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-7 4.-:7NASACP2092c.; 1 INASA Conference Publication 2092

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Aeropropulsion 1979

Proceedings of a conference held atNASA Lewis Research CenterCleveland, OhioMay 15-16, 1979

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NASA Conference Publication 2092

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Aeropropulsion 1979

Proceedings of a conference held atNASA Lewis Research CenterCleveland, OhioMay 15-16, 1979

National Aeronauticsand Space AdministrationScientific and TechnicalInformation Branch1979

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FOREWORDThe Lewis Research Center has a 37-year heri-tage of providing advances in aeronautical pro-pulsion from the research activities of its staffand its university and industrial grantees andcontractors. These advances have helped createthe preeminence in aeronautics that has contri-buted to our national defense, has provided swift,reliable transportation for our people and theirgoods, and has so greatly aided our position ininternational trade.Although the results of the Center's projectsand programs are reported as they are obtained,from time to time a conference such as this oneaffords the opportunity for a broad overview andan in-depth interpretation of a large body of re-sults and for informal discussions about the sub-ject. We hope you will find the material given here-in informative and useful and that you will in-quire further about whatever may be important toyou.

John F. McCarthy, Jr.Director

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PageXII. HYPERSONIC PROPULSIONH. Lee Beach, Jr.. . . . . . . . . . . . . . . . . . . . 387XIII. VERTICAL TAKEOFF AND LANDING (VTOL) PROPULSIONTECHNOLOGYCarl C. Ciepluch, John M. Abbott, Royce D. Moore, andJames F. Sellers . . . . . . . . . . . . . . . . . . . . 409

XIV. HIGH-PERFORMANCE-VEHICLE TECHNOLOGYLouis A. Povinelli . . . . . . . . . . . . . . . . . . . 445RESEARCH AND DEVELOPMENT CONTRACTORS AND GRANTEES . . . . 463

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I. AIRCRAFT ENERGY EFFICIENCY (ACEE) STATUS REPORTDonald L. Nored, James F. Dugan, Jr., Neal T. Saunders,and Joseph A. Ziemianski

National Aeronautics and Space AdministrationLewis Research Center

Efficient air transportation is of national concern since com-mercial aircraft constitute a primary segment of public trans-portation. However, the viability of the U.S. air transporta-tion industry and its ability to handle future traffic growthare threatened by rapidly escalating fuel prices. Figure I-l,taken from Civil Aeronautics Board (CAB) data (ref. l), illus-trates the recent trend in fuel prices. From 1973 to 1975, fuelprices essentially tripled internationally and doubled domesti-cally. They have since continued to increase about 12 percenteach year. Such increases are directly reflected in the air-craft direct operating cost (DOC).The major elements of the DOC are shown in figure I-2 (ref. 2).The elements are expressed in cents per available ton-mile ofthe mix of passengers and cargo for each year. Before 1973,these elements were about equal in their contribution to DOC.Starting in 1973, however, with the OPEC embargo and the subse-quent large increases in fuel prices, these historical relation-ships changed. Fuel prices began to escalate faster than therate of inflation and even faster than increased productivitycould reduce them. As a result, not only has the total DOC in-creased, but fuel costs have become a much larger percentage ofthe DOC. For example, as shown in figure I-3 for the Boeing 727flying commercially in the domestic market, in 1973, fuel con-tributed about 26 percent of the DOC. By 1977, this fuel per-centage had increased to about 41 percent.The percentage of the DOC related to fuel is expected to contin-ue to increase in the future. Large increases in aircraft fuelneeds are projected, as a result of an expected rapid growth inair travel. Aircraft fuel, however, is derived completely frompetroleum, a dwindling national resource. Hence, scarcity offuel and resulting higher prices are foreseen. This situationwill be aggravated by various artificial (e.g., OPEC initiated)price increases. Commercial aircraft currently use over 10 bil-lion gallons of fuel per year (ref. 1). This is conservativelyprojected to more than double by the year 2000. Obviously, anincrease in fuel supply would alleviate the problems of cost andavailability. For this reason, there is interest in producing

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synthetic jet fuel from our large national resources of coal andoil shale. However, such synthetic fuel will not be availablein large quantities for some years, and until it is, the air-craft industry must use its petroleum-based fuel more efficient-lY* Increases in fuel efficiency will help counteract the ef-fect of rising fuel prices on the DOC while also alleviating thecritical problem of future fuel availability.To answer this need, NASA started the Aircraft Energy Efficiency(ACEE) program in 1976. This program is a focused reponse tothe current importance of fuel effiency in aeronautics, for fuelconservation in general as well as for its effect on commercialaircraft operating economics. Included in the program are sixmajor projects aimed at providing technology for more fuel-conservative aircraft and propulsion systems for future commer-cial airline service (ref. 3). Three of the projects - in theareas of aerodynamics and aircraft structures - are managed bythe Langley Research Center. These aircraft-related projectsare Energy Efficient Transport, Laminar Flow Control, and Com-posite Structures. The other three projects are propulsion re-lated, as indicated in figure I-4, and are managed by the LewisResearch Center. They are (1) Engine Component Improvement(ECI) , directed at improving the engine components and the per-formance retention of existing engines; (2) Energy EfficientEngine (E3), directed at providing the technology base for thenext generation of turbofan engines; and (3) Advanced Turboprop,directed at advancing the technology of turboprop-powered air-craft to a point suitable for commercial airline service. It isthese three projects that are discussed in this paper in somedetail.ENGINE COMPONENT IMPROVEMENTThe EC1 project is investigating the potentiai for reducing fuelusage in existing engines. It is these engines that will usethe bulk of the commercial aircraft fuel between now and 1990.This project is expected to result in technology, by 1980 to1982, that will permit as much as a 5 percent fuel savings overthe operational life of the engines. Such fuel savings will beachieved by improving engine performance as well as by mini-mizing engine performance degradation in service. Thus, theproject has two parts: (1) performance improvement and (2) en-gine diagnostics (fig. I-5).Performance ImprovementThe performance improvement part of the EC1 project is directedat developing the technology for improved, more fuel-efficientengine components for early introduction into commercial ser-vice. The technical effort is being conducted by the manu-

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facturers (General Electric and Pratt & Whitney) of the threeengines that power most of the current commercial fleet (fig.I-6). These engines are the Pratt & Whitney JT8D - used on theBoeing 727 and 737 and the Douglas DC-9 - and the Pratt &Whitney JT9D and the General Electric CF6 - used on the Boeing747, Douglas DC-lo, and Airbus Industries A300. Derivatives ofthese latter two large, high-bypass-ratio engines will alsopower the newer airplanes such as the Boeing 757 and 767 and theAirbus Industries A310.The components being investigated are shown in figure I-7, andinclude most major components of the engine. The specific com-ponent improvements will be derived generally from improvedaerodynamics, reduced clearances, more effective cooling, andimproved materials.An extensive feasibility and screening analysis was conducted ona variety of concepts by both General Electric and Pratt &Whitney before a specific concept was selected (refs. 4 and 5).This analysis was a team effort, with each engine manufacturerbeing assisted by NASA, two aircraft manufacturers (Boeing andDouglas), and a number of airlines (United, American, TWA, PanAm, and Eastern). Technical merits (e.g., performance, weight,and maintenance) as well as economic merits (e.g., airline re-turn on investment, direct operating cost, and payback period)were investigated. From considerations such as potential fuelsavings, economic benefits, and cost of development, NASA se-lected 16 component improvement concepts for technology develop-ment by General Electric and Pratt & Whitney.For the JT8D engine, four concepts were selected (table I-l) fordevelopment:

(1) An improved outer air seal for the high-pressure turbine(2) A new high-pressure-turbine blade cooling concept, where-in the cooling air is discharged at the root of the blade(3) An aerodynamically improved DC9/JT8D reverser stangfairing that incorporates advanced composite materialsfor lighter weight(4) Longer blades and abradable trenched rubstrips on thehigh-pressure compressor

Concepts 1, 2, and 4 are being developed by Pratt & Whitney; thereverser stang fairing is to be developed by the DouglasAircraft Co.Technology development has been completed on the improved JT8Douter air seal. Figure I-8 shows some features of the seal.Through the use of honeycomb seal material plus additional

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knife-edges, an improved labyrinth arrangement was effectivelyadded to the seal. Also, the blade cooling flow, which had beendischarged completely from the blade tip, was rerouted to permitsome discharge from the blade suction surface. Results havebeen obtained from back-to-back engine tests of both the currentand improved configurations, as shown in figure I-9, over arange of thrust levels. At the normal (90 percent) cruisepoint, a specific fuel consumption (SFC) reduction of 0.6 per-cent was achieved. This reduction (though small) is equivalentto 6 percent of the total net income of the U.S. domestic air-lines in 1977 (at an average fuel cost of 50d/gailon). It isalso equivalent to an airline fuel savings of 200 million gal-lons by the end of the century, if the seal is incorporated bothinto new-production JT8D engines and into older engines throughretrofit, where economically feasible.For the JT9D engine, four concepts were also selected (tableI-l) for development by Pratt & Whitney:

(1) Active clearance control for the high-pressure turbine(2) A new fan incorporating only a single shroud along withlow-aspect-ratio blades(3) Ceramic thermal-barrier coatings on the vane end wallsin the high-pressure turbine(4) Ceramic outer air seals for the high-pressure turbine

Development of the active clearance control concept has beencompleted. As indicated in figure I-10, this is a technique toreduce tip clearances during cruise by cooling, and henceshrinking, the case. During the transient portions of flight(takeoff and landing), the cooling air is reduced or elimi-nated. This increases the tip clearances between the turbineblade and shroud and minimizes the potential for rubbing.The current and improved JT9D configurations are shown in figureI-11. The significant features of the improved configurationare

(1) A modified support ring that permits greater reductionin tip clearance(2) A doubling of the cooling flow(3) A new configuration for the air supply tubes thatprovides shorter impingement distances and a moreeffective use of the cooling air

Results are shown in figure I-12 for back-to-back tests of theimproved and current configurations. These results indicate4

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that cruise SFC can be reduced 0.65 percent over a wide range ofcruise thrust settings.For the CF6 engine, eight concepts (table I-2) were selected:

(1) A short-core nozzle that permits reduction in weight andscrubbing drag(2) A new front mount that improves the load distributionaround the compressor case and results in reduced tipclearances(3) A new fan with improved aerodynamic design(4) Improved aerodynamics for the high-pressure turbine(5) Improved roundness control for the high-pressure turbine(6) Reduced compressor bleed by recirculation of air in theDC-10 cabin air-conditioning system(7) Active clearance control for the high-pressure turbine(8) Active clearance control for the low-pressure turbine

Concept 6 is under development by the Douglas Aircraft Co.; allthe other concepts are to be developed by General Electric.Technology development has been completed on the first threeconcepts. The changes in the core nozzle resulted in a 0.9 per-cent SFC reduction at cruise, as demonstrated by flight tests onboth the DC-10 and A300 aircraft. The new front mount can pro-vide a 0.3 percent cruise SFC reduction. The improved fan canprovide a 2.0 percent cruise SFC reduction. Figure I-13 liststhe improved features for the new fan concept: basically, im-proved airfoils, rearward placement of the current single shroudto minimize blockage, and addition of a fan-case stiffening ringto keep the case round (hence, reducing tip clearances). Re-sults of sea-level engine tests are shown in figure I-14. Theindicated 3.2 percent SFC reduction at 40 000 pounds of sea-level thrust is equivalent to a 2 percent cruise SFC reduction.According to a projection based on various market predictions bythe engine manufacturers, the performance improvement part ofthe Engine Component Improvement project should provide the air-line industry with a cumulative fuel savings of at least 7 bil-lion gallons by the year 2000. This savings would occur if allthe concepts complete the technology development phase success-fully, are carried through the certification phase by the manu-facturers, and are successfully introduced into commercial ser-vice on future engine models or through retrofit.

