .0 0. 34 3 . 3 137. 0 .5 0 31. 7 .1 0 . 7 7 0.4 0 21 8 . 5 . 5 0 . 77 Tu r h o o ro o 66. 5.1 0.29 3 . 3 1 72 . -- 2 0 . 7 . 0 0 . 61 -- 206. 7 . 6 -- TO C - To t a l _ vne rs h i p Co st , $ S F C - S pec ifi c F ue l Consump ti on , l b / H P- H r AC - Acquisition Cost, $ We - EngineWeight,Ib / HP " Wf - Weight of Fuel, Ib Ce - EngineCost, $ / HP Cf - Fuel Cost, $ / Ib MC - EngineMaintenance Cost, S / F lightHr CD - CoolingDrag;%A / C Drag T AB L E A 3 . - S E NSITIVI T Y FOR MISSION 3, A 4-P L ACEU T ILITY W L = 800 l b H • 1 60 0 0 ft V = 1 6 0 k ts R = 800 N.M. W / S - 35 I b l ft 2 1980 Reclp TOC* = $85000 AC* = $44000 Wf* = 323Ib SHP* = 211
1 I "
BENE F ITSREL. TO 1980 RECIP. BASELI N EASSUMPTIONS . _ATOC _ /k AC _AWf SHP SFC W_ Ce . MC A d v. S.I. Recip 1 8.4 5.3 26.2 1 9 3 .3 63 1.58 4 2 .92 5 .6 9 Ver y A dv . S. I . R ec l p 2 4. 3 7 . 6 33 . 4 187 . 336 1 . 33 4 3 . 13 5. 50 A d v. R o tary 2 4.5 9. 2 25.6 1 73 .384 1 .36 43.89 4. 75 Very Adv. R o tary 3 4. 6 12. 9 39 .0 162 I . 33 4 .901 44. 32 3 .1 9 Adv. Die s el 2 5. 3 5.9 2 6. 9 192 .36 3 1.4 4 4 2 . 9 4 4. 23 V ery Adv . Dle s el 3 0.8 7.9 36 . 2 1 86 .3 2 3 1.3 1 43,24 3 . 63 Adv. Turb op r op 27 . 7 4. 9 6 . 8 171 . 55 4 . 723 4 3 . 89 2 . 33 SEN S I T I V I TY COEFF I C I E N T S ( % A PE R A)
_TOC _T O C _ T OC 6 TOC _T O C _ _AC _AC _AC 6AC _ __
c _
s_ _ _ _MC _ _CD _ "_ee_e 6CD _SFC
S.I. 7 4. 3.1 0.45 2.9 132. 0.51 2 i . 3 .8 0. 7 8 0 .39 214. 3 .8 0 .6 7 Rotary 6 7 . 2.2 0. 33 2.9 123. 0.42 18. 3 . 3 0. 7 2 0.3 7 194. 3. 7 0.64 Die s el 6 9 . 2.6 0.34 3 .0 12 7 . 0.39 18. 3 . 7 0 . 7 7 0.36 20 3 . 3 . 7 q. 67 Turbopr o p 7 1 . 2.4 0.2 9 2. 9 169. -- 31. 3 .4 0.6 1 -- 214. 2. 9 -- TOC -Tota ] OwnershipCost,$ SFC - SpecificFue l Con s umpti o n, I b / HP - Hr AC - Acquisition Cost, $ We - EngineWeight,Ib / HP , Wf - Weightof Fuel, lb Ce - Engi n eC o st, $ / NP Cf - Fuel Cost, $ / Ib MC - EngineMaintenance Co s t, S / FllghtHr CD - C oo lingDrag; %A / C Drag i T ABL E A 4. - SENS I T I V I TY FOR MISS I O N4 , 4- PL AC EUTI L I T Y W = 8 00 lb H - 2 5 0 00 ft V • 2 10 kts R = 34 00 N .H . W / S • 40 l b /ft 2 L
J1980 Recip TOC*- $137000 AC*- $70000 gf*- 723 lb SHP*- 365I
BENEFITSREL. TO 1980 RECIP. BASELIN E ASSUMPTIONS _ o Z_TOC _Z_AC _AWf SHP SFC W e Ce MC A dv . S.I. Re c lp 22.3 1 0. 3 2 8 . 4 325 .359 1.39 40.00 9.88 Very Adv . S . I. Re c i p 2 9 . 3 14 . 1 36 . 4 310 . 333 1 .17 4 0. 0 0 9 . 38 A d v. Rotar y 3 0.4 1 5.5 2 7 .9 • 28 9 . 3 80 1 . 18 4 0.00 7 . 5 4 VeryA d v. Rotary 41.4 21. 3 42.0 266 . 331 . l g 7 40.00 5 .0 3 Adv. Die s e l 3 0.2 11 .5 2 7 . 1 3 22 . 37 0 1 .22 40.00 6 .85 V e ry Adv . Die s el 3 6. 3 14 . 8 37 . 1 3 09 . 331 1. 1 1 40 . 0 0 5 . 84 A dv. Tur b op r op 34.3 5 .8 25 . 9 339 .4 56 .590 40.0 0 4.71 SENSI T I V I T Y C O EFFICIE NTS ( % A PE R A ) _TOC _TOC _TOC 6TOC _TOC _TOC _AC _AC _AC 6AC _Wf 6Wf _Wf &W_ aCe 6MC &Cf &C D _ _ee C_ee _CD _SFC S.I. Re c ip 9 1. 4.8 0.45 1.8 1 3 6. 0.6 7 44. 6.5 0.8 3 0.58 212. 4. 3 0. 73 Rotary 7 6. 2. 7 0. 31 1.8 12 3 . 0. 4 7 3 5. 4 . 7 0. 7 5 0.42 191. 4. 1 0.7 3 Die s el 85. 3 .8 0.34 1.8 129. 0.55 42. 5.4 0. 8 0 0.52 20 8 . 3 . 8 0 . 73 Turboprop 74, 6,1 0.39 1.8 14 3 . -- 44. 1 1 . 3 0.8 9 -- 192. 6.1 -- TOC - Total OwnershipCost, $ SFC - Specific Fuel Consumption, Ib / HP-Hr AC - Acquisition Cost, $ We - EngineWeight,Ib / HP Wf - Weightof Fuel, lb Ce - EngineCost, $ / HP Cf - Fuel Cost, S / Ib MC - EngineMaintenance C ost , S / FlightHr CD - CoolingDrag;%A / C Drag T AB LE AS. - S E NSITIVITY F OR MISSIONS , A 6- PL AC E U T I L I T Y W L = 1 200 I b H = 1 0000 ft V = 1 8 0 k t s R = 6 00 N.M. W / S- 2 5 lb l ft 2 1 1980 Recip TOC* - $136000 AC* = $84000 Wf* = 3391b SHP*- 342 1 BE N EFITSREL.TO 1980 RECIP. BAS E LI NE ASSUMPTIO N S _Z_TOC _AAC _ L _Wf SHP SFC We C € MC A d v. S.I. Recip 8.2 -2.4 21.2 3 12 .360 1.40 40.00 9 .44 Very Adv. S. I. Recip 16.0 1 .6 30.7 2 9 7 .333 1. 1 8 40.00 8.9 3 A d v. Rotary 1 7 . 4 3 .6 2 2.1 274 . 3 8 1 1 . 20 40 . 00 7 .1 5 Ver y Adv. Rotary 2 9 .0 9. 3 38.1 250 . 33 1 . 80 8 4 0.00 4 . 73 Adv. Die s el 1 7 .4 -0.2 2 3 .0 3 06 .35 9 1.25 40.00 6. 53 V e ryA d v. Di e s e l 2 3 . 6 2 . 5 33 . 9 294 . 3 2 0 1 . 13 40 . 00 5 . 5 6 A dv . Turboprop 2 6 .2 7.4 12.7 231 .4 9 6 .608 40.99 3.11 SENS I T IVI T Y COEFF I C I ENTS ( % A PER A ) _TOC 6TOC 6TOC 6T O C _TOC T_ _AC _AC _AC 6AC _Wf ___ _Wf _CD S_ &We aCe _MC C_ _C0 _ W_ee _e 6CD _SFC 6VI e S.I. Re clp 80. 7 .8 0.42 1.8 13 7 . 0.6 7 21. 9.6 0.6 7 0.4 7 224. 7 .8 0 .80 .
Rotary 7 0. 6.1 0.31 1.8 124. 0.4 7 .1 7 . 7 .5 0.60 0. 3 1 200. 5. 4 D.BO Die s el 7 5. 6.4 0. 3 2 1.8 131. 0. 47 18. 8.8 0.65 0. 33 214. 7 . 0 0 .80 Turboprop 63. 5.6 0.26 1.8 155. -, 36. 8.0 0.52 -- 205. 7 ,2 -- Toc - Total OwnershipCost, $ SF C - SpecificFuel Consumption, Ib_HP-Hr AC - Acquisition Cost, $ We - EngineWeight,Ib / HP Wf - Weightof Fuel, Ib Ce - EngineCost, $ / HP Cf - Fuel Cost, $ / ]b MC - E ngineMaintenance Cost, S / FllghtHr CD - CoolingDrag; %A / C Drag • J
T_LE A6 SE N SITIVITY F O R MISSI O N 6, A6P_CE .I-PERFOR_ N CE
,_. ,_O O lb .-_oo o, _ v.2_ok_sR-I oo o N ._. w/S-351b/ f t2
!19_ 0_eotp _0:*_ S_4_ODD AC*- S15_OOO Wf '_ 73, lb S N , -5_91
BENEFITSREL. TO 1980 RECIP. BASELINEASSUMPTIONS • %ATOC % /\ AC %_Wf SH P SFC We C_ MC Adv. S.I. Recip 24.3 12.8 32.3 493 .354 1.39 40.00 15.90 V e r y A d v . S . I . R eci p 3 2.2 18 . 2 41 . 5 4 6 2 . 3 28 1 .16 40 . 00 14 . 7 6 Adv. Rotary 3 5.3 21.2 35.0 420 . 37 6 1.0 3 40.0 0 1 0 . 9 6 Ver y A d v. Rotary 4 5. 7 2 7 .5 49.1 3 7 8 . 3 2 8 .926 40.00 7 .15 Adv. Die s el 3 4. 9 18.2 34. 3 468 .366 1.01 40.00 9 . 97 VeryAdv. Diesel 40.5 21. 3 44.0 446 .32 7 . 7 04 40.00 8.44 A d v. T u rbo p ro p 38 .6 15 .6 33 .4 4 9 5 . 5 26 . 5 4 9 40.0 0 7 .20 S E N S IT IVIT Y C O EFFIC I E N TS ( % a PE R A ) _TOC 6TOC 6TOC 6TOC _TOC T_ _AC _AC _AC 6AC _Wf _Wf 6Wf s_ aw e _C e _MC _ _C 0 S_ _ee C_ ee a C D _SF C F " .... " .....