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Engine DiagnosticsThe engine disagostics part of the EC1 project is directed atidentifying and quantifying the sources of the performance de-terioration that occurs with time in the JT9D and CF6 engines.The effort will also provide design methodology and maintenancepractices for minimizing such deterioration in current and fu-ture engines. As with performance improvement, NASA has majorcontracts with both Pratt & Whitney and General Electric fortesting, data analysis, and modeling.The major contributors to performance degradation are shown infigure I-15. Examples of these degradation mechanisms are shownin figure I-16. In the cold section of the engine, foreign ob-ject damage, erosion, and surface roughness are significant. Inthe hot section of the engine, thermal distortion is one of thepredominant degradation mechanisms - causing, for example,warpage or distortion of vanes. Clearance increases occurthrough the entire engine, as a result of blades rubbing withthe outer shrouds. These clearance increases result in effi-ciency losses.In investigating these contributors to performance degradation,the general approach is as follows:

(1) Gather existing (historical) data from airline in-flightrecordings and from ground test cells at both airlineand engine overhaul shops. (To date, this has been donefor about one-third of the world's aircraft that useJT9D and CF6 engines.) Also, conduct inspections on aselected used parts that contribute to performancedegradation.

(2) Augment this information with specia l engine tests andinspections in order to evaluate the effects of de-teriorated components, and subsequent refurbishment, onboth overall performance and engine module performance.(Overall performance deterioration was evaluated by se-curing an engine from an airline, testing it, conductinga complete teardown, and inspecting it. The engine wasthen refurbished and/or reassembled, retested, and re-turned to the airline. Module performance degradationwas evaluated by tests before and after used moduleswere replaced.)

(3) Assess the causes of performance degradation during thefirst flight or flights of the aircraft as the enginestructure first responds to the flight environment(4) Assess the causes of long-term performance degradationas a function of both cycles and time

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(5) Determine the effects of deteriorated parts on moduleperformance(6) Establish statistical trends, analytical models, anddesign criteria, with associated correlations of theeffect of maintenance practices on SFC losses

The results of the module replacement and refurbishment test onthe CF6 engine are shown in figure I-17. The numbers in paren-theses represent the number of test modules. Simply cleaningthe fan blades and recontouring the leading edges resulted in aspecific-fuel-consumption reduction of 0.3 percent. Replacingthe low-pressure turbine with new or refurbished modules result-ed in a SFC reduction of 0.4 percent. The high-pressure-turbineresuits are not available yet, since testing and analysis arestill under way.How the various performance degradation mechanisms - such asclearance increases, erosion and airfoil roughness, and thermaldistortion - affect the cruise SFC deterioration of the JT9Dengine modules is shown in figure I-18. These results, averagedfor a large number of engines, are based on used-parts analysisand prerepair tests (i.e., the historical data). Clearancechanges affect all the engine modules from the very firstflight. As the engine matures, however, thermal distortion be-comes of significance for the high-pressure turbine, and erosionand airfoil roughness become of significance for the high-pressure compressor. Although the data indicate that the high-and low-pressure-turbine modules do not deteriorate much beyondthe 1000th flight, this is only because these modules are nor-mally replaced every 1000 to 2000 flights. Thus, what is shownare deterioration values where the high-pressure turbine hasbeen replaced between the 1000th and 3000th flights.Percent cruise SFC deterioration as a function of the number offlight cycles (where a flight cycle is a takeoff, flight, andlanding) is shown in figure I-19. As evident, clearance in-creases cause more than 50 percent of the overall deteriorationexperienced in the JT9D engine.The major causes of these service-related changes in gas-pathclearances are the loads on the engine in flight. These arenacelle aerodynamic loads, experienced during airplane rotation,as well as inertia loads, such as "g" and gyroscopic loads.These loads, schematically shown in figure I-20, cause the en-gine to bend and distort and the blade to rub against the outershroud. They thus produce increased clearances and attendantperformance losses.Pratt & Whitney and Boeing - in a cooperative program - used aNASTRAN finite-element structural model (fig. I-21) to analyzethese flight-load effects on the JT9D engine. The NASTRAN model

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was used to calculate the deflections or changes in clearancesthat occur throughout the engine, for both steady-state and dy-namic conditions. These clearance changes were then convertedto performance losses. Shown in table I-3 are the NASTRAN theo-retical results. In the steady-state case, the nacelle aero-dynamic load affected all engine stages and contributed to about87 percent of the total performance loss. The other two loadingconditions were of lesser significance. In the dynamic cases,such as wind gust and hard landings, there were no significantchanges from the steady-state results.To provide a data base for validating the NASTRAN theoreticalsteady-state load results, Pratt & Whitney will subject a JT9Dengine to a simulated aerodynamic loads test. Aerodynamic loadswill be simulated by placing a series of supporting bands(bellybands) around the inlet. These bands will then be pulledto exert loads on the inlet to simulate nacelle loading condi-tions (fig. I-22). As the engine is run through a performancetest, X-ray and laser proximity probes will be used to determinechanges in tip clearances. Additional instrumentation will beadded to the engine to determine the engine and module perfor-mance changes that occur as the aerodynamic loads are simulated.A summary of engine diagnostics results to date for both theJT9D and CF6 engines is shown in figure I-23. Shown are thestatistically averaged engine performance deterioration trendsas a function of engine cycles. There is about a 3 percent in-crease in SFC after 3000 flight cycles, and about 1 percent isrecovered during a normal overhaul (which is typically done onthe engine hot section). Many factors influence the shape andvalues of the curves: Airline overhaul practice, route struc-ture, derated takeoff, engine module mix, airplane model, andengine location are but a few.The output of the engine diagnostics effort is expected to pro-vide industry with data necessary to cost effectively restorethe performance of current engines. The magnitude and locationof performance deterioration will also be pinpointed, and themodular performance analyses techniques required to diagnose therelated performance losses associated with the engine will beimproved. For derivative and future engines, this effort - byidentifying unique degradation mechanisms and developing usage-related deterioration models - will provide design tools forimproving performance retention.ENERGY EFFICIENT ENGINEThe objective of the Energy Efficient Engine (E3) project isto provide an advanced technology base for a new generation offuel-conservative turoofan engines for commercial transports.Specifically, this project involves aggressive development of

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advanced component technologies, followed by integration andtesting as complete systems. A technology readiness date of1983 has been set for completing these activities. At thattime, the advanced component technologies will have been devel-oped and demonstrated to a point where they are suitable for usein a future commercial engine development. In the late 1980's,E3 technology could appear in new advanced commercial turbofanengines, or perhaps even a few years sooner in advanced deriv-ative versions of current engines. Benefits from E3 tech-nology would then start to accrue and would grow rapidly duringthe 1990's.Goals were established at the outset of the project to guide theselection of engine cycles and configurations and to serve as afocus for the technology efforts. These goals recognized thatfuture engines must be not only fuel efficient, but also econom-ically attractive to the airlines and environmentally acceptableto the public. Goals for fuel savings are

(1) At least a 12 percent reduction in SFC for newly manu-factured engines(2) At least a 50 percent improvement in performance reten-tion over the lives of the engines

To provide economic incentives, NASA established a goal of atleast a 5 percent reduction in the DOC. These fuel and economicgoals are relative to current high-bypass-ratio turbofan engines(specifically, the JT9D-7A and the CF6-50C engine models).There is a high degree of uncertainty as to future environmentalrequirements. As a result, since any future advanced turbofanengine must meet the requirements prevailing at that time, NASAselected, as goals, the most stringent limits for noise andemissions that have been proposed to data by either the FederalAviation Administration or the Environmental Protection Agency.These are the FAA's FAR-36 (1978) standards for noise and theEPA's proposed 1981 standards for emissions.The E3 project is being accomplished through parallel con-tracts with the two U.S. manufacturers of large, high-bypass-ratio engines: General Electric and Pratt & whtney. The effortitself is structured and scheduled as shown in figure I-24. Thefirst element involves defining a baseline design for a futureadvanced propulsion system. This baseline design serves as afocus for defining technology needs, performance potential, com-ponent interfaces, and system trade-offs. To arrive at optimumcycle characteristics and engine configurations, both GeneralElectric and Pratt & Whitney conducted trade-off evaluations, inwhich they concentrated on fuel usage and DOC for E3 technol-ogy engines operating on potential 1990-era aircraft. Each com-pany was assisted by Boeing, Douglas, and Lockheed in conducting