S.I.Recip 8 3 . I 0 .0 0. 3 6 1.0 I16. 0.6 7 38 . 12.6 0.59 0.5 3 20 3 . 8. 7 0.80 R ot ary 6 7 . 6 . 8 0 .2 5 1. 0 1 01 . 0 . 45 2 9 . 9 . 5 0 . 5 1 0.41 17 3 . 6 .1 0 . 65 Die s el 7 4. 7 .6 0. 26 1.0 108. 0 .4 7 3 4. 10.8 0.5 6 0.4 7 1 9 1 7 .0 0, 67 Turb o pr o p 66. 1 0 .5 0.3 1 1.0 1 17 . -- 3 4. 1 6.7 0.6 1 -- 1 87 . 1 1 .4 -- TOC - Total OwnershipCost, $ SFC - SpecificFuel Con s umpti o n, Ib]HP-Hr AC - Acquisition Cost, $ We - Eng i neWeigh t ,lb / HP Wf - Weightof Fuel, Ib Ce - EngineC o st, $ / HP Cf - Fuel Co s t, $ / Ib MC - EngineMaintenan c e C ost, S / F llgh t Hr CD - CoolingDrag;%A / C Drag T ABLE A 7. - S E NSI T IVITY FOR MISSION7*, A 6 PLA C E HI-PERF W L = 1 40 0 lb H = 40000 ft V " 3 4 0 kt s R = 1600 N . M . W / S " 4 5 l b / ft 2 ( ADVANC ED TURBOPROP TOC*= $188000 AC*= $168000 Wf*= 572 Ib SHP* " 827( BENEFITS REL. TO ADV. TP BAS E LIN E ASSUMP T IONS % A TOC % AA C % AWf S H P SFC W e C e MC Ve ry A d va ncedS. I . R ectp 6 . 9 2 1.0 8 . 6 4 86 0. 3 2 7 1 . 16 4 0 ' 00 15 . 6 6 W t t h I deal i zed Tu rbo cha r ge r Ve ry Adv a nc e d S. l . R ec l p 1 . 2 17 _ 3 6 . 8 552 0 , 327 1 . 16 40 . 0 0 17 . 35 W lt h 2 1 000 F t C r itic a l T u rbo c ha r ger A d van c e d Tu rbopr op 0 0 0 8 2 7 0. 3 8 6 0.4 6 4 40.00 12. g l S E NSI T IVITY C OEFF I C I E N T S ( % A PER A ) 6TOC 6TOC 6TOC 6TOC 6TOC 6 T OC 6AC 6AC 6AC 6AC 6Wf 6Wf 6 Wf 6S F C 6We 6 C e 6 MC 6 O f 6C-'- D " 6 S FC _ _ 6 C --- D 6 SF C 6W'-- e 6C--' 0 S .I. Recfp** 102 1 0. 4 O . S O 1. 3 1 57 O. B2 38 11 . 3 0 . 53 0 . 64 341 1 4 . 4 1.12 Tu rbo p r op 95 21 . 5 0.69 i . 3 154 .... 45 24.1 0.94 .... 3 1 1 18 . 0 ....
TOC - To ta l O w ne r shipCos t, $ S F C . Speci f ic Fu el Co n su mp t i on , l b / HP - N r A C - A cqu i s iti on C os t, $ W , - En g ln e ' W el ght, Ib / MP Wf - Weight o f Fuel,Ib C _ - E ngine Cos t, $ / H P Cf - F u e l Cos t,$ / Ib MC- - E ngineMainten a n c e C os t,$ / F lf_htHr C O - Cool i n g D rag;% A I C D r a g *Adva n ced Tu rboprop B as eline E ngi n e.
* * Ba s e don VeryA d v a n ced S.I. R .wlth I d eali z e d Turb oc h a rger.
T AB LEA8 . - S E NSI T IVITY F OR MISSION8, A 4-P L AC EL IGH T TWIN W L = 800 I b H = 2 50 0 0 ft V = 3 00 kts R = 1400 N.M. W / S = 45 Ib l ft 2 I 1980 Recip TOC * = $264000 AC* = $179000 Wf*- 896 Ib SHP* = 3081 BENEFITSREL. TO 1980 RECIP. BASELINEASSUMPTIONS % _TOC _AC % z _Wf SHP SFC g e C¢ MC Adv. S.I. Re c ip 22.8 I0.5 3 1.6 265 . 3 61 1.44 40.00 15 .68 Very A d v. S. I. Recip 29.9 1 4.4 40.0 250 . 3 35 1. 2 1 40.00 1 4.68 A d v. R o tar y 2 9.9 14. 7 3 2.3 233 . 3 8 1 1 . 2 6 40.85 1 2. 30 Ver y Adv. Rotary 40.6 20.3 4 7 .1 209 . 333 .84 9 40.00 8.0 6 Adv. Die s el 28.9 I I.2 2 9. 9 2 6 3 . 372 1.32 40.00 I I.20 VeryA d v. Diese l 3 5. 1 1 4.6 40.5 249 . 33 3 1 .20 40.00 9 . 