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these trade-offs. The initial design effort has been completed,but lower-level work will be conducted during the remainder ofthe project to support the experimental efforts and to use datafrom those efforts for refining the initial propulsion-systembaseline design. Thus, an on-going appraisal of the technologystatus and potential will be obtained. Component technologyneeds have now been well defined, and work has started on thesecond element of the project. This element - which comprisesthe major portion of the project - involves design and extensiveexperimental development of the major engine components. Theintent is to substantially improve the performance of each com-ponent.In the third element, the components will be integrated to as-sess their interactive effects and to work on those mechanicaltechnologies that are unique to the integrated system. First,the high-pressure compressor, the combustor, and the high-pressure turbine will be assembled in a core (or high spool)package. Then the core will be integrated with the low-spoolcomponents (fan, booster or low-pressure compressor, low-pressure turbine, and mixer) and a metal "boilerplate" nacelle.This integrated core - low-spool package (ICLS) will then betested to assess the overall performance, component interac-tions, and design integrity of the mechanical systems.General Electric ConfigurationThe initial baseline propulsion-system design has been fin-ished. Figure I-25 illustrates General Electric's E3 config-uration, which resulted from that effort. Their configurationuses two spools plus a mixer and a long-duct nacelle. A mixeris included because, according to results from the initial per-formance analyses, exhaust-gas mixing offered the potential forabout a 3 percent reduction in SFC from that of an unmixed en-gine.Table I-4 lists the cycle characteristics of General Electric'sE3 configuration. As a comparison, values for General Elec-tric's reference engine - the CF6-50C - are also shown.bypass ratio of the E3 Thehas been increased substantially, ashas the compressor pressure ratio, over those of the CF6-50C.Turbine temperatures have been increased only moderately. Cycletrade-off studies indicated that, if the DOC and performance-retention goals are to be met, only moderate increases could betolerated even with improved materials. A thrust of 36 500pounds was used to size the hardware for this project. Thetechnologies, however, are scalable over a wide range; hence,commercial engines based on E3 technology could have whateverappropriate thrust is required to meet airline needs. For thisparticular configuration, General Electric predicts a 14.2 per-cent cruise SFC reduction over the CF6-50C SFC.

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The advanced technologies needed for the high-pressure-core com-ponents are listed in figure I-26. 'The aggressive compressordesign has a compression ratio of 23 in only 10 stages. So thatit can achieve a polytropic efficiency of over 90 percent, itwill have advanced, low-aspect-ratio, highly loaded airfoils andwill be made of advanced materials. In addition, active clear-ance control will be used on stages 6 to 10 in order to achievemuch tighter clearances than those used in current engines.The combustor is a two-zone, double-annular configuration. Thisrather complex configuration is required to meet the stringentemission goals, which will be very difficult to achieve on anengine with an overall pressure ratio of 36. This combustorrequires several technology advances, including improved dif-fuser designs and a segmented (or "shingled") liner, in order toachieve longer life.The high-pressure turbine has two stages and is designeu for anefficiency of over 92 percent at maximum cruise conditions.This high efficiency requires a number of technology advancesincluding low leakages, improved cooling concepts, and advancedactive clearance controls. Also, several advances in materialsand fabrication processes are required (directionally solidifiedblades, ceramic shrouds, and near-net-shape rotors that use hotisostatic pressing of powder-metallurgy alloys).Similar aggressiveness in technology advances is required forthe General Electric low-spool components, as shown in figureI-27. Particularly important are the fan and the booster. Thefan uses low tip speeds and a low placement of the midspan dam-per to achieve over 88 percent efficiency. The novel quarter-stage "island" booster design serves a dual function: It aidsmatching of the fan and core streams, and it centrifuges anyforeign objects away from the core stream and thereby helps re-duce foreign-object-damage erosion in the core. The five-stage,low-pressure turbine has been specially configured to reducenoise. And, as in the high-pressure compressor and turbine, anactive clearance control system will be used to permit closerrunning clearances. The mixer will be designed to achieve 75percent mixing effectiveness and very low pressure losses in ashort length.Pratt & Whitney ConfigurationThe Pratt & Whitney E3 design that resulted from the initialdesign and definition effort is illustrated in figure I-28. Itis similar to the General Electric design in that it is a two-spool, mixed-flow configuration. This design, like GeneralElectric's, has also stressed greater durability, performanceretention, and ease of maintenance. Both designs feature ashort, stiff, straddle-mounted core with five main bearings

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located in easily accessible bearing compartments. Specialattention also has been given to structural load paths in orderto minimize the bending and twisting encountered in current en-gines, thus improving performance retention. Both designs alsofeature careful attention to active and passive clearance con-trol as an aid in retaining performance.The cycle characteristics of Pratt & Whitney's baseline designare listed in table I-5. Again, for comparison, values fortheir reference engine - in this case the JT9D-i'A - are alsoshown. The predicted cruise SFC reduction for the E3 designis 14.9 percent below that of the JT9D-7A.Advanced-technology features in the Pratt & Whitney E3 coreconfiguration are shown in figure I-29. This core includes alo-stage compressor with a compression ratio of 14. A number ofadvanced-technology features - including supercritical airfoils,trenched cases, and active clearance controls - help achieve apolytropic efficiency level of over 91 percent. The combustoralso depends on a two-zone configuration for emissions control(in a similar manner to General Electric's combustor), but thePratt & Whitney design has two burning zones arranged axially inseries. Several technology advances are also planned for thiscombustor, including carburetor-type fuel nozzles and an ad-vanced liner configuration that uses improved cooling conceptsand segmented panels to extend life.Pratt & Whitney selected a single stage for the hign-pressureturbine in order to gain the cost and maintenance benefits asso-ciated with fewer hot-section parts. However, a cooled turbineefficiency of over 88 percent must be reached if the total en-gine performance required for fuel savings is to be achieved.Reaching this high efficiency in a single stage will require ad-vances in several areas, including advanced airfoil designs forthe transonic airflow in this turbine, improved cooling schemes,and several advanced materials. Pratt & Whitney's single-crystal alloys will be used for the blades and vanes. An im-proved active clearance-control system is also planned for thisturbine.Advanced-technolo y4 features in the iow-spool components of thePratt & Whitney E design are shown in figure I-30. The fanblades are unshrouded. Eliminating the midspan dampers, used inthe fan blades on all current engines, can improve efficiencyand hence reduce SFC nearly 1 percent. However, the designs ofunshrouded blades invariably result in very large, long-chordconfigurations that tend to increase engine weight. Thus, ad-vanced design approaches are needed to generate lighter weightfan-blade configurations that are resistant to both aeroelasticflutter and foreign-object damage. Pratt & Wnitney has devel-oped a promising design approach that uses hollow titanium air-foils, and work is now under way to develop the manufacturing

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for such unique blades. To date, the most promisingis to fabricate the hollow blades from diffusion-bondedother low-spool components include a four-stage, low-compressor driven by a four-stage, counter-rotating,The approach to exhaust-gas mixi,ng isextremely aggressive, calling for a design mixing effec-of 85 percent and low pressure losses. These designare to be achieved by using a short, scalloped mixer tothe weight and drag penaities of the required long-duct

component design phase is well under way, and both Generaland Pratt & Whitney are conducting technology investi-in support of their design work. Figure I-31 illus-an example of such work at General Electric in support offan-frame acoustic design. Current engines have aboutas many vanes as fan blades for minimum fan noise. Alsoan axial spacing of about 1 chord width between theand blades, along with separate struts to carry loads toan case. In the E3 design, General Electric will use avane-blade ratio, as indicated, and a much wider axialThis design permits the vanes to be integrated withstruts and hence reduces fan weight and cost.tests of these two configurations were recently con-and the results are shown in figure I-32 for perceivedlevel as a function of acoustic angle. The upper plotresults for the approach condition; the lower plot showsfor the takeoff condition. The two configurations gavesimilar acoustic results. Thus, GE's E3 fan is beingwith a vane-blade ratio of about 1.1 and very wideas illustrated in figure I-31.

mportant series of supporting technology tests con-by GE involves the combustor design. GE's prime com-design is the double-annular configuration, shown at thein figure I-33. Principles of this two-zone design wereunder NASA's Experimental Clean Combustor program.results of that program,ditions for E3, when extrapolated to the higherindicate that this configurationpotential for meeting the stringent E3 emissions goals.the two separate combustion zones require rather com-staging of the valves that control the fuel-flow to eachThus, GE proposed an alternative design involving a sim-single-annular configuration. A series of early tests wereon this single-annular configuration in order to assess itsfor meeting the E3 emissions goals.13

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The prime resutls of those tests are shown in figure I-34.Plotted are the emissions levels for carbon monoxide (CO), hy-drocarbons (HC), and oxides of nitrogen (NO,). The left barin each pair shows single-annular combustor results from testsrecently run as part of the E3 project. The right bar in eachpair shows estimates for the double-annular combustor: Thesepredictions are based on the earlier test results from the Ex-perimental Clean Combustor program. These results indicate thatthe single-annular combustor does not meet the E3 goals for COand NO, emissions. Also, these differences might be increasedwith future broad-specification fuels. Thus, NASA has concludedthat the more-complex double-annular combustor is needed to meetthe stringent E3 emissions goals. All e fforts are now directedat the double-annular configuration so that the low emissionslevels that might be required in future engines can be demon-strated.Another example of the component technology efforts completed todate involves Pratt & Whitney's single-stage turbine. Earlyemphasis was on aerodynamic tests in an uncooied turbine-rig sothat the best airfoil designs could be selected. Two combin-ations of vanes and blades were tested; figure I-35 shows thetwo blades. Results for the uncooled-rig tests are shown infigure I-36. Plotted are the measured efficiency levels asfunctions of airfoil span for the two airfoil combinations.Both designs achieved good efficiencies, but the 43-percent-reaction airfoils produced higher efficiencies. This reactionvalue was thus selected for the E3 turbine design.BenefitsThe fuel savings that have been predicted for the E3 enginesare shown in figure I-37. These are the projections made forboth engine designs by three major aircraft manufacturers: Boe-ing, Douglas, and Lockheed. Using the predicted SFC reductionsof 14 to 15 percent, the aircraft companies predicted fuelusages for their own versions of advanced aircraft for variousflight missions, and then they compared these to the predictedfuel usages for similar aircraft with current engines. Plottedare the differences in the fuel-usage estimates. (Note the or-dinate of the plotsavings for E3 starts at 10 percent.) The projected fuelranges from 13 to 22 percent. The fuel savingsincrease with longer flight distances since the prime fuel sav-ings is achieved in the cruise portions of each flight mission.However, even the shorter-range flights with relatively shortcruise durations show significant fuel savings.Projections were also made for the direct operating costs of thesame aircraft-engine combinations, and a plot of these resultsis shown in figure I-38. As might be expected, the projectedtrends for DOC are similar to those for fuel savings since fuel