4 2 Adv. Turboprop 3 3 .2 8.8 26. 7 261 .459 .60 1 40.00 7 .05 S E NSITIVITY COE FFICI E NTS (% A PE R A) _TOC _TOC _TOC _TOC _TOC _TOC _AC _AC _AC _AC _W £ .._ _Wf _Wf 6 C D _We _ _Ce _MC _ _CD _ W_ee _Ce 6CD _S F C S,I, R e c i p 8 3 . 7 .0 0 . 3 8 0 . 93 117 . 0 .54 3 5. 7 .8 0 .5 3 0 . 3 4 2 03 . 5 . 7 0 . 74 Rotary 7 0. 5.2 0.25 0. 97 1 03. 0.40 30. 6.5 0. 47 0. 31 1 82. 5 .2 0. 7 1 Die s e l 7 5. 5. 7 0.2 7 0.96 1 09. 0.42 35. 8.0 0.5 1 0. 3 4 20 0 . 7 . 4 0. 71 Turboprop 7 4. 7 . 6 0.32 0.9 7 129. -- 52. 1 0. 7 0.56 -- 20 7 . 8.4 - r TOC - Tota l OwnershipCost, $ SFC - SpecificFuel Consumption, Ib / H P -Hr AC - Acquisition Cost, $ We - EngineWeight, Ib / HP Wf - Weightof Fue l , Ib Ce - E ngineCost, $ / HP Cf - Fuel C ost, $ / Ib MC - E ngineMa i ntenance Co st , S / F l igh t Hr CD - Coo l ingDrag; %A / C Drag T _ LE Ag . - S E NS IT I VIT Y FOR MISS I ON 9 , A 6 -P L A C E M ED I U Ml_ I N
WL'I_O O Ib N-IOO OO ft V -23 O kt s Ro11 00 N ... WIS-3 O Ib . t2
1198 0 Reclp TOC*- S37SOO0 AC*= S253000 Wf*- 1324 Ib SliP* - 493 ]
BE N EFITSREL.TO 19BO RECIP. BASELINEASSUMPTIO N S %ATOC _Z_AC _Wf SHP SFC W e C¢ MC Adv . S . I . Re c ip 9 . 7 - 1.1 23 .6 4 5 0 . 35 6 1. 38 40. 0 0 28 . 6 4 Very A d v. S. I . Re c lp 1 9 .3 4 . 7 34.7 4 19 .32 9 1.16 40.0 0 2 6 .4 5 A d v . Ro tary 2 1 . 9 6. 8 26. 5 3_ . 377 1 . 07 40 . 00 2 0 .0 4 Very A d v.R o tary 3 4. 6 1 3 . 8 4 3 . 6 3 45 .32 9 . 7 26 40.00 13 .05 A d v. Die s e l 23 . 9 5 . 5 2 9 .2 415 . 3 5 6 1 .0 1 4 0.00 17 .68 V ery Adv . Die s e l 30.8 9.3 40.7 3 9 3 .3 1 7 . 991 40. 0 0 1 4. 8 6 A d v. Tu rb op r op 36 . 1 1 5.0 2 7.7 2 8 2 .4 55 . 62 7 40.00 7. 6 7 SEN S I T I V I T Y COEFF I C I E N T S ( g e PER A ) _TOC bTOC _TOC - aTOC "_TOC T_ LT_O_C_ _AC _AC _AC 6AC _W_f. _ _ _Wf c_ 6We ace 6MC _ 6C0 _ _e _Ce 6CO &-SFC S.I. Re c i p 106. 1 2 .2 0. 4 5 0.67 1 3 2. 0. 1 6 57. 12.6 0 .66 0 .58 240. 1 4 .8 0. 9 1 Rotary 89. 9 . 7 0. 3 4 0.6 7 115. 0.5 7 4 7 . II. 7 0.59 0.51 211. II.0 0 . 79 Diesel 98. 8.1 0.33 0.66 121. 0.52 51. 9.8 0.6 3 0.42 228. 8. 7 Q. 70 T _ur bo prop 8 5 , 7.6 0 .2 8 0 .5 7 128. - - 8 2. I0. 2 0 . 5 3 - - 22 0 . 8. 9 - - TOC - Total Ownership C ost, $ SFC - SpecificFuel C onsumptio n , Ib / HP - Hr AC - Acquisition C o st, $ We - EngineWeight,Ib / HP Wf - Weightof Fue l , I b Ce - EngineCost, $ / HP Cf - Fuel Cost, $ / 1b MC - EngineMain t enance Co st ,S / F lightHr CD - CoolingDrag;%A / C Drag I TABLE AI O . - SE N SITIVIT Y F ORMISSIO N 1 0 , A 6 -PLA C EBU S INE SS T W I N W L= 1 2 00 l b H = 25 0 0 0 f t V = 2 70 k t s R = 1 6 00 N.M, W / S = 5 0 l b / ft 2
1198 0 R e tp TO C *- S311 00 0 AC *- S233 000 f *- 1214 l b S H P *- 3 2 3 I
BENEFI T S REL . TO 1 9 80 RECIP. BAS E LIN E ASSUMP T IONS % Z_ T OC %Z_AC _Wf SHP SFC WR Ce MC Adv. S.I. Re c lp 1 9 .8 7 .8 3 0. 6 2 7 8 .3 61 1 .42 40.0 0 16. 5 1 Very A d v . S . I. R ec i p 26. 1 1 0.8 3 9 .2 2 63 . 334 1.2 0 4 0 . 00 1 5 . 5 6 Adv. Ro tary 26 .3 11.4 3 0 .7 2 46 .3 8 1 1.24 4 0 . 2 2 1 2 .87 V e r y Adv . Ro t a ry 35.6 15.4 45 . 6 22 1 .332 .8 3 9 41 . 5 1 8.48 Adv. D i es e l 25.5 8.4 29.2 2 76 .37 1 1.30 40 . 00 1 1 .77 Very Adv . Di ese l 30 . 9 11. 4 39 . 7 2 62 . 33 2 1 . 1 8 4 0 . 00 9. 92 A d v. Turboprop 28.0 6. 1 23 .1 2 89 .4 7 1 . 6 01 40.00 7.8 9 SE NSI T I VIT Y COEFFIC I E N T S ( % A PER A ) _TOC - 6TOC _ TOC 6 TOC _TOC _ _AC _AC _AC 6A C _ _ 6We _ _MC _Cf _CD _ _e _ 6CD _SFC _-- 6C0 S.I.Re c l p 76 . 5.2 0.3 3 0.82 I0 7 . 0.48 29. 5.2 0.44 0.2 7 214 6 .5 0. 73 Rotary 6 4. 4.0 0.2 3 0.81 96. 0. 37 24. 4. 6 0. 39 0.25 188. 5. 3 0. 6 2 Diesel 7 1. 5. 7 0.25 0,82 I01. 0.3 9 26. 5.0 0.43 0.2 7 203. 5 .6 0. 67 Turbopr o p 66. 5.8 0.29 0. 77 121. -- 2 9. 7 .5 0.48 -- 20 7 7 .5 -- TOC - Total _vnershipC o st, $ SF C - SpecificF u elCon s umpti o n, Ib / HP-Hr AC - Acq u isition Cost, $ We - EngineWeight,Ib / HP Wf - Weightof Fuel, Ib Ce - EngineCost, $ / HP Cf - Fuel Cost, $ / Ib MC - EngineMaintenan c e C o st, S / Fl i ght Hr CD - CoolingDrag; %A / C Drag TABLE A1 1 . - S E N SITIV I T Y F OR MISSION11 " , A 8 P LA CEEXEC U TIVE W L = 1 8 00 Ib H- 4 50 00 ft V - 3 8 0 kts R = 1700 N ,M. W / S = 60 Ib l ft 2 ( A DVANCED TURBOPROP TOC*- $460000 AC*- $471000 Wf*-I069 lb SHP*-973 J BEN EF I TS R EL . TO A DV. TP BA SEL IN E A SS UM P T ION S i - % ATOC % AAC % _Wf SH P S F C We Ce MC Ve ryA dv a nc e d S.I . R e c i p 13 . 9 23 . 4 7 . 4 4 61 0 . 3 3 1 . 16 40 . 00 2 9 . 46 With Ideali z e d Tu rb oc harger Ve ry A dvancedS .I . Re c t p 1.2 15 . 3 - 0.6 617 0.33 1.16 40.00 41.12 W i th 2 1 DO 0 Ft Cr i ti c al T u rb o charger Adv a ncedTurbop r op 0 0 0 9 73 0.373 0.421 40 .0 0 31.13 SEN S I T I V I TY COEFFI C I E N TS (% A PER _ ) ....... 6 TOC _TOC 6TOC 6TOC _TOC 6TOC 6AC 6 AC 6 AC _ AC _Wf _ l _f 6 Mr 6 SF C 6W e 6C e 6MC 6Cf 6C D 6S FC 6W e 6 Ce 6CD 6 SFC 6W e 6 CD S.I . Re cip * * 119 8 . 7 0 . 38 0.5 2 1 25 0 . 67 53 7.1 0.41 0. 4 0 335 13. 6 1.17 10 4 2 8.9 0. 6 6 0. 52 1 22 .... 4 7 2 9.9 0.85 30 6 2 7.6 ....
TOC- To ta l O wn e r s hi p Cos t , $ S F C- Spe cifi c Fuel Consump ti on , lb / HP - H r A C - Ac quisi ti on Cos t, $ W e - Engine W eig ht, I b / HP Wf - Weight o f Fue l , I b C e - E ng i neCo s t, $ / HP " Cf - Fu e lCos t, $ 1 1 b M C - Eng i ne Mai n t ena n c e Cos t, S / Fllg ht Hr CD - C ooll ng D rag;% A / C Drag *A d vance d T u r b opro p Ba s e ll ne E ngine.
**Ba s e d o n Ver y AdvancedS . I . R . w l th Idea l lzed Tur bo charger.
T AB LE A I2. - S E N S I T I V I TY FOR M I S S I O N12 , A 4- PLA C EHEL ICO PTE R W L = 8 00 Ib H = 2 0 00 ft - V = I 00 kts R- 3 00 N.M. W / S ....