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about 40 percent to the direct operatingAlso important in these DOC reductions was the emphasisboth General Electric and Pratt & Whitney placed on minimizingcosts. For example, the total numbers of airfoilsin their engines were drastically reduced from the numbers inhe reference engines.the current status of the E3 project, these benefits ap-

be realistic. Present E3 designs indicate that ailare achievable, with the possible exception of the NO,emissions goal. The advanced technologies required to meet theare scheduled to be demonstrated by 1983.

third ACEE propulsion effort, the Advanced Turboprop proj-ect, is directed at establish ing the feasibility of radicallyimproving propeller-driven propulsion systems to the point wherethey can be effectively applied to future commercial air trans-Advanced turboprop propulsion systems promise extremelylarge fuel savings. And, with an aggressive technology develop-program - demonstrating the advanced technology in the late980's - they would be available for the commercial air trans-of the 1990's.aircraft are not new, of course. Turbopropsused today in various military, business, and commuter air-and at one time were used in medium-range commercial airFigure I-39 shows the turboprop-powered Lockheedwhich was introduced into medium-range commercial ser-in the late 1950's. It was a major improvement over theeciprocating-engine, propeller-driven aircraft then in com-

such as the Boeing 707 shown in figure(the first domestic commercial jet), were also introducedn the late 1950's. Such jet-powered aircraft offered bettercomfort since they flew above the weather and had lessand vibration in the cabin than the Electra. They wereThis, a long with their larger size, providedr productivity and convenience. And, they had newer en-h lower maintenance costs. The 707, and others liket, swept the market. Their major disadvantage was the higherconsumption of the turbojet engines. But, in an era offuel - around 5 to 10 cents a gallon at that time -price of fuel was not a critical factor in propulsion-systemI-41 illustrates this last point. Shown is the fuel con-for various propulsion systems - expressed in gallonsmile to account for speed differences. These values are

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plotted relative to the Electra engine fuel consumption. Alsoplotted are the dates of introduction into service (or possibleintroduction). Going from the Mach 0.6 turboprop-poweredElectra to the Mach 0.8 turbojet-powered 707 resulted in a 50percent increase in fuel consumption.Since the first commercial turbojet was put into service, a num-ber of improvements have increased the efficiency of aircraftgas-turbine engines. Most significantly, the high-bypass-ratioturbofan engine has been a major factor in reducing fuel con-sumption. The Energy Efficient Engine promises another 15 per-cent reduction in fuel consumption, primarily througn betterthermal efficiency. But the curve does appear to be flatteningout, an indication of a maturing technology.Making the next significant step in fuel savings will necessi-tate a major increase in propulsive efficiency. Sucn an in-crease, about 15 percent at Mach 0.8, is offered by the advancedturboprop. Even larger savings would be possible at lowerspeeds. This propulsion-system concept will be considerablydifferent, however, from the older, more conventional, Mach 0.6Electra system. Figure I-42 illustrates this difference. Theadvanced propeller concept is compared with the Electra propel-ler, to the same scale. The shape, size, and number of bladesare all radically different.The major features of the advanced propeller concept are shownin figure I-43.lar'ge, The advanced propeller would be driven by amodern turboshaft engine and a gearbox. The blades ofthe propeller would be very thin and highly swept to minimizeboth compressibility losses and propeller noise during high-speed cruise. An area-ruled spinner and an integrated nacelleshape would also be used to minimize compressibility iosses inthe propeller-blade hub region. A high power loading, and hencea relatively small propeller diameter, would be achieved byusing 8 or 10 blades. A final feature worthy of note is thatmodern propeller blade fabrication techniques would be used tomake the thin, highly swept, twisted blades.The basic reason for the advanced turboprop being such an at-tractive concept is its potential for high propulsive efficiencyin the Mach 0.7 to Mach 0.8 speed range, as shown in figureI-44. Older model turboprops had relatively thick, unswept pro-peller blades and experienced rapid increases in compressibilitylosses above Mach 0.6. Current high-bypass-ratio turbofansexhibit their highest propulsive efficiency (about 65 percent)at cruise speeds somewhat above Mach 0.8. The new advancedturboprops are estimated to be about 20 percent more efficientthan high-bypass-ratio turbofans at Mach 0.8. At lower cruisespeeds, the efficiency advantage of the advanced turboprop iseven larger. This high propulsive efficiency of the advancedturboprop makes it an attractive powerplant for many applica-tions.

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A number of aircraft studies have been conducted to quantify thebenefits promised by the advanced turboprop. Figure I-45 showssome of the passenger aircraft studied to date. Two aircraftare shown: a medium-range wide-body transport with four turbo-props mounted on the wing, and a shorter-range narrow-bodytransport with two turboprops mounted at the rear of the fuse-lage. Cargo airplanes and military patrol aircraft could alsobenefit from using advanced turboprops. In every study, theairplanes powered by advanced turboprops used much less fuelthan competing turbofan-power airplanes.

studies are summarized in figure I-46, where the trend infuel savings with aircraft design range is shown. For short-haul aircraft, where takeoff and descent dominate the fuel frac-tion, the turboprop fuel savings can be as high as 30 percent.For medium-range aircraft, fuel savings are 15 to 20 percent.For very long-range aircraft, where cruise dominates the fuelfraction, turboprop fuel savings are 17 to 30 percent. Thesefuel savings for the turboprop are relative to a turbofan-powered aircraft with the same level of component technology.Thus, if an E3 turbofan will achieve a 15 percent fuel savingsover a conventional turbofan in a new medium-range transport, anew turboprop with an E 3 level of component technology wouldachieve a 30 to 35 percent fuei savings. It is this very largefuel savings potential that prompted NASA to include the Ad-vanced Turboprop project in its ACEE program.The objective of this project is to develop the technology forefficient, reliable, and acceptable operation of advancedturboprop-powered aircraft at cruise speeds between Macn 0.7 andMach 0.8. The turboprop goals of minimums of 15 percent fuelsavings and 5 percent DOC improvement are relative to turbofansat equivalent levels of technology. The third turboprop goal isa cabin comfort level (noise and vibration) the same as that inmodern turbofan-powered aircraft.To achieve these goals, Lewis began a phased Advanced Turbopropproject in 1978. The major elements of phase I (an enablingtechnology and research phase) are shown in figure I-47. Thefirst element - the propeller-nacelle - is concerned with pro-peller aerodynamics, acoustics, and structures. The second ele-ment addresses the cabin environment. Since the fuselage may bein the direct noise field of the propeller, the noise generatedby the propeller must be attenuated by the cabin wall in orderto provide a low-noise cabin environment. Also, the turbopropengines must be mounted in such a way that cabin vibration islow. The tnird major element of phase I, installation aerody-namics, is concerned with an accelerated, swirling propellerslipstream flowing over the wing. Here, there is the challengeof integrating propeller design with wing design to achieve thebest combination of engine efficiency and aircraft lift-dragratio. Also, airplanes powered by advanced turboprop engines

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must be configured to have adequate stability and control. Thefourth element involves the mechanical components of an advancedturboprop propulsion system: the engine drive, the gearbox, andthe advanced propeller. These components must be designed andpackaged in such a way that maintenance and reliability will bemuch improved over that experienced by the first generation ofturboprop-powered aircraft. Since these four elements are sostrongly interrelated, aircraft trade-off studies must be per-formed to obtain the match that will best achieve the goals oflow fuel consumption, low operating cost, and passenger accep-tance.Propeller-NacelleThe advanced propeller technology program in phase I is outlinedin table I-6. The plan is to test a matrix of eight 2-foot-diameter wind tunnel models. The range of the propeller designparameters is indicated in the table. Eight or 10 blades areused to keep the propeller diameter relatively small. Sweep atthe blade tip as high as 60 will be investigated to determineits effect on propeller-generated noise at cruise. The tip-speed range results in transonic relative tip Mach numbersduring Mach 0.8 cruise. The high-cruise power loading also re-sults in relatively small propeller diameters. Besides the ex-perimental work on propellers in phase I, an extensive analyti-cal effort is under way to upgrade prediction and design capa-bilities.Three blades have been tested and are shown in figure I-48.Each blade is extremely thin over much of its span for minimumcompressibility losses during high-speed operation. Also, theblade aspect ratio is relatively low as a result of designingfor a relatively small propeller diameter (i.e., high powerloading). The blade on the left in the figure is straight, withOo of tip sweep. The middle blade has 300 of tip sweep.The blade on the right, with 450 tip sweep, was designed withboth refined aerodynamic and acoustic design methodology. In-deed, acoustics dictated the rather unique blade planform shape,which has a longer chord and more sweep than the other twoblades. This high-sweep propeller model is shown in figure I-49installed in the Lewis 8- by 6-Foot Wind Tunnel. All three pro-peller models were tested in this wind tunnel, and all weredriven by the strut-mounted air-turbine propeller test rig shownin the background.Efficiency. - The measured efficiencies of all three propellersare shown in figure I-50 as a function of Mach number. At Mach0.8 conditions, measured efficiency improved as the tip sweep ofthe models increased. This confirms that blade sweep does re-duce compressibility losses. Also, efficiency improved as Machnumber was reduced.