11980 Recip TOC*= $213000 AC*= $106000 gf*- 245 Ib SHP*- 354 l I !
BENEFITS REL. TO 1980 RECIP. BASELIN E ASSUMPTIONS % A TOC % Z_AC _Z_Wf SHP SFC Wp Ce M£: A d v. S.I. Re c ip 6. 9 1 .8 i 6. 6 348 . 3 5 9 1.4 2 40.00 1 0. 67 Very A d v. S. I. Re ci p 13 . 6 4. 2 24.8 33 8 . 33 2 1 .20 40.00 10 . 33 A d v. Rot a ry 13 .8 4. 1 1 4.8 33 8 . 3 7 9 1. 12 4 0.00 8 . 8 0 V e r y A d v . R o t a ry 2 2.2 7 . 9 2 9 .2 3 22 . 3 2 9 . 745 4 0.00 6 . 10 A d v. D|e s e ] 1 5. 9 3 .8 1 8. 7 339 . 3 5 8 1 . 1 9 40.00 7 .22 V eryA d v. D ie s e l 20.0 5 . 5 2 9. 2 333 .3 1 9 1 .07 40.00 6.29 Adv. Tur bosha ft 21.3 13 . 8 -12 .8 26 1 .52 6 . 47 6 40 . 00 3 . 43 .............. ..... S E N S I T I V ITY COE FFI C I E NTS ( % A PER A ) &TOC _TOC 6TOC 6TOC _TOC _TOC _AC _AC _AC 6AC _Wf _Wf _wf._ &We _C e _MC _ &CO _ _-W e _C e 6 C D _SFC _W e 6CO S.I. Re c l p 51 . 9 .5 0.50 1. 17 9 2. -- 1 5.4 8.4 0.8 3 -- 2 36 . 8. 3 -- Rotary 4 7 . 8.0 0.4 1 1 . 17 85. -- 13 . 9 7 . 6 0. 79 -- 224. 7.8 -- Diese l 4 9 . 8.2 0.4 3 1 . 17 88 . - - 14 . 1 8.0 0.82 -- 23 3 . 6. 3 -- Turboprop 51. 8. 9 0.3 3 1.17 1 22. -- 18. 6 8. 6 0.64 -- 2 31 . 1 2.2 -- TOC - Tota l Ownership C ost, $ SFC - SpecificFue l Consumpti o n, I b / HP-Hr AC - Acquisition Cost, $ We - EngineWeigh t ,Ib / HP Wf Weightof Fuel, Ib Ce - E ngineCost, $ / HP C f - Fuel Cost, $ / Ib MC - E ngineMaintenance Co s t, S / Fl igh t Hr CD - CoolingDrag; %A / C Drag T AB LE A 13 . - SE N S I TIV I T Y F OR MI SS ION 13 , A 6 - PLACE TWIN HEL I COPTER W L = 1200 I b H = 20 0 0 f t V = 1 30 k t s R = 5 0 0 N.M . W / S ....
119 80 R ec l p T OC* - $ 503 000 AC" - $2 4 6000 Wf "-g 2 3 ib SHP*- 398 ( BENEFIT S REL.TO 1980 RECIP. BASELIN E A SS UMPTIONS i_ o ZSTOC % LSAC _Wf SHP SFC W e Ce MC Adv. S.I. Recip 16.0 7 . 3 2 5 .8 3 6 7 . 3 58 1.4 1 40.00 22 .73 Ve r y A d v . S . I . R ec i p 23 . 8 l O . g 34 . 1 352 . 33 1 1.1 8 4 0 . 00 2 1. 63 A d v.Rotary 2 3 . 3 10 .4 2 4. 7 3 5 3 . 37 8 1 . 10 4 0 . 00 1 8.44 Ver y A d v. R o tary 3 2.8 15.8 38 . 6 3 30 . 32 g . 73 8 40.00 12 . 5 0 A dv. D ie se l 2 5 .5 10.2 28.5 358 .357 1. 1 7 4 0 . 00 15.10 V e ry Adv . Di es e I 30. 2 1 2 .8 38.0 3 44 . 319 1.06 40.00 1 3 . 00 Adv. T urboshaft 27. 8 14.5 7.1 3 2 8 .48 5 . 488 40. 0 0 8.87 SEN S I T I V IT Y COEFF I C IEN TS ( % 6 PER A ) &TOC &TOC 6TOC 6TOC _TOC _TOC _AC _AC _AC 6AC '6Wf _W S_"_ - _ &We _Ce 6Me _ &CD _ _ee _ 6C O _SFC S. I . R ecip 6 5. 11. 6 _ 0 . 44 0 .5 0 119 . -- 32 " i ii. o o .75 -- 216 . 8.9 - - ..
R o tary 58. 9 .6 0. 3 6 0.50 1 0 4 . -- 2 8. 3 9 ' 8 0 .7 1 - - 202. 8. 2 -- Die s e l 61 . 1 0. 3 0. 37 0.50 1 10. -- 2 9 . 7 i 0.5 0. 7 3 -- 20 9 . 8 . 3 , - Turboprop 6g. 14. 1 0 .3 5 0.50 166. -- 4 1 . 3 1 4.8 0. 7 0 -- 21 8. 15 . 7 ,-- TOC - Total Ownershipcost, $ SF C - Specific F uel C on s umption, I b / HP-Hr AC - Acquisition Cost, $ We - EngineWeight, Ib / HP Wf - Weightof Fuel, Ib Ce - EngineCo s t , $ / HP Cf - Fuel Cost, $ / Ib MC - E n g ineMainte n ance Cos t, S / F l i g ht Hr CD - CoolingDrag; %A / C Drag j TA B LE A14 . - E STI I _ T EDANDA CTUA L VA L UES OF NISSION1 0 A IR CRA F T I MPROVENENTS FO R VERY ADVANCED TECHNOLOGY ENG I NES. EST IM ATES DER I VED FRO M ADVANCED TECHNOLOGY BA S EL IN E VALUE S PLUSSENSITIV I TYDAT A OFTABLE A I O E sti m at ed V al ue A c t u a l Va l ue S .I.R . % e T OC 28.2 26. 1 S.I. R . % _ AC 10.0 1 o.8 S.I. R. % _Wf 38 . 9 39. 2 D ie s el% A T O C 3 1.0 3 0. 9 D ie s el% AA C 10.4 11.0 D i ese l% 6 Wf 3 B . B 39.7 R ota ry% A T OC 3 5. 6 3 5. 6 R o tary% AAC 14 . 9 15 . 4 R o tary% AWf 44 . 2 45 . 6 .........• .... .. , . , _ .... _ . .. ° , _ °.............
Current Technology Advanced Technology Veryadvanced technology • Nat. aspir, below 10000ft •Turbocompound - all altitudes -Turbocompound _ all altitudes • Turbocharged over10000ft • Lean-burn • Stratifiedcharge ,,Homogeneous charge • Homogeneous charge * Multifuel • AvGAS • AvGAS ,65 0/ o coolingdragreduction • Cooling drag - 10 % of totalA / C ,50 "/ o coolingdrag reduction Figure1 . - Design featuresfor spark-ignited reciprocating (SIR)engine. All designspresented arefour-strokewith horizontallyopposed cylinders.
Advanced Technology VeryAdvanced T echnology • Stratifiedcharge • Stratifiedcharge • Multifuel capability • Multifuel capability • Turbocharged • VATturbocharged / pressure compounded • Liquidcooled (6 . 5 % drag • Liquidcooled(zerocoolingdrag) reduction) • Retracting apexseals • Conventional apexdrag Figure2. - Design featuresfor advanced rotary engines .
'\ C - 79-5 3 7 Advanced Technology VeryAdvanced Technology • Radial design • Radialdesign • Two-stroke cycle • Two-stroke cycle • Limited cooling(.50 % drag * Limited cooling(65 % drag reduction) reductbn) • Turbocharged (7 . 25:1P / P • Turbocharged (g:l P / Pat at 25000ft) 25000ft) • Conventional combustor • Catalytic combustor in in T.C . loop T.C. loop • Conventional lubrication • Syntheticoil Figure3. - Design featuresfor lightweight dieselengines .
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Figure 5. - Design features flldvanced turtllne engines.
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.6-- Turtx _ rop (I" turr e t t K hnOlOgy " VA D Ve r y ad_ r _ mc ed tech n ology •5 -- Sl _ rk - lgnltion piston Rotary Z011, Obs e l I_ ADV A dva n c e d te ch nology _ '/// r /// . 3 y // , r /i l • F i// f // J , ii , 111J ,izs F Ill CT ADV CT ADVVAD CT ADV VAD CT ADV VAD Figure 6. - Projected BS F C I m tx'ov _ ments foradvanc ed engines at 350 m aximu m ratedpower. F light conditions ar e ;5 000foot a ltitude end ;5 0 knots endIncl u d e s pr e ssurlz a tlon andauxiliary power p e nalti e s.
;5 0 0 0 Feet cruise altitude .50- TP Tu r bop r op S S . I . Recip R Ro ta ry D Die se l & •. - . ..[p . 40 -- *V e ryad v anc e d t e ch n ol o g y " 5 _ R D _ S R _
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"'very ilrlvanted technoloqy " lU ..... ---'----I.---.J-----I ~oo 300 100 200 300 400 Maximum rated. hp Ma~lmum rilted. hp figure 9 .• EIIed of sin on englnll Wllighl for Figure 10. • 1.S$umel! tnglne cosl 1197111.
thll advanced engines.
CT Baseline nonexistent Turboprop Diesel Rotary CT Current technology ADV Advanced technology VAD Very advanced technology 5.1. Reclp CT AOV VAO Figure It - ProJected time between OY'erhaul (TIlOI for advanced engines at 350 maximum rated power. ' Excludes OY'erhaul reserve Yoo 2DO 3DO Maximum rated shaft horsepower, SLS Figure 12. - Assumed engine maintenance cosl 50000-- • O To d ay's N C & m issio n s • Futuris t ic N C & m issi on s 40 0 0O -- & Helicopter m issions 30000-- Cu rr e ntaltitud e limit _= _ ' 20000-- Sin g l e s " 3
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0 100 2 00 3 00 . 400 Des i gn c rui ses pe ed , k not s Figure 1 3.- Cruis e spe ed a n da lti t ude spectru m con s ide r ed.
Cruise a lt itude, ft 10 0 (3 0 1 6 0 0 0 25 000 A ._ A 50 -- ZI 0 Adva nced technology • Veryadvanc ed technology
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1 2 5 9 3 4 6 10 8 Mis s i on Fi gur e1 4 . - Mi s si on fu el re du dl on_ e n tl a l o f ad v an ce d e ng i n e s .
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Figure 18. - Helicopter Improvement potential.
1.4 _ ,- L4 _ B es elln ez T ur bo proplt L4 L 2 L 2 -- L2 _ ., _ ' 45000 t t '_ .8 _ .8 -- _ _..,,.,_. .8 _ i __ Turboprop == --' = S I R- Z lO00 --..i SIR- Ideal .6 -- .6 _ . 6 3 0 3 5 40 3 0 3 5 40 4 5 30 3 5 40 103 D e sign cruls e al t itude, ft Figure19. - Eff e c t ofhighaltitude onfore c ast e d aircraftImp r mmm on ts. Missi on 1118-pl a ce executiv e tw in, 3 80knots , 1 1 00 n . m.L 1!80 SIR Bas e lin e 50-- 40 -- 40 _° zc- _ ..= 2 c 0 S.T. R e clp 1 3 Di e s e l It -- I C -- 0 Rotar y Z _ Turboprop 0 Advanc e d tec hn ology I I i I " _eryadvanced t e chnoloc_y I 0.3 .4 .5 0' 3 .4 .5 BS F C, ro / hp - hr Figur e 20. - Effect of vary i ng engin e BS F C on T OC a n dmi s si on f uel. Mi s sion 1016 plac e bu s in e s s twin).