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fuselage wall and the cabin interior trim. These two conditionsare interrelated, as shown in figure I-55.The maximum-cabin-noise goal is 75 decibels on the A scale(dBA). For a typical propeller design, sound pressure level inthe cabin would be 90 decibels. With current conventional cabintreatment, exterior noise could be as high as 110 decibels.With improved fuselage design and cabin treatment, exteriornoise of 140 decibels or more could be tolerated according toinitial studies conducted by the Langley Research Center. How-ever, the price that must be paid for the higher noise attenu-ation is an increase in fuselage weight.Influencing cabin noise, of course, is the propeller noiselevel. At the moment, there is uncertainty in predicting noiselevel and also uncertainty in interpreting wind tunnel noisemeasurements. The absolute noise level of the three modelstested thus far is somewhere in the shaded area with a possiblelower bound of approximately 140 decibels. With improved pro-peller design, it may be possible to reduce the noise level toabout 130 decibels. Since the two bars overlap, a number ofsolutions are possible for achieving the desired maximum noiselevel in the cabin. Thus, it will be necessary to conduct air-craft trade-off studies. Engine location is, of course, anothervariable in such studies. The optimizations are quite differ-ent, for example, when the turboprops are mounted at the rear ofthe fuselage rather than on the wing.Installation AerodynamicsThe third major element of the Advanced Turboprop project isinstallation aerodynamics. So far only the question of drag hasbeen addressed. Drag becomes of concern with the advanced tur-boprop because, at the Mach 0.7 to 0.8 cruise speeds of inter-est, large compressibility losses could be encountered on thatarea of the wing washed by the accelerated slipstream of thepropeller.Preliminary results have been obtained with a propeller slip-stream simulator (fig. I-56). This figure shows a wing-bodyaircraft model installed in an Ames wind tunnel. There were nomodel turboprops directly on the aircraft model; instead, theejector device directed a simulated propeller slipstream overthe wing. By using this device, slipstream Mach number andslipstream swirl angle were varied independently. It is thesetwo parameters that are peculiar to a turboprop installation.The results are shown in figure I-57 for two cruise Mach num-bers. The aircraft drag at a lift coefficient typical of cruiseis the zero reference point. When the simulator was turned on,generating a Oo- swirl slipstream, the increase in drag was

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about 2.5 percent. This level was maintained to swirl angles of70. At higher swirl angles, drag decreased as the wing re-covered some of the swirl and actually became negative between7O and 110. Since the amount of swirl of the slipstreamincreases as power loading is increased, propeller design se-lection must consider these slipstream effects. Wing designmust also be optimized for the recovery of swirl. Such integra-tion may yield an installation drag for the turboprop equal to -or perhaps even lower than - that of the turbofan.Turboprop installation drag will be accurately assessed in thenear future by testing an aircraft model with a 2-foot-diameterpropeller, as shown in the artist's sketch of figure I-58. Datafrom this test will guide the integration of propeller designwith wing design.

Mechanical ComponentsThe fourth major element in the Advanced Turboprop project in-volves the mechanical components. This element was of partic-ular concern at the beginning of the project because of the highmaintenance costs for the turboprop propulsion system on theElectra and the general feeling that turbofan maintenance costswould always be substantially lower. For this reason, the LewisResearch Center placed a contract with Detroit Diesel A llison(with Hamilton Standard as a subcontractor) to evaluate theElectra system. Results indicated that the 1970-era turbopropmaintenance costs were indeed higher than those for the JT8Dturbofan (a representative turbofan), but the difference wasmainly in the core engine. The core engine of the turboprop wasexpensive to maintain largely because it was an older technologycore originally designed for military application. It wastherefore concluded that overall advanced-turboprop maintenancecosts can be competitive with those of an advanced turbofan.Core maintenance costs, based on equal levels of technology,should be about equal. Preliminary conceptual designs were madeto evaluate how turboprop maintenance costs couid be reduced.The advanced propeller and gearbox can be greatly improved byperforming on-condition maintenance instead of scheduledoverhauls, by using modular construction, and by emphasizingsimplicity and reliability in design.Future PhasesResults of the subscale testing part of phase I has been verypromising. Progress has been achieved in developing design andprediction methodologies. This phase is an enabling technologyeffort that will take about 3 years to accomplish, with the in-tent being to establish the feasibility of concepts and a database for future efforts. As a logical next step, NASA is plan-

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ning large-scale component verification tests that will lead toflight verification (fig. I-59). To be evaluated are propel-ler fabrication, propeller flutter, scale effects relative topropeller-generated noise, and advanced fuselage-noise-attenuation concepts. Operational effects for a full range offlight conditions would be an important feature in demonstratingtechnological readiness. Thus, at the conclusion of the Ad-vanced Turboprop project, technology would be available for usein designing very efficient turboprop-powered commercial air-craft and in establishing commercial acceptance of this advancedsystem.CONCLUDING REMARKSThe technology readiness dates for the three Aircraft Energy Ef-ficiency propulsion projects are shown in figure I-60. By thedates shown, the technology required for beginning commercialdevelopment will be demonstrated. Such developments would beexpected to yield the fuel savings indicated in the figure.Figure I-61 indicates the potential benefits of the three proj-ects in terms of both direct operating cost and fuel savings.DOC savings are, of course, important since future propulsionsystems must be economically attractive to the airlines as we1.Las fuel efficient.There could be about a 5 percent fuel savings and a 3 percentDOC savings in derivative engines that use Engine Component Im-provement (ECI) technology. With a technology readiness data of1980 to 1982, this project will yield results that are extremelyapplicable to the near-term needs of the airlines. Energy Ef-ficient Engine (E3) benefits are about a 15 to 20 percent fuelsavings and a consequent 5 to 10 percent DOC savings. Thesebenefits could be realized by the late 1980's in new enginesthat use E3 technology to be demonstrated by 1983. Advancedturboprops are expected to provide about 30 to 35 percent fuelsavings over current engines, if E3 technology is used for theturboshaft engine, with a possible 10 to 15 percent DOC sav-ings. These major gains could begin to be achieved in the1990's, if the technology is successfully demonstrated in thelate 1980's. In a future where fuel prices (and possibly evenfuel allocations) will influence aircraft design and selection,advanced turboprops are expected to play an important role.These three projects differ not only in their technology readi-ness dates but also in the level and type of technology thateach is addressing. In going from the EC1 project to the E3project to the Advanced Turboprop project, the emphasis shiftsfrom the near term to the far term, with progressively largerbenefits, but also with more unknown technology, with higherdegrees of risk, with more difficulties in integration with theaircraft, and with ultimate applications being clouded as usual

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by market and economic uncertainties. The ACEE effort, however,represents a comprehensive approach to resolving these diffi-culties and unknowns and offers a number of technology options.And, if the ACEE technology advances prove as attractive as ex-pected, this program should have a major impact on future com-mercial air transportation systems.REFERENCES1. Fuel Cost and Consumption, Monthly Reports. Civil AeronauticsBoard.2. Trends in Airline Cost Elements. Second ed., Civil Aero-nautics Board, July 1977 plus Supplement, Jan. 3, 1979; andTrends in Airline Unit Costs. Fourth ed., Civil Aero-nautics Board, Feb. 1979.3. Aircraft Fuel Conservation Technology - Task Force Report.NASA-Office of Aeronautics and Space Technology, Sept. 10,1975.4. Fasching, W. A.: CF6 Jet Engine Performance Improvement Pro-gram: Task I - Feasibility Analysis. (Ri'9AEG295, GeneralElectric Co.: NASA Contract NAS3-20629.) NASA CR-159458,1979.5. Gaffin, W. 0.; and Webb, D. E.: JT8D and JT9D Jet Engine Per-formance Improvement Program: Task I - Feasibility Anal-

ysis. (PWA-5518-38, Pratt & Whitney Aircraft Group; NASAContract NAS3-20630.) NASA CR-159449, 1979.

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.PERFORMANCE IMPROVEMENT CONCEPTSPRATT & WHITNEY ENGINES

CONCEPTS

P & WA JTSD ENGINEHPT OUTER AIR SEALNEW HPT BLADEREVERSER STANG FAIRINGTRENCHED ABRADABLE HPCP & WA JT9D ENGINEHPT ACTIVE CLEARANCE CONTROLNEWFANTECHNOLOGY

HPT VANE THERMAL COATINGHPT ABRADABLE OUTER SEAL

% SKREDUCTION

0. 6.9.5.9

.71. 3::IL

COMPLETIONDATE

SEP 1978JUL 1980SEP 1980DEC 1980

NOV 1978JUN 1979FEB 1981OCT 1981cs-79-1534

Table I-l

PERFORMANCE IMPROVEMENT CONCEPTSGENERAL ELECTRIC ENGINE

CONCEPTS

GE CF6 ENGINESHORT CORE NOZZLENEW FRONT MOUNTIMPROVED FANHPT AERO REFINEMENTHPT ROUNDNESS CONTROLREDUCED HPC BLEEDLPT ACTIVE CLEARANCE CONTROLHPTACTIVE CLEARANCE CONTROL

% SFC COMPLETIONREDUCTION DATE

0. 9 FEB 1979.3 MAR 1979

2. 0 MAY 19791.3 JUL 1979.4 MAY 1980.7 SEP 1980.3 OCT 1980.6 AUG 1981

cs-79-1535

Table I-2

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PROPULSION SYSTEM LOADS ANALYSIS RESULTS

STEADY STATEAREA AFFECTED % OF TOTALSFC LOSSALL STAGES 87HPT STAGES 8FAN STAGE 5

DYNAMICNO SIGNIFICANT CHANGE FROM STEADY STATE

cs-79-2089

Table I-3

CYCLE CHARACTERISTICS OFGENERAL ELECTRICS ENGINES

AT MAX CRUISE CONDlTlONS

EXHAUST CONFIGURATIONBYPASS RATIOFAN PRESS. RATlOCOMPRESSOR PRESS. RATIOOVERALL PRESS. RATIOTURBINE TEMP, OFHOT-DAY TAKEOFFMAX CRUISETAKEOFF THRUST, SLS-lbINSTALLED SFC, lbmlhrllbfcs-79-2065

EEE CF6-5DCMIXED SEPARATE6. 9 4.31. 61 1. 7222.6 12.936.1 32. 0

2450 24002170 208036 500 50 2500.572 0.666L=-14.%A

Table I-4

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CYCLE CHARACTERISTICS OFPRATT 8, WHITNEY'S ENGINES

AT MAX CRUISE CONDITIONS

EEE JT9D-7AEXHAUST CONFIGURATIONBYPASS RATIOFAN PRESS. RATIOCOMPRESSOR PRESS. RATIOOVERALL PRESS. RATIOTURBINE TEMP, OFHOT-DAY TAKEOFFMAX CRUISETAKEOFF THRUST, SLS-lbINSTALLED SFC, lbmlhrllbf

CS-79-2067

MIXED SEPARATE6. 6 5.11. 71 1.5813. 9 10. 037. 4 25.42495 23002200 2000

41 115 46 3000.576 0.677IA = -14.9% 3

Table I-5

ADVANCED PROPELLER TECHNOLOGY PROGRAMPHASE I

DESIGN MATRIX OF EIGHT 2-ft DIAM WIND TUNNEL MODELSNO. OF BLADES 8 OR 10SWEEP AT BLADE TIP, deg 0 TO 60TIP SPEED, ftlsec 600 TO 800(RELATIVE TIP MACH NO. 1 (1.0 TO 1.15)POWER LOADING, hplft2 26 TO 37.5(RELATIVE DIAMJ (1. 2 TO 1.0)