1960 SIRBas e line 50-- 50-- I 40-- 40-- 0 S. L R _ lp O D i esel t O -- 10-- O Ro t ary L _ Turboprop 0 Advanced t e chnolo g y • Veryadva n ced technology
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0 l 2 0 1 2 Engine specific weigh t,r d hp Rgure2L - Effedofvarying e ngine Install ed sp e cific M ightm T OC e nd m lsslmfuelr e duction pot e n- tial. Mission 101 6 place busin e ss twin).
1980 SIRBaseline 40-- 0 S.I. R ec lp C] Di e sel lO- -- 0 Rotary Z _ Turbopr op 0 Advanc ed t ec hnology 40 50 O lO 20
% I I I I • ivery ed ,, nled tec .no l_ I
EngineInitial OEM cost, t / hp Engine main t e n ance c os t,t / f t-hr F igure 22. - Effect of va ryi n g e n gine OEM andmaint e na n ce cost on T OC. Mission 10(6place business twin).
'\ 1 _ 80 SIRB a s e ll n , 't _ 2 0 -- d i s p lac e ment e ngi n es . _ _ Rot a ry - T op SIR- Bottom J . Ve r yId v t n c ed p m ltlve < ---- -- A ¢ ,,anc _ turboprop
- - l i i I I
50 -- - //
!
Ba se lin e _ 1( /// B aseli n e 0 200 250 3 0 0 : 1 5 0 20 0 250 3 0 0 3 5 0 400 Crui s e s p eed , knots Figure 23. - Eff e ct ofcruis e s peed onfor e cast e d aircraftImprovem e nt s . Mis s i on ]0 (6place business tw i n) 30 -- 1 980 SIRBaseline V e rya d vance d positive _J Z O -- dlsp_ e m e n! e ngin e s . _ Di e s e l - M i d dl e l O /////////// ///// /////// /_ _ Ro t a ry - Top " _ 0 -- SIR - Bott o m _ --- A dv a nc ed turbop r op
., - 1 o I •
/ ; 50_ 50-- 4 o 4 ( _ -///////////////////// _ .
" = Bas e line 10 Baseline
_! _ z _ -
6 _ _ =_ l O -- ] , J I
1900 13 00 1600 1900 Rang e , N ml F igur e Z4. - Eff e ct of range onfo re cast e d aircra ft Improv e ment s . Mlsslon 10 , ( 6 pl a ce b u sin e ss twin 1.
- 5 0 I _ 0 SIRBas e line 20 S.I. Reclp dies e l I / / ' // " II Veryadvanced r /f// @ V e ry ad vanced rotary 10 _ _ Advanc ed turboprop I / I Base il ne I 0 :' 500 10130 Utilization,hr l yr Figure25. - Effect ofaircraft utilization on aircraft Improve- m e nts.Misslm10 . ,(6place business twin).
\ 1 . R eport No . 2. Governmen t Acc l m l o n No . 3. Rec l piem' s Cata l og No .
NASA TM-81 5 84 4. Title e nd S u btit l e 5. Rep o rt Dire COMPA RISO N SO F F O UR A LTE RN AT IVE POWE RP LAN T O ct o ber 1980 TYPE S FO R F UTUR E G E NER A L A VI AT IO N A I R CR A FT 6. i ,,. ' f orm l og Or p. i . z ,ti o . Cod , 7 . A u t ho r (s) 8. Pro' form ing Or gan iz at io n R e por t No.
T. O . Wlcke n he i ser , G. Kntp , R. M . P lencner , and W . C . Strack E - 561 10. Work Unit No.
9 , Performing Org a nization Name a nd Address National Aeronautics and Space Administration 11. C on tract or Grant No.
Lewis R esearch Center Cleveland , O h i o 44135 1 3 . Type of Report _ n clPeriod Cove _ ed 12 . S po n s oringAgency Na me Imd Address T ec hni c al M e moran d um National A e ron a u t i c s a ndSp ace Ad m ini s tration 14. SponsoringAgency Code Washington , D. C . 2 0 546 15. Su pplementaryNotes 16. Abs t ract Re centl y c o mpleted NAS A -spons o red co n cept ual st udies have cul min ated in the ide n t i f i c ation o f pro mi s ing new technologies for future spark ignition , diesel, rotary, an d t ur bine engines. This paper re po rts the res ul t s of a NASA in -hou s e preli min ary assess m ent study th at co m pares these fo u r pow erpl an t t y pes in several gener al avi at ion applic ati ons. The ev al u at ion con s ist ed of in - stall in g each po werpla n t ty pe in " ru b beriz ed" aircraft w hich are si z ed to acco m plish fixed mis- s io ns . The pri m ar y ev al u at ion criter i a in clude projected a i rcr af t cost , tot al ownersh i p cost, and mis s ion fuel.
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17 . K e y W o rds (Suggested by Author( s )) 1 8. Disl _ 'ibution State me nt T ur bo pr o p e n gine s ; Die s el e n gines; Uncl as s i fi ed - unlimi t ed Rotary engines; Reciproc ating engines; STAR C at egory 07 Co m p ari son 19. Security Oe s s if. (of thisreport) 20. SecurityCl a sslf. (of this page) 21. No. of Pages 22. Price" Un classifi ed U n cl as s if ied " Forsale bythe National Technical Inf or mation Service , S pringfield , Vi rg inia 22161 ,r National Aeronaut i cs and S PE C IAL FO URTH C LA SS MAIL Po s tag e an( ' ¢ _ Space Administration 8OOK N ati o n a l A e ro n a u ti cs a n d Spa ce Admin is tr a t i on Washington, D.C. NASA451 20546 O ff i cial B usi n e ss P e nal ty for Priva t e Us e, $300 r