USE DATA TO UPGRADE AERO, ACOUSTIC, & STRUCTURAL ANALYSIS &PREDICTION PROGRAMS CS-79-2155

Table I-6

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FUEL PRICE,algal

U.S. AIRLINEJET FUELMONTHLY AVERAGES

50rPRICE

-4't

- INTERNATIONAL TRUNKS10 --- DOMESTIC TRUNKS

- 1977 I978YEAR B-79-2295Figure I-l

DIRECT PERATINGOST LEMENTSDOMESTIC OPERATIONS OFTHE DOMESTIC TRUNKS

-----FUEL'- -MAINTENANCE {DIRECT7- AND INDIRECT] /---CREW / /

.- -.OTHER iDEPRECIATION. .0'6-- HENlALS. INSlJRANCEt ././,5- //CENTS PER /AVAILABLE 4-TON-MILE

7-l-

I;69 1970 -1971i 1972 1973 1974 1975 1976 1977 :978YEAR ENDFigure I-2

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DIRECTOPERATING OSTSB727-200; DOMESTIC OPERATION

FUELCREWMAINTOTHER

Figure I-3

ACEE PROPULSION PROJECTS

=-ENGINE COMPONENT _= \=IMPROVEMENT 1 X= ENERGY EFFICIENTENGINE=

ADVANCED TURBOPROP

Figure I-4

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- -

ENGINE COMPONENT IMPROVEMENT PROJECT

PERFORMANCE IMPROVEMENT ENGINE DlAGNOSTlCSGOAL OF % FUEL SAVINGS cs-79-2099Figure I-5

PERFORMANCE IMPROVEMENT

P & WA JT8D P & WA JT9DDEVELOP TECHNOLOGY FOR COMPONENTS TO REDUCEFUEL CONSUMPTlON IN NEW PRODUCTION ORRETROFIT OF CURRENT ENGINES BY 1980-1982

Figure I-6

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COMPONENT AREASHIGH PRESS. COMPRESSOR (HPCh :FAN-/ :

7 LOW PRESS. TURBINE (LPT)\\\EXHAUST NOZZLE

CS-79-2096

FigureI-

JT8D HPT OUTER AIR SEALwURRENT

ADDED KNIFE-EDGE SEALADDED HONEYCOMB SEALCOOLING AIR DISCHARGE r-l.RELOCATEDCRUISE SFC REDUCTION, 0.6%DEMONSTRATED I I

IMPROVEDc-79-2144

Figure I-8

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JT8DHPT OUTER IR SEALENGINE EST

CRUISESFCREDUCTION, -70020 40 60 80% MAX CRUISE THRUST tS 7g ,;;- -

Figure I-9

ACTIVECLEARANCEONTROL

CRUISE CLEARANCETAKEOFF CLEARANCESHROUD & CASE

NO AIR COOLING AIR CS-79-208b

Figure I-10

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. .,..---.--.--w

JTSD HPT ACTIVE CLEARANCE CONTROL

CURRENTMODIFIED SEAL SUPPORT RINGINCREASED CASE COOLINGAIR FLOWSHORTER IMPINGEMENTDISTANCECRUISE SFC REDUCTION, 0.6R. DEMONSTRATED IMPROVED

Figure I-11

JT9DHPTACTIVE LEARANCEONTROLTESTDATA1.0 -

CRUISE (-)ooOo c\SFCREDUCTION, 5 NORMAL% CRUISE0 I60 70 80 90 100

cs-79-2086 % MAX CRUISE THRUST

Figure I-12

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CF6FAN

IMPROVED AIRFOILSREARWARD SHROUDFAN CASE STIFFFNFRCRUISE SFC REDUCTION, % DEMONSTRATED CURRENT IMPROVED

cs-79-2097

Figure I-13

CF6ENGINE ESTCOMPARISON--o-- CURRENT FAN ENGINES

I --- IMPROVED FAN ENGINES

SFC

I I I I35 ooo 40 000 45 OaJ 50 cm0SEA LEVEL THRUST. lbcs-7g-2083

Figure I-14

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CONTRIBUTORSO ENGINE ERFORMANCEETERIORATION

f CLEARANCE INCREASES

Figure 1-15

EXAMPLESFENGINE ERFORMANCEIjETERlORATlON

F.O. D. EROSION,fjHROUD

--WARPAGE

CLEARANCE INCREASE THERMAL DISTORTION cs-7g-zog4Figure I-16

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CF6MODULE EPLACEMENT/REFURBISHMENTESTRESULTSCRUISE CONDlTlONS

REPLACE VANES/SEALSmm.-% ASFC

NEW MODULE0.4% ASFC

FAN (2)CLEAN BLADESlRECONTOUR

LEADING EDGE0.3% ASFC

Figure I-17

JTSDMODULAR ERFORMANCEETERIORATIONESULTS

CRUISE SFCDETERIORATION,%

cs-79-2091

USED PARTS DATA & PRE-REPAIR TESTS0 CLEARANCE INCREASEm EROSION/AIRFOILROUGHNESSlX$l THERMAL DISTORTION

1OOOTH LIGHT

FAN Lt% Hlk HPT CPT(ENGINE MODULE)Figure I-18

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JT9D PERFORMANCEETERIORATIONBY DAMAGEMECHANISMCRUISE CONDITIONS

4r EROSION(FAN, LPC, HPC)y

CRUISE SFCDETERIORATION,%

cs-79-2088

3- /'

> CLEARANCEINCREASE1 1 ) jj (ALL)0 1000 2DcQ 3000

FLIGHT CYCLES

FigureI-

JTSDEXTERNALPPLIEDLOADS ND REACTIONS

THRUST

NACELLE AERODYNAMIC 8"PRESS. DISTRIBUTION >'

LOADS

*NACELLE AERODYNAMIC LOADS*INERTIA LOADS (s--".. J:

FigureI-

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JT9D/747 PROPULSIONYSTEM TRUCTURALODEL

INLET

11ODDSTATIC FREEDOMS4 000 ELEMENTSREVERSER, COWLING, ETCINCLUDED IN ANALYSIS

135-79-2092

Figure I-21

JTSDENGINEN PRAll &WHITNEYX-RAY NGINE EST ACILITY

FigureI-

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AVERAGE NGINEPERFORMANCEDETERIORATIONRENDSJT9D & CF6

CRUISE SFCDETERIORATION,%

cs-79-2085

4 r3 TYPICALOVERHAUL2 RECOVERYCURRENT1 Lcz i UNRESTOREDP RFORMANCE

0 1000 2000 3000 4MKlUSAGE, CYCLES

Figure I-23

ENERGY EFFICIENT ENGINE PROJECTSUMMAl?cSCHEDULE

IcyPROPULSION SYSTEMDEFINITION & DESIGN

COMPONENT TECHNOLOGIES

SYSTEMS INTEGRATIONTECHNOLOGIES

cs-79-2071 GENERAL ELECTRICPRAlT & WHITNEY

Figure I-24

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..-..-_.. .-.-.-... ,....._._._. -

ENERGY EFFICIENT ENGINE CONFIGURATIONGENERAL LECTRIC

Figurer-25

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ENERGY FFICIENT NGINE GENERAL LECTRIC ONFIGURATIONCORE COMPONENT TECHNOLOGIES

/ i/ I/ I/ ITEN-STAGE H.P. COMPRESSOR TWO-ZONE COMBUSTOR23~1 PRESS. RATIO LOW EMISSIONSHIGH AIRFOIL LOADINGS ADVANCED DIFFUSERACTIVE CLEARANCE CONTROL SEGMENTED LINERADVANCED MATERIALS DIGITAL-CONTROL STAGlNG(35-79-2187

TWO-STAGE H.P. TURBINEVERY HIGH EFFICIENCYLOW LEAKAGEIMPROVED COOLINGADVANCED MATERIALSACTIVE CLEARANCE CONTROL

Figure I-26

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ENERGY FFICIENT NGINE GENERAL LECTRIC ONFIGURATIONLOW-SPOOL COMPONENT TECHNOLOGIES

MIXERHIGH EFFECTIVENESSLOW PRESS. LOSS

/

:-.; :- I ., :I_. i

.:. __.. -.: .. . -

iA-SINGLE-STAGE FAN l/4-STAGE ISLAND BOOSTER FIVE-STAGE L P. TURBINELOW TIP SPEED AUTOMATlC CORE MATCHING LOW NOISE CONFIGURATIONINTEGRAL STATORlFRAME REDUCED CORE FOD ACTIVE CLEARANCE CONTROL cs-79-Z t36SlEELlKEVlAR CONTAINMENT

Figure I-27

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ENERGY EFFICIENT ENGINE CONFIGURATIONPRATT & WHITNEY

Figure I-78

ENERGY FFICIENT NGINE PRATT& WHITNEY ONFIGURATIONCORE COMPONENT TECHNOLOGIES

/ I \TEN-STAGE H. P. COMPRESSOR TWO-ZONE COMBUSTOR SINGLE-STAGE H. P. TURBINESUPERCRITKAL AIRFOILS LOW EMISSIONS HIGH EFFKIENCYTRENCHED CASES CARBURETOR FUEL NOZZLES TRANSONIC FLOWACTlVE CLEARANCE CONTROL ADVANCED LINER IMPROVED COOLINGADVANCED MATERIALS DIGITAL-CONTROL STAGING ADVANCED MATERIALS

cs-79-2189 ACTlVE CLEARANCE CONTROLFigure I-29

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ENERGY FFICIENT NGINE PRATT& WHITNEY ONFIGURATIONLOW-SPOOL COMPONENT TECHNOLOGIES

\MIXER

/I \ \HIGH EFFECTIVENESSLOW PRESS, LOSS

SINGLE-STAGE FAN 4-STAGE L P. COMPRESSOR 4-STAGE L P. TURBINESHROUDLESS BLADES HIGH LOADING COUNTER-ROTATINGSTRUCTURAL VANES CANTED AIRFOILS LOW LEAKAGEAllKEVlAR CONTAINMENT ACTIVE CLEARANCE CONTROLCS-79-2lA8

Figure 1-N

ENERGY FFICIENT NGINE GENERAL LECTRICFAN-FRAME ACOUSTIC CONFIGURATIONS

CONVENTIONALCONFlGURATlON

0 0 0STRUTSVANkBLADE RATIO -2

Figure I-31

E3CONFIGURATION

-1.1 CC.-79-07-

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ENERGY FFICIENT NGINE GENERAL LECTRICCOMBUSTOR EMISSION CHARACTERISTICS; 4% IDLE THRUST

m SINGLE-ANNULAR COMBUSTOR (TEST)5 r m DOUBLE-ANNULAR COMBUSTOR (PREDICTED)4

POUNDS PER1000 Ibs 3 --THRUSTHOURS PER 2CYCLE

GOAL.- -----

1n co HC

Figure I-34 cs-79-2073

ENERGY FFICIENT NGINE PRATT& WHITNEYH.P. TURBINE UNCOOLED RIG HARDWARE

3% REACTlON BLADE 43% REACTlON BLADEFigure I-35

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ENERGY FFICIENT NGINE PRATT& WHITNEYH.P. TURBINE UNCOOLED-RIG TEST RESULTS

EFF,%8684

% EFFREACTION (UNCOOLED)0 35 90. 10 43 90. 6

8210 10 20 30 40 50 60 70 80 90 100AIRFOIL SPAN, %cs-79-2069

FigureI-

POTENTIALUELSAVINGS F ENERGY FFICIENT NGINESRELATIVE TO THE SAME AIRCRAFT WITH SCALED JT9D-7A OR CF6-50C

25 r

0 BOEINGCl DOUGLASA LOCKHEED

BLOCK 20FUELSAVINGS,% 15

10 I0 1000 2000 3000 4000 5000 6000 7000FLIGHT DISTANCE, n mi cs-?g-?o7FFigureI-

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POTENTIAL OC MPROVEMENTSF ENERGYEFFICIENT NGINESRELATlVE TO THE SAME AIRCRAFT WITH SCALED JT9D-7A OR CF6-5DC

15r0 BOEING0 DOUGLASn LOCKHEED

DOCIMPROVEMENT,%

I I Ics-79-207. 0 1000 2000 3000 4000 5000 6000 7000FLIGHT DISTANCE, n mi

Figure I-38

LOCKHEEDLECTRAIRCRAFT

Figure I-39

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BOEING 07 AIRCRAFT

Figure I-40

PROPULSIONYSTEM ELATIVE UELCONSUMPTION1.6rI vFlRST COMMERCIAL TURBOJETS1.4

REL4TlVE l-2FUELCONSUMPTlON, l..

gal/mile.8

fl

vLOW BYPASS RATIO TURBOFANS

.HIGH BYPASS RATIO TURBOFANS\-.. ENERGY EFF;"::",, ENGlNE

.' _,". ,.:, ::,.mvmcm TURBOPROP. L73 ENGIN (MACH 0.8)1960 I970 1980 1990 2000

?i-7y?-Z2U YEAR INTRODUCED INTO SERVICEFigure I-41

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SIZE/CONFIGURATION COMPARISON

ADVANCEDROPELLER ELECTRAROPELLERcs-79-2149figure I-42

ADVANCEDURBOPROPROPULSIONYSTEM/) .c-THlN SWEPT BIADES

L LARGE, MODERNLOADING &ODERN BlADE TURBOSHAFT(8 OR 10 BLADES) FABRICATION ENGINE & GEARBO:

CS-79-2165

Figure I-43

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INSTALLED ROPULSIVEFFICIENCYT CRUISEHIGH BYPASS100

INSTALLED 90PROPULSIVE 80EFF,Lw 70/60511

cs-79-2153 -15CRUISE MACH NO.

Figure I-44

ADVANCED TURBOPROP PASSENGER AIRCRAFT

Figure1-45

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TRENDOF POTENTIALUELSAVINGS ORADVANCEDTURBOPROP-POWEREDIRCRAFTRELATIVE TO TURBOFAN-POWERED AIRCRAFT WITH

SAME LEVEL OF CORE TECHNOLOGY

% 10 t- TAKEOFF & CRUISE FUELDESCENT DOMINATFO IIOMINATFD- -_... - -....... ..-- - - . . . - -I

0 2000 4000 6000DESIGN RANGE, n mi _._~.,_ _.

Figure I-46

ADVANCEDURBOPROP ROJECT

nPHASE I - MAJOR ELEMENTS

AIRCRAFT TRADEOFFSb MECHANICALCOMPONENTS

0 ENGINE0 GEARBOX0 PROPELLER

9 GOALSINSTALLATION 0 LOW FUEL CONSUMPTION0 LOW OPERATING COST CABIN ENVIRONMENTAERODYNAMICS 0 PASSENGER ACCEPTANCE . NOISE0 DRAG . VIBRATIONl STABILITY CONTROL cs-79-2162

Figure I-47

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PROPELLER MODEL EFFICIENCYLEWIS 8 x 6 WlND TUNNEL

85r POWER LOADING37.5 hp/ft* AT M = 0.8

MEASUREDPROPELLEREFF,%

cs-79-2163

80

75

70.70 .75 .80 .85MACH NO.

Figure I-50

PROPELLERERFORMANCET MACH0.8MODEL WITH 45 TIP SWEEP; LEWIS 8 x 6 WIND TUNNEL

81r37.5 POWER LOADING. hplft*

76i650 700 740 a00

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NOISECOMPARISONF ADVANCEDPROPELLER ODELSMO = D.8 DESIGN CONDITIONS; LEWIS 8 x 6 WIND TUNNEL

cs-7gd.156 6c /-- /- uO R--//--

ON-, -.REDUCTION IN 2 %-.PEAK SPL OFBLADE PASSING 0FREQUENCY,dB 4

Figure I-52

Figure I-53

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POWERED-PROPELLER SEMI-SPAN AIRCRAFT MODEL

Figure I-58

ADVANCEDURBOPROP

FUTURE PHASESLARGE-SCALE COMPONENTS - DEVELOPMENT& VERIFICATlON. PROPELLER FABRICATION0 PROPELLER AEROlACOUSTlClSTRUCTURAL SCALING VERlFlCATlON0 FUSELAGE SEGMENTS/CABIN ENVIRONMENT. ADVANCED GEARBOX & PITCH CHANGE SYSTEM

l AIRFRAME INTERACTIONS0 INSTALLED PERFORMANCEl OPERATlONAL EFFECTS0 ADVANCED ENGINE/GEARBOX/PITCH CHANGE SYSTEMPERFORMANCE

Figure I-59 -,-7,-, 1. ,

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ACEEPROPULSION ROJECTSPROJECTED FUEL SAVINGS & TECHNOLOGY READINESS DATES

40

F

. MED RANGEMISSION ADVANCED. SAVINGS RELATIVE TURBOPROP30 TO CURRENT (ATPlENGINES

EYERGY EFFICIENT ENGINE(E Ilo-

ENGINE COMPONENT IMPROVEMENT(EC110 I I I1980 1985 1990 1995

(35-79-2293 TECHNOLOGY READINESS DATE

Figure I-60

POTENTIAL ENEFITS F ACEEPROPULSIONROJECTSSAVINGS RELATIVE TO CURRENT ENGINES

MACH 0.8 MEDIUM-RANGE MISSION4030

MISSION F& iy~~~~~~~~~~~ 1 : ..:G>,,gr .&.: :,. ENERGY E F F ICIE NT E NGINE

10 -a ENGINE COMPONENT IMPROVEMENTI I

0 5 10 15 20DOC SAVINGS, % cs-79-2291

FigureI-

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itemized on the right of figure 11-2, all result in a markedreduction in carbon monoxide and hydrocarbon emissions.Increasing the burning zone residence time will allow time forthe complete consumption of the hydrocarbon fuel and carbon mon-oxide. Reducing the flow velocity is one way to increase theresidence time. Delaying the injection of dilution air will re-sult in a longer primary zone and hence increased residencetime. Delayed mixing also reduces quenching effects caused byrapid cooling of the burning mixture. Increasing the burning-zone equivalence ratio, that is, adding more fuel to the burningzone, results in a higher local temperature, which acceleratesthe combustion reactions. Improvements in fuel atomization anddistribution prevent large pockets of fuel-rich mixtures fromoccurring and make for smaller fuel droplets which can be morerapidly consumed.At high power conditions, typically takeoff and climb, the majorpollutants are oxides of nitrogen and smoke. Carbon monoxide andhydrocarbons are present only in very small amounts, as com-bustion efficiency is virtually 100 percent. Oxides of nitrogenand smoke emissions are formed when residence times are long,when flame temperatures become very high due to the high airpressures and temperatures (which are typical of high-power oper-ation), and where there is poor local fuel distribution, whichcauses very-high-temperature, fuel-rich pockets or zones. Thesepollutants can be reduced by decreasing the combustor residencetime so that they do not have time to form in any significantamount. This can be done by increasing the velocity or by en-hancing mixing, which effectively reduces the length of the burn-ing zone. Decreasing the equivalence ratio, that is, operatingfuel lean, reduces the maximum flame temperature and the rate atwhich these pollutants are formed. Improving fuel atomizationand distribution results in a more uniform mixture of the fueland air.An examination of the emission-reduction techniques reveals aninteresting conflict between those that reduce idle emissions andthose that reduce high-power emissions (fig. 11-3): Witn tneexception of improved fuel atomization and distribution, the re-duction techniques are in opposition. As an example, to reducelow-power emissions, increased residence time is needed; however,to reduce high-power emissions, decreased residence time isneeded. This poses a difficult combustor design dilemma. It isas if two different combustors, one for low power and one forhigh power, are required.

he problems, then, were to determine if these conflictingapproaches could be integrated into a real engine combustor with-out compromising performance and, if so, to determine to whatlevel the resulting pollutants could be reduced. These were theuestions to be answered by our near-term emission-reduction pro-Put another way, the objectives of this program were to61

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investigate new combustor concepts with the potential for signif-icantly lower emission levels and to measure the emissions re-duction by engine test. Contracts were awarded to major aircraftengine manufacturers to devise and investigate new combustor con-cepts. In general, these contracts were for a multiphase pro-gram. The first phase consisted of the screening of a variety ofnew combustor concepts to determine those with the greatestemission-reduction potential. Those concepts so identified wererefined in the second phase, and finally the best or most "engineready" combustor design was tested in an engine to measure theemission reduction obtainable. Table II-2 shows the engine manu-facturers and the engines selected. The engines are arranged inorder of increasing compressor pressure ratio. The EPA engineclass designation is shown in the left column.Near-Term Program ResultsThe results of the programs conducted with the engines in EPAclass T2 are discussed herein. In general, these results arequite similar to those obtained for combustors in the other en-gine classes. Figure II-4 shows the Vorbix combustor used in theJT9D-7 engine. Vorbix is an acronym meaning vortex burning andmixing. A cross section of this combustor (fig. 11-4) shows thattnere has been a drastic departure from conventional combustors.The Vorbix combustor consists of two burning stages arranged inseries: a pilot stage, for low-power emission control, and amain stage, for operation at all engine conditions beyond idle.The main stage is separately fueled; the pilot stage is the ig-nition source for the main stage. The bar graphs below the draw-ing compare the emissions of a production JT9D-7 combustor withthose obtained with the Vorbix combustor tested in JT9D-7 en-gine. Carbon monoxide emissions were reduced by more than one-half; total hydrocarbon emissions were reduced by over a factorof 10; and oxides of nitrogen emissions were reduced by over one-half. A photograph of the Vorbix combustor is shown in figure11-5.Figure II-6 shows the double-annular combustor tested in an ex-perimental CF~-50 engine. This combustor also has two stages,but, here, the pilot and main stages are arranged in parallel,resulting in two annular burning zones. The pilot zone is usedat all operating conditions and is designed to control low-powerpollutants. The main zone is functional at all engine conditionsabove idle and is designed to reduce the high-power pollutants.The bar graphs compare the production CF6-50 combustor emissionswith those of the double-annular combustor. Carbon monoxideemissions were reduced by about 40 percent; total hydrocarbons bya factor of 10; and oxides of nittrogen by about 30 percent. Aphotograph of the double-annular combustor is presented in figure11-7.62

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This then forms the basis for the far-term emission reductionprogram.The objective of our far-term program is to evoive the technologyneeded for the development of combustors with minimum pollutantlevels. The achievement of this objective relies heavily on con-tinuing basic and applied research. The degree of risk and over-all level of complexity associated with the adaptation of ad-vanced techniques is more severe than that of the near-term pro-grams. Fundamental studies are viewed as a requirement to clos-ing the gaps in our understanding of key areas and to bringingthe new technology to a point where a new approach to combustordesign is practical. As discussed previously two techniquesappear particularly attractive: The lean, premixed, prevaporizedand the catalytic combustion techniques. In late 1979 contractswill be awarded to evolve and evaluate both of these combustorconcepts.Before lean, premixed, prevaporized combustors can be used inaircraft engines, additional research is required in severalareas. Shown in figure II-17 is a conceptual drawing of a lean,premixed, prevaporized combustor, which is a staged type ofdesign. The pilot stage has been configured to include features,such as a hot-wall liner to minimize idle pollutants. The mainstage looks much like a flame-tube rig and contains a fuel in-jector, a premixing and prevaporizing section, and a flameholder. To maintain a wide operating range, while burning aslean as possible, controi of the airflow as well as the fuel flowbetween the two stages may be required. To achieve this requiredairflow control, a variable geometry device has been included inthe diffuser section.Key areas requiring additional study are also shown in the figure11-17. Combustor inlet airflow characteristics must be known toassure uniform fuel-air distributions. Engine transient cnar-acteristics must be identified and studied to avoid autoignitionand flashback in the fuel-air mixing passage. Practical scnemesfor varying the combustor geometry and controlling the operationof the combustor must be identified. Techniques for predictingand achieving the required fuei distribution and vaporization inthe premixing section of the main stage as needed. Autoignitionand flashback may also be problems there. More data on thesephenomena are needed over the full range of engine operating con-ditions, including transients. Other areas of the combustor alsorequire study. Lean stability and altitude relight capabilityneed special attention with these advanced concepts. Becausemost of the combustor airflow must pass through the main stage tosatisfy the lean burning requirement, the amount of air availableto cool the combustor liner will be less tnan tnat of currenttechnology combustors. It therefore appears likely that advancedliner cooling schemes will be required to avoid liner durabilityproblems.

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Digital engine controls will like ly be required for the addi-tional complexity of variable geometry. It is expected thatfull-authority digital-control technology will be available inthe future. However, additional study is needed to examine thecontrol aspects of variable geometry combustors and to establishtransient response requirements.Figure II-18 indicates the areas requiring research for catalyticcombustors. In general, all of the problem areas associated withpremixed combustion apply equally well to the catalytic con-cepts. Problems unique to this concept include the activity ofthe catalytic materials over wide operating ranges, long-termdegradation and poisoning of the catalyst, and thermal durabilityof the catalyst during continuous and cyclic operation of thecatalyst bed. While considerable progress has been made in thelast few years on research into catalyst and substrate materials,considerably more is required. It is not the purpose of thispaper to fully document the results to date of these fundamentalstudies in support of our far term program. However, threestudies that highlight some of the activities have been selectedfor discussion.Studies of the influence of flame-holder geometry on emissionsand performance have been undertaken. Figure II-19 shows theflame zone structure for six flame-holder designs that have beenevaluated. The designs tested included wire grids, perforatedplates, cones, and C-gutters. These open-duct burning photo-graphs were taken only for visualization purposes. Actual test-ing was done at high pressure in an enclosed flame-tube rig.Figure II-20 presents data from a high-pressure, lean, premixed,prevaporized flame-tube experiment. Emissions of oxides of ni-trogen are shown as a function of the adiabatic flame temper-ature. When this experiment was completed, several milestoneshad been passed. First, flame-tube experiments were for thefirst time conducted at pressures well above 10 atmospheres. Infact, these flame-tube experiments were successfully conducted atpressures to 30 atmospheres and inlet-air temperatures comparableto those of modern engines. Second, although previous data hadshown an inconsistentency in oxides of nitrogen emissions withincreasing pressure, this experiment demonstrated that, from 10to 30 atmospheres, pressure had no effect on oxides of nitrogenemissions in lean, premixed, prevaporized, combustors when cor-related against adiabatic flame temperature. Third, this experi-ment verified the emission levels projected from lower pressuretests.Tests are also underway to determine the characteristics ofboundary-layer autoignition and flashback phenomena in premixedfuel-air streams. As discussed earlier, an understanding ofautoignition and flashback phenomena are important in the designand operation of the premixer section. Figure II-21 is a photo-

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GASEOUS EMISSION STANDARDS

ENGINE CLASS POLLUTANT"HC CO NO,

Tl TURBOJET/TURBOFAN LESS THAN 8000 lb THRUST 1.6 9.4 3.7T2 TURBOJET/TURBOFAN GREATER THAN 8000 lb .8 4.3 3.0THRUSTT3 P&WJT3D .8 4.3 3.0T4 P & W JT8D .8 4.3 3.0T5 TURBOJET/TURBOFAN ENGINES FOR SUPERSONIC 3.9 30.1 9.0AIRCRAFTP2 TURBOPROP ENGINES 4. 9 26.8 12. 9"T STANDARDS AS lb/l000 lb THRUST-hr/CYCLEP STANDARDS AS lb/l000 hp-hr/CYCLE

Table II-Ics-/Y-175s

NEAR TERM EMISSION REDUCTION PROGRAMSCOPE

ENGINE MANUFACTURER ENGINECLASSP2 DETROIT-DIESEL-ALLISON 501-D22-ATl GARREll AIRESEARCH TFE-731-2T4 PRAll & WHITNEY JT8D-17T2 PRAll & WHITNEY JT9D-7T2 GENERAL ELECTRIC CF6-50

cs-79-1747Table II-Z

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CURRENTCOMBUSTORS

PRIMARY ZONE -/ & FILM-COOLING AIR10 5 10

5 5

0LI'Itl 0 0

2Or

0SMOKE NO.CS-cl-: j j,Y

CO, EPAP Tl-lC, EPAP NO,, EPAPFigure II-1

COMBUSTORMISSIONSOPERAilNGCwDlTlONS COMBUSTlONCHARACTERlSllCS POLLUTANTS REDUCTlONTECHNIQUES

LOW POWER/IDLE QUENCHING CARBON MONOXIDE INCREASE RESIDENCE TIMEPOOR COMBUSTION UNBURNED REDUCE FLOW VELOCITYSTABILITY HYDROCARBONS & DELAY MIXINGPOOR FUEL INCREASE EQUIV RATIOATOMIZATION & IMPROVE FUEL ATOMIZATIONDISTRIBUTlON 8, DISTRIBUTION

HIGH POWER/TAKEOFF EXCESS RESIDENCETIMEHIGH FLAME TEMPPOOR LOCAL FUELDISTRIBUTION

OXIDES OF DECREASE RESIDENCE nMENITROGEN INCREASE FLOW VELOCITYSMOKE ENHANCE MIXINGDECREASE EQUIV RATlOIMPROVE LOCAL FUELDlSTRlBUllONFigure II-2 cs-79-2054

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EMISSIONREDUCTIONROBLEM

LOW POWERIDLE

INCREASE RESIDENCE TIMEREDUCE FLOW VELOCITYDELAY MIXINGINCREASE EQUIVALENCE RATlO

IMPROVE FUEL ATOMIZATION & DISTRIBUTION

HIGH POWERTAKEOFF

DECREASERESIDENCE TIMEINCREASE FLOW VELOCITYENHANCE MIXINGDECREASE EQUIV RATIOIMPROVE LOCAL FUEL DISTRIBUTION

Figure II-3 cs-79-2042

VORBIXCOMBUSTORONCEPTORJTSD-7ENGINE

PRODUCTlON

CO, EPAP MC, EPAP NO,, EPAPFigure II-4

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PROTOTYPE VORBIX COMBUSTOR

Fiaure II-5

DOUBLE/ANNULAROMBUSTOROR 36-50 ENGINEFUEL-, tl

- PRODUCTIONa DOUBLE ANNULAR

CO, EPAP THC, EPAP NO,, EPAP ,- ,I. ,,IFigure II-6

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-,

HOTWALLCOMBUSTORCONCEPT NO. 1

SHELLS WITH IMPINGEMENTCOOLED LINERS 7 n

SWIRL CUP AIR mucs-79-1751

PRIMARYDILUTION

Figure II-9

SECbNDARYDILUTION

RECUPERATIVEOOLING OMBUSTORCONCEPT NO. 2

SWIRLER AIR SHELLS WITH IMPINGEMENTWITH ENHANCED COOLED LINERS7 m nTEMP I \FUEL INTOSIMPLEX -NOZZLES

PRIMARYDILUTION AIR iiiAi HIGHER TEMP jFigure II-10

Ill SECONDARY /IDILUTION uCS-,-:(/

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CATALYTICONVERTEROMBUSTORCONCEPT NO. 3

SHELLS WITH IMPINGEMENT ,-HONEYCOMB

LlNITIAL BURNING ZONEFigure II-11

CATALYTICOMBUSTORSSEMBLY

Figu re II -12

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EMISSI